Tracking performance and global stability guaranteed neural control of uncertain hypersonic flight vehicle

Directory of Open Access Journals (Sweden)

Tao Teng

2016-02-01

Full Text Available In this article, a global adaptive neural dynamic surface control with predefined tracking performance is developed for a class of hypersonic flight vehicles, whose accurate dynamics is hard to obtain. The control scheme developed in this paper overcomes the limitations of neural approximation region by employing a switching mechanism which incorporates an additional robust controller outside the neural approximation region to pull the transient state variables back when they overstep the neural approximation region, such that globally uniformly ultimately bounded stability can be guaranteed. Especially, the developed global adaptive neural control also improves the tracking performance by introducing an error transformation mechanism, such that both transient and steady-state performance can be shaped according to the predefined bounds. Simulation studies on the hypersonic flight vehicle validate that the designed controller has good velocity modulation and velocity stability performance.

Application of CFD to a generic hypersonic flight research study

Science.gov (United States)

Green, Michael J.; Lawrence, Scott L.; Dilley, Arthur D.; Hawkins, Richard W.; Walker, Mary M.; Oberkampf, William L.

1993-01-01

Computational analyses have been performed for the initial assessment of flight research vehicle concepts that satisfy requirements for potential hypersonic experiments. Results were obtained from independent analyses at NASA Ames, NASA Langley, and Sandia National Labs, using sophisticated time-dependent Navier-Stokes and parabolized Navier-Stokes methods. Careful study of a common problem consisting of hypersonic flow past a slightly blunted conical forebody was undertaken to estimate the level of uncertainty in the computed results, and to assess the capabilities of current computational methods for predicting boundary-layer transition onset. Results of this study in terms of surface pressure and heat transfer comparisons, as well as comparisons of boundary-layer edge quantities and flow-field profiles are presented here. Sensitivities to grid and gas model are discussed. Finally, representative results are presented relating to the use of Computational Fluid Dynamics in the vehicle design and the integration/support of potential experiments.

The HYTHIRM Project: Flight Thermography of the Space Shuttle During the Hypersonic Re-entry

Science.gov (United States)

Horvath, Thomas J.; Tomek, Deborah M.; Berger, Karen T.; Zalameda, Joseph N.; Splinter, Scott C.; Krasa, Paul W.; Schwartz, Richard J.; Gibson, David M.; Tietjen, Alan B.; Tack, Steve

2010-01-01

This report describes a NASA Langley led endeavor sponsored by the NASA Engineering Safety Center, the Space Shuttle Program Office and the NASA Aeronautics Research Mission Directorate to demonstrate a quantitative thermal imaging capability. A background and an overview of several multidisciplinary efforts that culminated in the acquisition of high resolution calibrated infrared imagery of the Space Shuttle during hypervelocity atmospheric entry is presented. The successful collection of thermal data has demonstrated the feasibility of obtaining remote high-resolution infrared imagery during hypersonic flight for the accurate measurement of surface temperature. To maximize science and engineering return, the acquisition of quantitative thermal imagery and capability demonstration was targeted towards three recent Shuttle flights - two of which involved flight experiments flown on Discovery. In coordination with these two Shuttle flight experiments, a US Navy NP-3D aircraft was flown between 26-41 nautical miles below Discovery and remotely monitored surface temperature of the Orbiter at Mach 8.4 (STS-119) and Mach 14.7 (STS-128) using a long-range infrared optical package referred to as Cast Glance. This same Navy aircraft successfully monitored the Orbiter Atlantis traveling at approximately Mach 14.3 during its return from the successful Hubble repair mission (STS-125). The purpose of this paper is to describe the systematic approach used by the Hypersonic Thermodynamic Infrared Measurements team to develop and implement a set of mission planning tools designed to establish confidence in the ability of an imaging platform to reliably acquire, track and return global quantitative surface temperatures of the Shuttle during entry. The mission planning tools included a pre-flight capability to predict the infrared signature of the Shuttle. Such tools permitted optimization of the hardware configuration to increase signal-to-noise and to maximize the available

Numerical study on aerodynamic heat of hypersonic flight

Directory of Open Access Journals (Sweden)

Huang Haiming

2016-01-01

Full Text Available Accurate prediction of the shock wave has a significant effect on the development of space transportation vehicle or exploration missions. Taking Lobb sphere as the example, the aerodynamic heat of hypersonic flight in different Mach numbers is simulated by the finite volume method. Chemical reactions and non-equilibrium heat are taken into account in this paper, where convective flux of the space term adopts the Roe format, and discretization of the time term is achieved by backward Euler algorithm. The numerical results reveal that thick mesh can lead to accurate prediction, and the thickness of the shock wave decreases as grid number increases. Furthermore, most of kinetic energy converts into internal energy crossing the shock wave.

Tracking error constrained robust adaptive neural prescribed performance control for flexible hypersonic flight vehicle

Directory of Open Access Journals (Sweden)

Zhonghua Wu

2017-02-01

Full Text Available A robust adaptive neural control scheme based on a back-stepping technique is developed for the longitudinal dynamics of a flexible hypersonic flight vehicle, which is able to ensure the state tracking error being confined in the prescribed bounds, in spite of the existing model uncertainties and actuator constraints. Minimal learning parameter technique–based neural networks are used to estimate the model uncertainties; thus, the amount of online updated parameters is largely lessened, and the prior information of the aerodynamic parameters is dispensable. With the utilization of an assistant compensation system, the problem of actuator constraint is overcome. By combining the prescribed performance function and sliding mode differentiator into the neural back-stepping control design procedure, a composite state tracking error constrained adaptive neural control approach is presented, and a new type of adaptive law is constructed. As compared with other adaptive neural control designs for hypersonic flight vehicle, the proposed composite control scheme exhibits not only low-computation property but also strong robustness. Finally, two comparative simulations are performed to demonstrate the robustness of this neural prescribed performance controller.

Flight Experiments for Hypersonic Vehicle Development (Experimentations envol pour le developpement d'un vehicule hypersonique) (CD-ROM)

National Research Council Canada - National Science Library

2007-01-01

...: 1 CD-ROM; 4 3/4 in.; 124 MB. ABSTRACT: This RTO-AVT/VKI Lecture Series brought together specialists from Europe, USA, and Russia to discuss flight experiments that pertain to the development of hypersonic vehicles...

Determination of the hypersonic-continuum/rarefied-flow drag coefficient of the Viking lander capsule 1 aeroshell from flight data

Science.gov (United States)

Blanchard, R. C.; Walberg, G. D.

1980-01-01

Results of an investigation to determine the full scale drag coefficient in the high speed, low density regime of the Viking lander capsule 1 entry vehicle are presented. The principal flight data used in the study were from onboard pressure, mass spectrometer, and accelerometer instrumentation. The hypersonic continuum flow drag coefficient was unambiguously obtained from pressure and accelerometer data; the free molecule flow drag coefficient was indirectly estimated from accelerometer and mass spectrometer data; the slip flow drag coefficient variation was obtained from an appropriate scaling of existing experimental sphere data. Comparison of the flight derived drag hypersonic continuum flow regime except for Reynolds numbers from 1000 to 100,000, for which an unaccountable difference between flight and ground test data of about 8% existed. The flight derived drag coefficients in the free molecule flow regime were considerably larger than those previously calculated with classical theory. The general character of the previously determined temperature profile was not changed appreciably by the results of this investigation; however, a slightly more symmetrical temperature variation at the highest altitudes was obtained.

High-Lift Propeller Noise Prediction for a Distributed Electric Propulsion Flight Demonstrator

Science.gov (United States)

Nark, Douglas M.; Buning, Pieter G.; Jones, William T.; Derlaga, Joseph M.

2017-01-01

Over the past several years, the use of electric propulsion technologies within aircraft design has received increased attention. The characteristics of electric propulsion systems open up new areas of the aircraft design space, such as the use of distributed electric propulsion (DEP). In this approach, electric motors are placed in many different locations to achieve increased efficiency through integration of the propulsion system with the airframe. Under a project called Scalable Convergent Electric Propulsion Technology Operations Research (SCEPTOR), NASA is designing a flight demonstrator aircraft that employs many "high-lift propellers" distributed upstream of the wing leading edge and two cruise propellers (one at each wingtip). As the high-lift propellers are operational at low flight speeds (take-off/approach flight conditions), the impact of the DEP configuration on the aircraft noise signature is also an important design consideration. This paper describes efforts toward the development of a mulit-fidelity aerodynamic and acoustic methodology for DEP high-lift propeller aeroacoustic modeling. Specifically, the PAS, OVERFLOW 2, and FUN3D codes are used to predict the aerodynamic performance of a baseline high-lift propeller blade set. Blade surface pressure results from the aerodynamic predictions are then used with PSU-WOPWOP and the F1A module of the NASA second generation Aircraft NOise Prediction Program to predict the isolated high-lift propeller noise source. Comparisons of predictions indicate that general trends related to angle of attack effects at the blade passage frequency are captured well with the various codes. Results for higher harmonics of the blade passage frequency appear consistent for the CFD based methods. Conversely, evidence of the need for a study of the effects of increased azimuthal grid resolution on the PAS based results is indicated and will be pursued in future work. Overall, the results indicate that the computational

Weakly Ionized Plasmas in Hypersonics: Fundamental Kinetics and Flight Applications

International Nuclear Information System (INIS)

Macheret, Sergey

2005-01-01

The paper reviews some of the recent studies of applications of weakly ionized plasmas to supersonic/hypersonic flight. Plasmas can be used simply as means of delivering energy (heating) to the flow, and also for electromagnetic flow control and magnetohydrodynamic (MHD) power generation. Plasma and MHD control can be especially effective in transient off-design flight regimes. In cold air flow, nonequilibrium plasmas must be created, and the ionization power budget determines design, performance envelope, and the very practicality of plasma/MHD devices. The minimum power budget is provided by electron beams and repetitive high-voltage nanosecond pulses, and the paper describes theoretical and computational modeling of plasmas created by the beams and repetitive pulses. The models include coupled equations for non-local and unsteady electron energy distribution function (modeled in forward-back approximation), plasma kinetics, and electric field. Recent experimental studies at Princeton University have successfully demonstrated stable diffuse plasmas sustained by repetitive nanosecond pulses in supersonic air flow, and for the first time have demonstrated the existence of MHD effects in such plasmas. Cold-air hypersonic MHD devices are shown to permit optimization of scramjet inlets at Mach numbers higher than the design value, while operating in self-powered regime. Plasma energy addition upstream of the inlet throat can increase the thrust by capturing more air (Virtual Cowl), or it can reduce the flow Mach number and thus eliminate the need for an isolator duct. In the latter two cases, the power that needs to be supplied to the plasma would be generated by an MHD generator downstream of the combustor, thus forming the 'reverse energy bypass' scheme. MHD power generation on board reentry vehicles is also discussed

In-flight measurements of propeller blade deformation on a VUT100 cobra aeroplane using a co-rotating camera system

Science.gov (United States)

Boden, F.; Stasicki, B.; Szypuła, M.; Ružička, P.; Tvrdik, Z.; Ludwikowski, K.

2016-07-01

Knowledge of propeller or rotor blade behaviour under real operating conditions is crucial for optimizing the performance of a propeller or rotor system. A team of researchers, technicians and engineers from Avia Propeller, DLR, EVEKTOR and HARDsoft developed a rotating stereo camera system dedicated to in-flight blade deformation measurements. The whole system, co-rotating with the propeller at its full speed and hence exposed to high centrifugal forces and strong vibration, had been successfully tested on an EVEKTOR VUT 100 COBRA aeroplane in Kunovice (CZ) within the project AIM2—advanced in-flight measurement techniques funded by the European Commission (contract no. 266107). This paper will describe the work, starting from drawing the first sketch of the system up to performing the successful flight test. Apart from a description of the measurement hardware and the applied IPCT method, the paper will give some impressions of the flight test activities and discuss the results obtained from the measurements.

Nonlinear Robust Control of a Hypersonic Flight Vehicle Using Fuzzy Disturbance Observer

Directory of Open Access Journals (Sweden)

Lei Zhengdong

2013-01-01

Full Text Available This paper is concerned with a novel tracking controller design for a hypersonic flight vehicle in complex and volatile environment. The attitude control model is challengingly constructed with multivariate uncertainties and external disturbances, such as structure dynamic and stochastic wind disturbance. In order to resist the influence of uncertainties and disturbances on the flight control system, nonlinear disturbance observer is introduced to estimate them. Moreover, for the sake of high accuracy and sensitivity, fuzzy theory is adopted to improve the performance of the nonlinear disturbance observer. After the total disturbance is eliminated by dynamic inversion method, a cascade system is obtained and then stabilized by a sliding-mode controller. Finally, simulation results show that the strong robust controller achieves excellent performance when the closed-loop control system is influenced by mass uncertainties and external disturbances.

Scaled Rocket Testing in Hypersonic Flow

Science.gov (United States)

Dufrene, Aaron; MacLean, Matthew; Carr, Zakary; Parker, Ron; Holden, Michael; Mehta, Manish

2015-01-01

NASA's Space Launch System (SLS) uses four clustered liquid rocket engines along with two solid rocket boosters. The interaction between all six rocket exhaust plumes will produce a complex and severe thermal environment in the base of the vehicle. This work focuses on a recent 2% scale, hot-fire SLS base heating test. These base heating tests are short-duration tests executed with chamber pressures near the full-scale values with gaseous hydrogen/oxygen engines and RSRMV analogous solid propellant motors. The LENS II shock tunnel/Ludwieg tube tunnel was used at or near flight duplicated conditions up to Mach 5. Model development was strongly based on the Space Shuttle base heating tests with several improvements including doubling of the maximum chamber pressures and duplication of freestream conditions. Detailed base heating results are outside of the scope of the current work, rather test methodology and techniques are presented along with broader applicability toward scaled rocket testing in supersonic and hypersonic flow.

Experimental Investigation of Brazilian 14-X B Hypersonic Scramjet Aerospace Vehicle

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João Felipe de Araujo Martos

2017-01-01

Full Text Available The Brazilian hypersonic scramjet aerospace vehicle 14-X B is a technological demonstrator of a hypersonic airbreathing propulsion system based on the supersonic combustion (scramjet to be tested in flight into the Earth’s atmosphere at an altitude of 30 km and Mach number 7. The 14-X B has been designed at the Prof. Henry T. Nagamatsu Laboratory of Aerothermodynamics and Hypersonics, Institute for Advanced Studies (IEAv, Brazil. The IEAv T3 Hypersonic Shock Tunnel is a ground-test facility able to produce high Mach number and high enthalpy flows in the test section close to those encountered during the flight of the 14-X B into the Earth’s atmosphere at hypersonic flight speeds. A 1 m long stainless steel 14-X B model was experimentally investigated at T3 Hypersonic Shock Tunnel, for freestream Mach numbers ranging from 7 to 8. Static pressure measurements along the lower surface of the 14-X B, as well as high-speed Schlieren photographs taken from the 5.5° leading edge and the 14.5° deflection compression ramp, provided experimental data. Experimental data was compared to the analytical theoretical solutions and the computational fluid dynamics (CFD simulations, showing good qualitative agreement and in consequence demonstrating the importance of these methods in the project of the 14-X B hypersonic scramjet aerospace vehicle.

Flight demonstration of new thruster and green propellant technology on the PRISMA satellite

Science.gov (United States)

Anflo, K.; Möllerberg, R.

2009-11-01

The concept of a storable liquid monopropellant blend for space applications based on ammonium dinitramide (ADN) was invented in 1997, within a co-operation between the Swedish Space Corporation (SSC) and the Swedish Defense Research Agency (FOI). The objective was to develop a propellant which has higher performance and is safer than hydrazine. The work has been performed under contract from the Swedish National Space Board and ESA. The progress of the development has been presented in several papers since 2000. ECAPS, a subsidiary of the Swedish Space Corporation was established in 2000 with the aim to develop and market the novel "high performance green propellant" (HPGP) technology for space applications. The new technology is based on several innovations and patents w.r.t. propellant formulation and thruster design, including a high temperature resistant catalyst and thrust chamber. The first flight demonstration of the HPGP propulsion system will be performed on PRISMA. PRISMA is an international technology demonstration program with Swedish Space Corporation as the Prime Contractor. This paper describes the performance, characteristics, design and verification of the HPGP propulsion system for PRISMA. Compatibility issues related to using a new propellant with COTS components is also discussed. The PRISMA mission includes two satellites in LEO orbit were the focus is on rendezvous and formation flying. One of the satellites will act as a "target" and the main spacecraft performs rendezvous and formation flying maneuvers, where the ECAPS HPGP propulsion system will provide delta-V capability. The PRISMA CDR was held in January 2007. Integration of the flight propulsion system is about to be finalized. The flight opportunity on PRISMA represents a unique opportunity to demonstrate the HPGP propulsion system in space, and thus take a significant step towards its use in future space applications. The launch of PRISMA scheduled to 2009.

Experimental Studies on Hypersonic Stagnation Point Chemical Environment

National Research Council Canada - National Science Library

Chazot, O

2006-01-01

Development of space transportation is a very challenging task. Hypersonic flight should be investigated in details to allow designing spacecraft according to the severe environment of their flight conditions...

In-flight measurements of aircraft propeller deformation by means of an autarkic fast rotating imaging system

Science.gov (United States)

Stasicki, Boleslaw; Boden, Fritz

2015-03-01

The non-intrusive in-flight measurement of the deformation and pitch of the aircraft propeller is a demanding task. The idea of an imaging system integrated and rotating with the aircraft propeller has been presented on the 30th International Congress on High-Speed Imaging and Photonics (ICHSIP30) in 2012. Since then this system has been constructed and tested in the laboratory as well as on the real aircraft. In this paper we outline the principle of Image Pattern Correlation Technique (IPCT) based on Digital Image Correlation (DIC) and describe the construction of a dedicated autarkic 3D camera system placed on the investigated propeller and rotating at its full speed. Furthermore, the results of the first ground and in-flight tests are shown and discussed. This development has been found by the European Commission within the 7th frame project AIM2 (contract no. 266107).

Experimental Investigation of Brazilian 14-X B Hypersonic Scramjet Aerospace Vehicle

OpenAIRE

de Araujo Martos, João Felipe; da Silveira Rêgo, Israel; Pachon Laiton, Sergio Nicholas; Lima, Bruno Coelho; Costa, Felipe Jean; de Paula Toro, Paulo Gilberto

2017-01-01

The Brazilian hypersonic scramjet aerospace vehicle 14-X B is a technological demonstrator of a hypersonic airbreathing propulsion system based on the supersonic combustion (scramjet) to be tested in flight into the Earth’s atmosphere at an altitude of 30 km and Mach number 7. The 14-X B has been designed at the Prof. Henry T. Nagamatsu Laboratory of Aerothermodynamics and Hypersonics, Institute for Advanced Studies (IEAv), Brazil. The IEAv T3 Hypersonic Shock Tunnel is a ground-test facility...

Guidance Law and Neural Control for Hypersonic Missile to Track Targets

Directory of Open Access Journals (Sweden)

Wenxing Fu

2016-01-01

Full Text Available Hypersonic technology plays an important role in prompt global strike. Because the flight dynamics of a hypersonic vehicle is nonlinear, uncertain, and highly coupled, the controller design is challenging, especially to design its guidance and control law during the attack of a maneuvering target. In this paper, the sliding mode control (SMC method is used to develop the guidance law from which the desired flight path angle is derived. With the desired information as control command, the adaptive neural control in discrete time is investigated ingeniously for the longitudinal dynamics of the hypersonic missile. The proposed guidance and control laws are validated by simulation of a hypersonic missile against a maneuvering target. It is demonstrated that the scheme has good robustness and high accuracy to attack a maneuvering target in the presence of external disturbance and missile model uncertainty.

Retooling CFD for hypersonic aircraft

Science.gov (United States)

Dwoyer, Douglas L.; Kutler, Paul; Povinelli, Louis A.

1987-01-01

The CFD facility requirements of hypersonic aircraft configuration design development are different from those thus far employed for reentry vehicle design, because (1) the airframe and the propulsion system must be fully integrated to achieve the desired performance; (2) the vehicle must be reusable, with minimum refurbishment requirements between flights; and (3) vehicle performance must be optimized for a wide range of Mach numbers. An evaluation is presently made of flow resolution within shock waves, transition and turbulence phenomenon tractability, chemical reaction modeling, and hypersonic boundary layer transition, with state-of-the-art CFD.

TBCC Discipline Overview. Hypersonics Project

Science.gov (United States)

Thomas, Scott R.

2011-01-01

The "National Aeronautics Research and Development Policy" document, issued by the National Science and Technology Council in December 2006, stated that one (among several) of the guiding objectives of the federal aeronautics research and development endeavors shall be stable and long-term foundational research efforts. Nearly concurrently, the National Academies issued a more technically focused aeronautics blueprint, entitled: the "Decadal Survey of Civil Aeronautics - Foundations for the Future." Taken together these documents outline the principles of an aeronautics maturation plan. Thus, in response to these overarching inputs (and others), the National Aeronautics and Space Administration (NASA) organized the Fundamental Aeronautics Program (FAP), a program within the NASA Aeronautics Research Mission Directorate (ARMD). The FAP initiated foundational research and technology development tasks to enable the capability of future vehicles that operate across a broad range of Mach numbers, inclusive of the subsonic, supersonic, and hypersonic flight regimes. The FAP Hypersonics Project concentrates on two hypersonic missions: (1) Air-breathing Access to Space (AAS) and (2) the (Planetary Atmospheric) Entry, Decent, and Landing (EDL). The AAS mission focuses on Two-Stage-To-Orbit (TSTO) systems using air-breathing combined-cycle-engine propulsion; whereas, the EDL mission focuses on the challenges associated with delivering large payloads to (and from) Mars. So, the FAP Hypersonic Project investments are aligned to achieve mastery and intellectual stewardship of the core competencies in the hypersonic-flight regime, which ultimately will be required for practical systems with highly integrated aerodynamic/vehicle and propulsion/engine technologies. Within the FAP Hypersonics, the technology management is further divided into disciplines including one targeting Turbine-Based Combine-Cycle (TBCC) propulsion. Additionally, to obtain expertise and support from outside

Hypersonic drone design: A multidisciplinary experience

Science.gov (United States)

1988-01-01

Efforts were focused on design problems of an unmanned hypersonic vehicle. It is felt that a scaled hypersonic drone is necessary to bridge the gap between present theory on hypersonics and the future reality of the National Aerospace Plane (NASP) for two reasons: to fulfill a need for experimental data in the hypersonic regime, and to provide a testbed for the scramjet engine which is to be the primary mode of propulsion for the NASP. Three areas of great concern to NASP design were examined: propulsion, thermal management, and flight systems. Problem solving in these areas was directed towards design of the drone with the idea that the same design techniques could be applied to the NASP. A seventy degree swept double delta wing configuration, developed in the 70's at NASA Langley, was chosen as the aerodynamic and geometric model for the drone. This vehicle would be air-launched from a B-1 at Mach 0.8 and 48,000 feet, rocket boosted by two internal engines to Mach 10 and 100,000 feet, and allowed to cruise under power of the scramjet engine until burnout. It would then return to base for an unpowered landing. Preliminary energy calculations based upon the flight requirements give the drone a gross launch weight of 134,000 lb. and an overall length of 85 feet.

CFD analysis of hypersonic, chemically reacting flow fields

Science.gov (United States)

Edwards, T. A.

1993-01-01

Design studies are underway for a variety of hypersonic flight vehicles. The National Aero-Space Plane will provide a reusable, single-stage-to-orbit capability for routine access to low earth orbit. Flight-capable satellites will dip into the atmosphere to maneuver to new orbits, while planetary probes will decelerate at their destination by atmospheric aerobraking. To supplement limited experimental capabilities in the hypersonic regime, computational fluid dynamics (CFD) is being used to analyze the flow about these configurations. The governing equations include fluid dynamic as well as chemical species equations, which are being solved with new, robust numerical algorithms. Examples of CFD applications to hypersonic vehicles suggest an important role this technology will play in the development of future aerospace systems. The computational resources needed to obtain solutions are large, but solution adaptive grids, convergence acceleration, and parallel processing may make run times manageable.

Disturbance observer-based L1 robust tracking control for hypersonic vehicles with T-S disturbance modeling

Directory of Open Access Journals (Sweden)

Yang Yi

2016-11-01

Full Text Available This article concerns a disturbance observer-based L1 robust anti-disturbance tracking algorithm for the longitudinal models of hypersonic flight vehicles with different kinds of unknown disturbances. On one hand, by applying T-S fuzzy models to represent those modeled disturbances, a disturbance observer relying on T-S disturbance models can be constructed to track the dynamics of exogenous disturbances. On the other hand, L1 index is introduced to analyze the attenuation performance of disturbance for those unmodeled disturbances. By utilizing the existing convex optimization algorithm, a disturbance observer-based proportional-integral-controlled input is proposed such that the stability of hypersonic flight vehicles can be ensured and the tracking error for velocity and altitude in hypersonic flight vehicle models can converge to equilibrium point. Furthermore, the satisfactory disturbance rejection and attenuation with L1 index can be obtained simultaneously. Simulation results on hypersonic flight vehicle models can reflect the feasibility and effectiveness of the proposed control algorithm.

Elevator Sizing, Placement, and Control-Relevant Tradeoffs for Hypersonic Vehicles

Science.gov (United States)

Dickeson, Jeffrey J.; Rodriguez, Armando A.; Sridharan, Srikanth; Korad, Akshay

2010-01-01

Within this paper, control-relevant vehicle design concepts are examined using a widely used 3 DOF (plus flexibility) nonlinear model for the longitudinal dynamics of a generic carrot-shaped scramjet powered hypersonic vehicle. The impact of elevator size and placement on control-relevant static properties (e.g. level-flight trimmable region, trim controls, Angle of Attack (AOA), thrust margin) and dynamic properties (e.g. instability and right half plane zero associated with flight path angle) are examined. Elevator usage has been examine for a class of typical hypersonic trajectories.

Surface Heat Flux and Pressure Distribution on a Hypersonic Blunt Body With DEAS

Science.gov (United States)

Salvador, I. I.; Minucci, M. A. S.; Toro, P. G. P.; Oliveira, A. C.; Channes, J. B.

2008-04-01

With the currently growing interest for advanced technologies to enable hypersonic flight comes the Direct Energy Air Spike concept, where pulsed beamed laser energy is focused upstream of a blunt flight vehicle to disrupt the flow structure creating a virtual, slender body geometry. This allies in the vehicle both advantages of a blunt body (lower thermal stresses) to that of a slender geometry (lower wave drag). The research conducted at the Henry T. Nagamatsu Laboratory for Aerodynamics and Hypersonics focused on the measurement of the surface pressure and heat transfer rates on a blunt model. The hypersonic flight conditions were simulated at the HTN Laboratory's 0.3 m T2 Hypersonic Shock Tunnel. During the tests, the laser energy was focused upstream the model by an infrared telescope to create the DEAS effect, which was supplied by a TEA CO2 laser. Piezoelectric pressure transducers were used for the pressure measurements and fast response coaxial thermocouples were used for the measurement of surface temperature, which was later used for the estimation of the wall heat transfer using the inverse heat conduction theory.

Method and system for control of upstream flowfields of vehicle in supersonic or hypersonic atmospheric flight

Science.gov (United States)

Daso, Endwell O. (Inventor); Pritchett, II, Victor E. (Inventor); Wang, Ten-See (Inventor); Farr, Rebecca Ann (Inventor)

2012-01-01

The upstream flowfield of a vehicle traveling in supersonic or hypersonic atmospheric flight is actively controlled using attribute(s) experienced by the vehicle. Sensed attribute(s) include pressure along the vehicle's outer mold line, temperature along the vehicle's outer mold line, heat flux along the vehicle's outer mold line, and/or local acceleration response of the vehicle. A non-heated, non-plasma-producing gas is injected into an upstream flowfield of the vehicle from at least one surface location along the vehicle's outer mold line. The pressure of the gas so-injected is adjusted based on the attribute(s) so-sensed.

Geometry Modeling and Adaptive Control of Air-Breathing Hypersonic Vehicles

Science.gov (United States)

Vick, Tyler Joseph

Air-breathing hypersonic vehicles have the potential to provide global reach and affordable access to space. Recent technological advancements have made scramjet-powered flight achievable, as evidenced by the successes of the X-43A and X-51A flight test programs over the last decade. Air-breathing hypersonic vehicles present unique modeling and control challenges in large part due to the fact that scramjet propulsion systems are highly integrated into the airframe, resulting in strongly coupled and often unstable dynamics. Additionally, the extreme flight conditions and inability to test fully integrated vehicle systems larger than X-51 before flight leads to inherent uncertainty in hypersonic flight. This thesis presents a means to design vehicle geometries, simulate vehicle dynamics, and develop and analyze control systems for hypersonic vehicles. First, a software tool for generating three-dimensional watertight vehicle surface meshes from simple design parameters is developed. These surface meshes are compatible with existing vehicle analysis tools, with which databases of aerodynamic and propulsive forces and moments can be constructed. A six-degree-of-freedom nonlinear dynamics simulation model which incorporates this data is presented. Inner-loop longitudinal and lateral control systems are designed and analyzed utilizing the simulation model. The first is an output feedback proportional-integral linear controller designed using linear quadratic regulator techniques. The second is a model reference adaptive controller (MRAC) which augments this baseline linear controller with an adaptive element. The performance and robustness of each controller are analyzed through simulated time responses to angle-of-attack and bank angle commands, while various uncertainties are introduced. The MRAC architecture enables the controller to adapt in a nonlinear fashion to deviations from the desired response, allowing for improved tracking performance, stability, and

Requirements for facilities and measurement techniques to support CFD development for hypersonic aircraft

Science.gov (United States)

Sellers, William L., III; Dwoyer, Douglas L.

1992-01-01

The design of a hypersonic aircraft poses unique challenges to the engineering community. Problems with duplicating flight conditions in ground based facilities have made performance predictions risky. Computational fluid dynamics (CFD) has been proposed as an additional means of providing design data. At the present time, CFD codes are being validated based on sparse experimental data and then used to predict performance at flight conditions with generally unknown levels of uncertainty. This paper will discuss the facility and measurement techniques that are required to support CFD development for the design of hypersonic aircraft. Illustrations are given of recent success in combining experimental and direct numerical simulation in CFD model development and validation for hypersonic perfect gas flows.

Hypersonic drone vehicle design: A multidisciplinary experience

Science.gov (United States)

1988-01-01

UCLA's Advanced Aeronautic Design group focussed their efforts on design problems of an unmanned hypersonic vehicle. It is felt that a scaled hypersonic drone is necesary to bridge the gap between present theory on hypersonics and the future reality of the National Aerospace Plane (NASP) for two reasons: (1) to fulfill a need for experimental data in the hypersonic regime, and (2) to provide a testbed for the scramjet engine which is to be the primary mode of propulsion for the NASP. The group concentrated on three areas of great concern to NASP design: propulsion, thermal management, and flight systems. Problem solving in these areas was directed toward design of the drone with the idea that the same design techniques could be applied to the NASP. A 70 deg swept double-delta wing configuration, developed in the 70's at the NASA Langley, was chosen as the aerodynamic and geometric model for the drone. This vehicle would be air launched from a B-1 at Mach 0.8 and 48,000 feet, rocket boosted by two internal engines to Mach 10 and 100,000 feet, and allowed to cruise under power of the scramjet engine until burnout. It would then return to base for an unpowered landing. Preliminary energy calculations based on flight requirements give the drone a gross launch weight of 134,000 pounds and an overall length of 85 feet.

Sliding mode disturbance observer-enhanced adaptive control for the air-breathing hypersonic flight vehicle

Science.gov (United States)

An, Hao; Wang, Changhong; Fidan, Baris

2017-10-01

This paper presents a backstepping procedure to design an adaptive controller for the air-breathing hypersonic flight vehicle (AHFV) subject to external disturbances and actuator saturations. In each step, a sliding mode exact disturbance observer (SMEDO) is exploited to exactly estimate the lumped disturbance in finite time. Specific dynamics are introduced to handle the possible actuator saturations. Based on SMEDO and introduced dynamics, an adaptive control law is designed, along with the consideration on ;explosion of complexity; in backstepping design. The developed controller is equipped with fast disturbance rejection and great capability to accommodate the saturated actuators, which also lead to a wider application scope. A simulation study is provided to show the effectiveness and superiority of the proposed controller.

Investigation of piloting aids for manual control of hypersonic maneuvers

Science.gov (United States)

Raney, David L.; Phillips, Michael R.; Person, Lee H., Jr.

1995-01-01

An investigation of piloting aids designed to provide precise maneuver control for an air-breathing hypersonic vehicle is described. Stringent constraints and nonintuitive high-speed flight effects associated with maneuvering in the hypersonic regime raise the question of whether manual control of such a vehicle should even be considered. The objectives of this research were to determine the extent of manual control that is desirable for a vehicle maneuvering in this regime and to identify the form of aids that must be supplied to the pilot to make such control feasible. A piloted real-time motion-based simulation of a hypersonic vehicle concept was used for this study, and the investigation focused on a single representative cruise turn maneuver. Piloting aids, which consisted of an auto throttle, throttle director, autopilot, flight director, and two head-up display configurations, were developed and evaluated. Two longitudinal control response types consisting of a rate-command/attitude-hold system and a load factor-rate/load-factor-hold system were also compared. The complete set of piloting aids, which consisted of the autothrottle, throttle director, and flight director, improved the average Cooper-Harper flying qualities ratings from 8 to 2.6, even though identical inner-loop stability and control augmentation was provided in all cases. The flight director was determined to be the most critical of these aids, and the cruise turn maneuver was unachievable to adequate performance specifications in the absence of this flight director.

Robust Switching Control for Hypersonic Vehicles, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Flight in the hypersonic regime is critical to NASA's goals because access to earth orbit and re-entry from orbit to earth or to other planets with atmospheres...

Flight Investigation of the Performance of a Two-stage Solid-propellant Nike-deacon (DAN) Meteorological Sounding Rocket

Science.gov (United States)

Heitkotter, Robert H

1956-01-01

A flight investigation of two Nike-Deacon (DAN) two-stage solid-propellant rocket vehicles indicated satisfactory performance may be expected from the DAN meteorological sounding rocket. Peak altitudes of 356,000 and 350,000 feet, respectively, were recorded for the two flight tests when both vehicles were launched from sea level at an elevation angle of 75 degrees. Performance calculations based on flight-test results show that altitudes between 358,000 feet and 487,000 feet may be attained with payloads varying between 60 pounds and 10 pounds.

Hypersonic CFD applications for the National Aero-Space Plane

Science.gov (United States)

Richardson, Pamela F.; Mcclinton, Charles R.; Bittner, Robert D.; Dilley, A. Douglas; Edwards, Kelvin W.

1989-01-01

Design and analysis of the NASP depends heavily upon developing the critical technology areas that cover the entire engineering design of the vehicle. These areas include materials, structures, propulsion systems, propellants, integration of airframe and propulsion systems, controls, subsystems, and aerodynamics areas. Currently, verification of many of the classical engineering tools relies heavily on computational fluid dynamics. Advances are being made in the development of CFD codes to accomplish nose-to-tail analyses for hypersonic aircraft. Additional details involving the partial development, analysis, verification, and application of the CFL3D code and the SPARK combustor code are discussed. A nonequilibrium version of CFL3D that is presently being developed and tested is also described. Examples are given of portion calculations for research hypersonic aircraft geometries and comparisons with experiment data show good agreement.

Heating Augmentation for Short Hypersonic Protuberances

Science.gov (United States)

Mazaheri, Ali R.; Wood, William A.

2008-01-01

Computational aeroheating analyses of the Space Shuttle Orbiter plug repair models are validated against data collected in the Calspan University of Buffalo Research Center (CUBRC) 48 inch shock tunnel. The comparison shows that the average difference between computed heat transfer results and the data is about 9.5%. Using CFD and Wind Tunnel (WT) data, an empirical correlation for estimating heating augmentation on short hypersonic protuberances (k/delta less than 0.3) is proposed. This proposed correlation is compared with several computed flight simulation cases and good agreement is achieved. Accordingly, this correlation is proposed for further investigation on other short hypersonic protuberances for estimating heating augmentation.

Local and System Level Considerations for Plasma-Based Techniques in Hypersonic Flight

Science.gov (United States)

Suchomel, Charles; Gaitonde, Datta

2007-01-01

The harsh environment encountered due to hypersonic flight, particularly when air-breathing propulsion devices are utilized, poses daunting challenges to successful maturation of suitable technologies. This has spurred the quest for revolutionary solutions, particularly those exploiting the fact that air under these conditions can become electrically conducting either naturally or through artificial enhancement. Optimized development of such concepts must emphasize not only the detailed physics by which the fluid interacts with the imposed electromagnetic fields, but must also simultaneously identify system level issues integration and efficiencies that provide the greatest leverage. This paper presents some recent advances at both levels. At the system level, an analysis is summarized that incorporates the interdependencies occurring between weight, power and flow field performance improvements. Cruise performance comparisons highlight how one drag reduction device interacts with the vehicle to improve range. Quantified parameter interactions allow specification of system requirements and energy consuming technologies that affect overall flight vehicle performance. Results based on on the fundamental physics are presented by distilling numerous computational studies into a few guiding principles. These highlight the complex non-intuitive relationships between the various fluid and electromagnetic fields, together with thermodynamic considerations. Generally, energy extraction is an efficient process, while the reverse is accompanied by significant dissipative heating and inefficiency. Velocity distortions can be detrimental to plasma operation, but can be exploited to tailor flows through innovative electromagnetic configurations.

Prediction of forces and moments for flight vehicle control effectors. Part 1: Validation of methods for predicting hypersonic vehicle controls forces and moments

Science.gov (United States)

Maughmer, Mark D.; Ozoroski, L.; Ozoroski, T.; Straussfogel, D.

1990-01-01

Many types of hypersonic aircraft configurations are currently being studied for feasibility of future development. Since the control of the hypersonic configurations throughout the speed range has a major impact on acceptable designs, it must be considered in the conceptual design stage. The ability of the aerodynamic analysis methods contained in an industry standard conceptual design system, APAS II, to estimate the forces and moments generated through control surface deflections from low subsonic to high hypersonic speeds is considered. Predicted control forces and moments generated by various control effectors are compared with previously published wind tunnel and flight test data for three configurations: the North American X-15, the Space Shuttle Orbiter, and a hypersonic research airplane concept. Qualitative summaries of the results are given for each longitudinal force and moment and each control derivative in the various speed ranges. Results show that all predictions of longitudinal stability and control derivatives are acceptable for use at the conceptual design stage. Results for most lateral/directional control derivatives are acceptable for conceptual design purposes; however, predictions at supersonic Mach numbers for the change in yawing moment due to aileron deflection and the change in rolling moment due to rudder deflection are found to be unacceptable. Including shielding effects in the analysis is shown to have little effect on lift and pitching moment predictions while improving drag predictions.

Computations of Viking Lander Capsule Hypersonic Aerodynamics with Comparisons to Ground and Flight Data

Science.gov (United States)

Edquist, Karl T.

2006-01-01

Comparisons are made between the LAURA Navier-Stokes code and Viking Lander Capsule hypersonic aerodynamics data from ground and flight measurements. Wind tunnel data are available for a 3.48 percent scale model at Mach 6 and a 2.75 percent scale model at Mach 10.35, both under perfect gas air conditions. Viking Lander 1 aerodynamics flight data also exist from on-board instrumentation for velocities between 2900 and 4400 m/sec (Mach 14 to 23.3). LAURA flowfield solutions are obtained for the geometry as tested or flown, including sting effects at tunnel conditions and finite-rate chemistry effects in flight. Using the flight vehicle center-of-gravity location (trim angle approx. equals -11.1 deg), the computed trim angle at tunnel conditions is within 0.31 degrees of the angle derived from Mach 6 data and 0.13 degrees from the Mach 10.35 trim angle. LAURA Mach 6 trim lift and drag force coefficients are within 2 percent of measured data, and computed trim lift-to-drag ratio is within 4 percent of the data. Computed trim lift and drag force coefficients at Mach 10.35 are within 5 percent and 3 percent, respectively, of wind tunnel data. Computed trim lift-to-drag ratio is within 2 percent of the Mach 10.35 data. Using the nominal density profile and center-of-gravity location, LAURA trim angle at flight conditions is within 0.5 degrees of the total angle measured from on-board instrumentation. LAURA trim lift and drag force coefficients at flight conditions are within 7 and 5 percent, respectively, of the flight data. Computed trim lift-to-drag ratio is within 4 percent of the data. Computed aerodynamics sensitivities to center-of-gravity location, atmospheric density, and grid refinement are generally small. The results will enable a better estimate of aerodynamics uncertainties for future Mars entry vehicles where non-zero angle-of-attack is required.

A 3D imaging system for the non-intrusive in-flight measurement of the deformation of an aircraft propeller and a helicopter rotor

Science.gov (United States)

Stasicki, Bolesław; Boden, Fritz; Ludwikowski, Krzysztof

2017-02-01

The non-intrusive in-flight deformation measurement and the resulting local pitch of an aircraft propeller or helicopter rotor blade is a demanding task. The idea of an imaging system integrated and rotating with the air-craft propeller has already been presented at the 30th International Congress on High-Speed Imaging and Photonics (ICHSIP30) in 2012. Since then this system has been designed, constructed and tested in the laboratory as well as in-flight on the Cobra VUT100 of Evektor Aerotechnik, Kunovice (CZ). The major aim of the EU FP7 project AIM2 ("Advanced In-flight Measurement techniques 2" - contract No. 266107) was to ascertain the feasibility of this technique under extreme conditions - vibration and large centrifugal forces - to real flight testing. Based on the gained experience a new rotating system for the application on helicopter rotors has recently been constructed and tested on the whirl tower of Airbus Helicopters, Donauwoerth (D). In this paper the principle of the applied Image Pattern Correlation Technique (IPCT), a specialized type of Digital Image Correlation (DIC), is outlined and the construction of both rotating 3D image acquisition systems dedicated to the in-flight deformation measurement of the aircraft propeller and helicopter rotor are described. Furthermore, the results of the ground and in-flight tests of these systems will be shown and discussed. The obtained results will be helpful for manufacturers in the design of their future aircrafts.

Uncertainty Propagation in Hypersonic Vehicle Aerothermoelastic Analysis

Science.gov (United States)

Lamorte, Nicolas Etienne

Hypersonic vehicles face a challenging flight environment. The aerothermoelastic analysis of its components requires numerous simplifying approximations. Identifying and quantifying the effect of uncertainties pushes the limits of the existing deterministic models, and is pursued in this work. An uncertainty quantification framework is used to propagate the effects of identified uncertainties on the stability margins and performance of the different systems considered. First, the aeroelastic stability of a typical section representative of a control surface on a hypersonic vehicle is examined. Variability in the uncoupled natural frequencies of the system is modeled to mimic the effect of aerodynamic heating. Next, the stability of an aerodynamically heated panel representing a component of the skin of a generic hypersonic vehicle is considered. Uncertainty in the location of transition from laminar to turbulent flow and the heat flux prediction is quantified using CFD. In both cases significant reductions of the stability margins are observed. A loosely coupled airframe--integrated scramjet engine is considered next. The elongated body and cowl of the engine flow path are subject to harsh aerothermodynamic loading which causes it to deform. Uncertainty associated with deformation prediction is propagated to the engine performance analysis. The cowl deformation is the main contributor to the sensitivity of the propulsion system performance. Finally, a framework for aerothermoelastic stability boundary calculation for hypersonic vehicles using CFD is developed. The usage of CFD enables one to consider different turbulence conditions, laminar or turbulent, and different models of the air mixture, in particular real gas model which accounts for dissociation of molecules at high temperature. The system is found to be sensitive to turbulence modeling as well as the location of the transition from laminar to turbulent flow. Real gas effects play a minor role in the

The NASA-sponsored Maryland center for hypersonic education and research

Science.gov (United States)

Lewis, Mark J.; Gupta, Ashwani K.

1995-01-01

The Office of Aeronautics of the National Aeronautics and Space Administration has established a program to support university programs in the field of hypersonic flight. Beginning in the fall of 1993, three universities, including the University of Maryland at College Park, were selected to participate in this activity. The program at the University of Maryland includes faculty in the Department of Aerospace Engineering and Department of Mechanical Engineering, and provides a multidisciplinary environment for graduate and undergraduate students to study and conduct research in the field of hypersonic flight. Ongoing projects cover the range of applications from cruisers through transatmospheric and reentry vehicles. Research activities, focused on propulsion, fluid dynamics, inverse design, and vehicle optimization and integration, are conducted in conjuntion with industrial partners and government laboratories.

HIFiRE-5 Flight Test Preliminary Results (Postprint)

Science.gov (United States)

2013-11-01

CFD . 15. SUBJECT TERMS Boundary layer transition, hypersonic , flight test 16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF ABSTRACT: SAR 18...consistent with prior CFD 33,34 and wind tunnel measurements 28,29 at hypersonic conditions that indicated that the centerline is more unstable... Hypersonic Sciences Branch High Speed Systems Division NOVEMBER 2013 Approved for public release; distribution unlimited

Mars Ascent Vehicle-Propellant Aging

Science.gov (United States)

Dankanich, John; Rousseau, Jeremy; Williams, Jacob

2015-01-01

This project is to develop and test a new propellant formulation specifically for the Mars Ascent Vehicle (MAV) for the robotic Mars Sample Return mission. The project was initiated under the Planetary Sciences Division In-Space Propulsion Technology (ISPT) program and is continuing under the Mars Exploration Program. The two-stage, solid motor-based MAV has been the leading MAV solution for more than a decade. Additional studies show promise for alternative technologies including hybrid and bipropellant options, but the solid motor design has significant propellant density advantages well suited for physical constraints imposed while using the SkyCrane descent stage. The solid motor concept has lower specific impulse (Isp) than alternatives, but if the first stage and payload remain sufficiently small, the two-stage solid MAV represents a potential low risk approach to meet the mission needs. As the need date for the MAV slips, opportunities exist to advance technology with high on-ramp potential. The baseline propellant for the MAV is currently the carboxyl terminated polybutadiene (CTPB) based formulation TP-H-3062 due to its advantageous low temperature mechanical properties and flight heritage. However, the flight heritage is limited and outside the environments, the MAV must endure. The ISPT program competed a propellant formulation project with industry and selected ATK to develop a new propellant formulation specifically for the MAV application. Working with ATK, a large number of propellant formulations were assessed to either increase performance of a CTPB propellant or improve the low temperature mechanical properties of a hydroxyl terminated polybutadiene (HTPB) propellant. Both propellants demonstrated potential to increase performance over heritage options, but an HTPB propellant formulation, TP-H-3544, was selected for production and testing. The test plan includes propellant aging first at high vacuum conditions, representative of the Mars transit

Analytical and computational investigations of a magnetohydrodynamics (MHD) energy-bypass system for supersonic gas turbine engines to enable hypersonic flight

Science.gov (United States)

Benyo, Theresa Louise

Historically, the National Aeronautics and Space Administration (NASA) has used rocket-powered vehicles as launch vehicles for access to space. A familiar example is the Space Shuttle launch system. These vehicles carry both fuel and oxidizer onboard. If an external oxidizer (such as the Earth's atmosphere) is utilized, the need to carry an onboard oxidizer is eliminated, and future launch vehicles could carry a larger payload into orbit at a fraction of the total fuel expenditure. For this reason, NASA is currently researching the use of air-breathing engines to power the first stage of two-stage-to-orbit hypersonic launch systems. Removing the need to carry an onboard oxidizer leads also to reductions in total vehicle weight at liftoff. This in turn reduces the total mass of propellant required, and thus decreases the cost of carrying a specific payload into orbit or beyond. However, achieving hypersonic flight with air-breathing jet engines has several technical challenges. These challenges, such as the mode transition from supersonic to hypersonic engine operation, are under study in NASA's Fundamental Aeronautics Program. One propulsion concept that is being explored is a magnetohydrodynamic (MHD) energy- bypass generator coupled with an off-the-shelf turbojet/turbofan. It is anticipated that this engine will be capable of operation from takeoff to Mach 7 in a single flowpath without mode transition. The MHD energy bypass consists of an MHD generator placed directly upstream of the engine, and converts a portion of the enthalpy of the inlet flow through the engine into electrical current. This reduction in flow enthalpy corresponds to a reduced Mach number at the turbojet inlet so that the engine stays within its design constraints. Furthermore, the generated electrical current may then be used to power aircraft systems or an MHD accelerator positioned downstream of the turbojet. The MHD accelerator operates in reverse of the MHD generator, re-accelerating the

Multiple-step fault estimation for interval type-II T-S fuzzy system of hypersonic vehicle with time-varying elevator faults

Directory of Open Access Journals (Sweden)

Jin Wang

2017-03-01

Full Text Available This article proposes a multiple-step fault estimation algorithm for hypersonic flight vehicles that uses an interval type-II Takagi–Sugeno fuzzy model. An interval type-II Takagi–Sugeno fuzzy model is developed to approximate the nonlinear dynamic system and handle the parameter uncertainties of hypersonic firstly. Then, a multiple-step time-varying additive fault estimation algorithm is designed to estimate time-varying additive elevator fault of hypersonic flight vehicles. Finally, the simulation is conducted in both aspects of modeling and fault estimation; the validity and availability of such method are verified by a series of the comparison of numerical simulation results.

Applications of an implicit HLLC-based Godunov solver for steady state hypersonic problems

International Nuclear Information System (INIS)

Link, R.A.; Sharman, B.

2005-01-01

Over the past few years, there has been considerable activity developing research vehicles for studying hypersonic propulsion. Successful launches of the Australian Hyshot and the US Hyper-X vehicles have added a significant amount of flight test data to a field that had previously been limited to numerical simulation. A number of approaches have been proposed for hypersonics propulsion, including attached detonation wave, supersonics combustion, and shock induced combustion. Due to the high cost of developing flight hardware, CFD simulations will continue to be a key tool for investigating the feasibility of these concepts. Capturing the interactions of the vehicle body with the boundary layer and chemical reactions pushes the limits of available modelling tools and computer hardware. Explicit formulations are extremely slow in converging to a steady state; therefore, the use of implicit methods are warranted. An implicit LLC-based Godunov solver has been developed at Martec in collaboration with DRDC Valcartier to solve hypersonic problems with a minimum of CPU time and RAM storage. The solver, Chinook Implicit, is based upon the implicit formulation adopted by Batten et. al. The solver is based on a point implicit Gauss-Seidel method for unstructured grids, and includes fully implicit boundary conditions. Preliminary results for small and large scale inviscid hypersonics problems will be presented. (author)

Modeling and Analysis of an Air-Breathing Flexible Hypersonic Vehicle

Directory of Open Access Journals (Sweden)

Xi-bin Zhang

2014-01-01

Full Text Available By using light-weighted material in hypersonic vehicle, the vehicle body can be easily deformed. The mutual couplings in aerodynamics, flexible structure, and propulsion system will bring great challenges for vehicle modeling. In this work, engineering estimated method is used to calculate the aerodynamic forces, moments, and flexible modes to get the physics-based model of an air-breathing flexible hypersonic vehicle. The model, which contains flexible effects and viscous effects, can capture the physical characteristics of high-speed flight. To overcome the analytical intractability of the model, a simplified control-oriented model of the hypersonic vehicle is presented with curve fitting approximations. The control-oriented model can not only reduce the complexity of the model, but also retain aero-flexible structure-propulsion interactions of the physics-based model and can be applied for nonlinear control.

Global neural dynamic surface tracking control of strict-feedback systems with application to hypersonic flight vehicle.

Science.gov (United States)

Xu, Bin; Yang, Chenguang; Pan, Yongping

2015-10-01

This paper studies both indirect and direct global neural control of strict-feedback systems in the presence of unknown dynamics, using the dynamic surface control (DSC) technique in a novel manner. A new switching mechanism is designed to combine an adaptive neural controller in the neural approximation domain, together with the robust controller that pulls the transient states back into the neural approximation domain from the outside. In comparison with the conventional control techniques, which could only achieve semiglobally uniformly ultimately bounded stability, the proposed control scheme guarantees all the signals in the closed-loop system are globally uniformly ultimately bounded, such that the conventional constraints on initial conditions of the neural control system can be relaxed. The simulation studies of hypersonic flight vehicle (HFV) are performed to demonstrate the effectiveness of the proposed global neural DSC design.

Terahertz Quantum Cascade Laser-Based Sensors for Hypersonic Flows (7274-050), Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Ground test facilities are used by NASA to simulate the conditions present during flight at hypersonic velocities, to test thermal protection materials for existing...

The Hypersonic Revolution. Case Studies in the History of Hypersonic Technology. Volume III: The Quest for the Orbital Jet: The National Aero-Space Plane Program (1983-1995)

National Research Council Canada - National Science Library

Schwelkart, Larry

1998-01-01

... that could fly fast enough to attain orbital velocity, is considered a success by many of the participants.1 They contend that by "showing up," NASP survived long enough to produce what many deem critical technologies for hypersonic flight...

Hyper-X and Pegasus Launch Vehicle: A Three-Foot Model of the Hypersonic Experimental Research Vehic

Science.gov (United States)

1997-01-01

The configuration of the X-43A Hypersonic Experimental Research Vehicle, or Hyper-X, attached to a Pegasus launch vehicle is displayed in this three-foot-long model at NASA's Dryden Flight Research Center, Edwards, California. Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43

Flow-Tagging Velocimetry for Hypersonic Flows Using Fluorescence of Nitric Oxide

Science.gov (United States)

Danehy, P. M.; OByrne, S.; Houwing, A. F. P.

2001-01-01

We investigate a new type of flow-tagging velocimetry technique for hypersonic flows. The technique involves exciting a thin line of nitric oxide molecules with a laser beam and then, after some delay, acquiring an image of the displaced line. One component of velocity is determined from the time of flight. This method is applied to measure the velocity profile in a Mach 8.5 laminar, hypersonic boundary layer in the Australian National Universities T2 free-piston shock tunnel. The velocity is measured with an uncertainty of approximately 2%. Comparison with a CFD simulation of the flow shows reasonable agreement.

Numerical analysis of exhaust jet secondary combustion in hypersonic flow field

Science.gov (United States)

Yang, Tian-Peng; Wang, Jiang-Feng; Zhao, Fa-Ming; Fan, Xiao-Feng; Wang, Yu-Han

2018-05-01

The interaction effect between jet and control surface in supersonic and hypersonic flow is one of the key problems for advanced flight control system. The flow properties of exhaust jet secondary combustion in a hypersonic compression ramp flow field were studied numerically by solving the Navier-Stokes equations with multi-species and combustion reaction effects. The analysis was focused on the flow field structure and the force amplification factor under different jet conditions. Numerical results show that a series of different secondary combustion makes the flow field structure change regularly, and the temperature increases rapidly near the jet exit.

A review of research in low earth orbit propellant collection

Science.gov (United States)

Singh, Lake A.; Walker, Mitchell L. R.

2015-05-01

This comprehensive review examines the efforts of previous researchers to develop concepts for propellant-collecting spacecraft, estimate the performance of these systems, and understand the physics involved. Rocket propulsion requires the spacecraft to expend two fundamental quantities: energy and propellant mass. A growing number of spacecraft collect the energy they need to execute propulsive maneuvers in-situ with solar panels. In contrast, every spacecraft using rocket propulsion has carried all of the propellant mass needed for the mission from the ground, which limits the range and mission capabilities. Numerous researchers have explored the concept of collecting propellant mass while in space. These concepts have varied in scale and complexity from chemical ramjets to fusion-driven interstellar vessels. Research into propellant-collecting concepts occurred in distinct eras. During the Cold War, concepts tended to be large, complex, and nuclear powered. After the Cold War, concepts transitioned to solar power sources and more effort has been devoted to detailed analysis of specific components of the propellant-collecting architecture. By detailing the major contributions and limitations of previous work, this review concisely presents the state-of-the-art and outlines five areas for continued research. These areas include air-compatible cathode technology, techniques to improve propellant utilization on atmospheric species, in-space compressor and liquefaction technology, improved hypersonic and hyperthermal free molecular flow inlet designs, and improved understanding of how design parameters affect system performance.

Flight Experiment Verification of Shuttle Boundary Layer Transition Prediction Tool

Science.gov (United States)

Berry, Scott A.; Berger, Karen T.; Horvath, Thomas J.; Wood, William A.

2016-01-01

Boundary layer transition at hypersonic conditions is critical to the design of future high-speed aircraft and spacecraft. Accurate methods to predict transition would directly impact the aerothermodynamic environments used to size a hypersonic vehicle's thermal protection system. A transition prediction tool, based on wind tunnel derived discrete roughness correlations, was developed and implemented for the Space Shuttle return-to-flight program. This tool was also used to design a boundary layer transition flight experiment in order to assess correlation uncertainties, particularly with regard to high Mach-number transition and tunnel-to-flight scaling. A review is provided of the results obtained from the flight experiment in order to evaluate the transition prediction tool implemented for the Shuttle program.

Development and implementation of a propeller test capability for GL-10 "Greased Lightning" propeller design

Science.gov (United States)

Duvall, Brian Edward

Interest in small unmanned aerial vehicles has increased dramatically in recent years. Hybrid vehicles which allow forward flight as a fixed wing aircraft and a true vertical landing capability have always had applications. Management of the available energy and noise associated with electric propeller propulsion systems presents many challenges. NASA Langley has developed the Greased Lightning 10 (GL-10) vertical takeoff, unmanned aerial vehicle with ten individual motors and propellers. All are used for propulsion during takeoff and contribute to acoustic noise pollution which is an identified nuisance to the surrounding users. A propeller test capability was developed to gain an understanding of how the noise can be reduced while meeting minimum thrust requirements. The designed propeller test stand allowed for various commercially available propellers to be tested for potential direct replacement of the current GL-10 propellers and also supported testing of a newly designed propeller provided by the Georgia Institute of Technology. Results from the test program provided insight as to which factors affect the noise as well as performance characteristics. The outcome of the research effort showed that the current GL-10 propeller still represents the best choice of all the candidate propellers tested.

High-speed Imaging of Global Surface Temperature Distributions on Hypersonic Ballistic-Range Projectiles

Science.gov (United States)

Wilder, Michael C.; Reda, Daniel C.

2004-01-01

The NASA-Ames ballistic range provides a unique capability for aerothermodynamic testing of configurations in hypersonic, real-gas, free-flight environments. The facility can closely simulate conditions at any point along practically any trajectory of interest experienced by a spacecraft entering an atmosphere. Sub-scale models of blunt atmospheric entry vehicles are accelerated by a two-stage light-gas gun to speeds as high as 20 times the speed of sound to fly ballistic trajectories through an 24 m long vacuum-rated test section. The test-section pressure (effective altitude), the launch velocity of the model (flight Mach number), and the test-section working gas (planetary atmosphere) are independently variable. The model travels at hypersonic speeds through a quiescent test gas, creating a strong bow-shock wave and real-gas effects that closely match conditions achieved during actual atmospheric entry. The challenge with ballistic range experiments is to obtain quantitative surface measurements from a model traveling at hypersonic speeds. The models are relatively small (less than 3.8 cm in diameter), which limits the spatial resolution possible with surface mounted sensors. Furthermore, since the model is in flight, surface-mounted sensors require some form of on-board telemetry, which must survive the massive acceleration loads experienced during launch (up to 500,000 gravities). Finally, the model and any on-board instrumentation will be destroyed at the terminal wall of the range. For these reasons, optical measurement techniques are the most practical means of acquiring data. High-speed thermal imaging has been employed in the Ames ballistic range to measure global surface temperature distributions and to visualize the onset of transition to turbulent-flow on the forward regions of hypersonic blunt bodies. Both visible wavelength and infrared high-speed cameras are in use. The visible wavelength cameras are intensified CCD imagers capable of integration

Feasibility study of a nonequilibrium MHD accelerator concept for hypersonic propulsion ground testing

International Nuclear Information System (INIS)

Lee, Ying-Ming; Simmons, G.A.; Nelson, G.L.

1995-01-01

A National Aeronautics and Space Administration (NASA) funded research study to evaluate the feasibility of using magnetohydrodynamic (MHD) body force accelerators to produce true air simulation for hypersonic propulsion ground testing is discussed in this paper. Testing over the airbreathing portion of a transatmospheric vehicle (TAV) hypersonic flight regime will require high quality air simulation for actual flight conditions behind a bow shock wave (forebody, pre-inlet region) for flight velocities up to Mach 16 and perhaps beyond. Material limits and chemical dissociation at high temperature limit the simulated flight Mach numbers in conventional facilities to less than Mach 12 for continuous and semi-continuous testing and less than Mach 7 for applications requiring true air chemistry. By adding kinetic energy directly to the flow, MHD accelerators avoid the high temperatures and pressures required in the reservoir region of conventional expansion facilities, allowing MHD to produce true flight conditions in flight regimes impossible with conventional facilities. The present study is intended to resolve some of the critical technical issues related to the operation of MHD at high pressure. Funding has been provided only for the first phase of a three to four year feasibility study that would culminate in the demonstration of MHD acceleration under conditions required to produce true flight conditions behind a bow shock wave to flight Mach numbers of 16 or greater. MHD critical issues and a program plan to resolve these are discussed

Flight Control Laws for NASA's Hyper-X Research Vehicle

Science.gov (United States)

Davidson, J.; Lallman, F.; McMinn, J. D.; Martin, J.; Pahle, J.; Stephenson, M.; Selmon, J.; Bose, D.

1999-01-01

The goal of the Hyper-X program is to demonstrate and validate technology for design and performance predictions of hypersonic aircraft with an airframe-integrated supersonic-combustion ramjet propulsion system. Accomplishing this goal requires flight demonstration of a hydrogen-fueled scramjet powered hypersonic aircraft. A key enabling technology for this flight demonstration is flight controls. Closed-loop flight control is required to enable a successful stage separation, to achieve and maintain the design condition during the engine test, and to provide a controlled descent. Before the contract award, NASA developed preliminary flight control laws for the Hyper-X to evaluate the feasibility of the proposed scramjet test sequence and descent trajectory. After the contract award, a Boeing/NASA partnership worked to develop the current control laws. This paper presents a description of the Hyper-X Research Vehicle control law architectures with performance and robustness analyses. Assessments of simulated flight trajectories and stability margin analyses demonstrate that these control laws meet the flight test requirements.

An Overview of the NASA FAP Hypersonics Project Airbreathing Propulsion Research

Science.gov (United States)

Auslender, A. H.; Suder, Kenneth L.; Thomas, Scott R.

2009-01-01

The propulsion research portfolio of the National Aeronautics and Space Administration Fundamental Aeronautics Program Hypersonics Project encompasses a significant number of technical tasks that are aligned to achieve mastery and intellectual stewardship of the core competencies in the hypersonic-flight regime. An overall coordinated programmatic and technical effort has been structured to advance the state-of-the-art, via both experimental and analytical efforts. A subset of the entire hypersonics propulsion research portfolio is presented in this overview paper. To this end, two programmatic research disciplines are discussed; namely, (1) the Propulsion Discipline, including three associated research elements: the X-51A partnership, the HIFiRE-2 partnership, and the Durable Combustor Rig, and (2) the Turbine-Based Combine Cycle Discipline, including three associated research elements: the Combined Cycle Engine Large Scale Inlet Mode Transition Experiment, the small-scale Inlet Mode Transition Experiment, and the High-Mach Fan Rig.

Hypersonic wind-tunnel free-flying experiments with onboard instrumentation

KAUST Repository

Mudford, Neil R.; O'Byrne, Sean B.; Neely, Andrew J.; Buttsworth, David R.; Balage, Sudantha

2015-01-01

Hypersonic wind-tunnel testing with "free-flight" models unconnected to a sting ensures that sting/wake flow interactions do not compromise aerodynamic coefficient measurements. The development of miniaturized electronics has allowed the demonstration of a variant of a new method for the acquisition of hypersonic model motion data using onboard accelerometers, gyroscopes, and a microcontroller. This method is demonstrated in a Mach 6 wind-tunnel flow, whose duration and pitot pressure are sufficient for the model to move a body length or more and turn through a significant angle. The results are compared with those obtained from video analysis of the model motion, the existing method favored for obtaining aerodynamic coefficients in similar hypersonic wind-tunnel facilities. The results from the two methods are in good agreement. The new method shows considerable promise for reliable measurement of aerodynamic coefficients, particularly because the data obtained are in more directly applicable forms of accelerations and rates of turn, rather than the model position and attitude obtained from the earlier visualization method. The ideal may be to have both methods operating together.

Three-Dimensional Aeroelastic and Aerothermoelastic Behavior in Hypersonic Flow

Science.gov (United States)

McNamara, Jack J.; Friedmann, Peretz P.; Powell, Kenneth G.; Thuruthimattam, Biju J.; Bartels, Robert E.

2005-01-01

The aeroelastic and aerothermoelastic behavior of three-dimensional configurations in hypersonic flow regime are studied. The aeroelastic behavior of a low aspect ratio wing, representative of a fin or control surface on a generic hypersonic vehicle, is examined using third order piston theory, Euler and Navier-Stokes aerodynamics. The sensitivity of the aeroelastic behavior generated using Euler and Navier-Stokes aerodynamics to parameters governing temporal accuracy is also examined. Also, a refined aerothermoelastic model, which incorporates the heat transfer between the fluid and structure using CFD generated aerodynamic heating, is used to examine the aerothermoelastic behavior of the low aspect ratio wing in the hypersonic regime. Finally, the hypersonic aeroelastic behavior of a generic hypersonic vehicle with a lifting-body type fuselage and canted fins is studied using piston theory and Euler aerodynamics for the range of 2.5 less than or equal to M less than or equal to 28, at altitudes ranging from 10,000 feet to 80,000 feet. This analysis includes a study on optimal mesh selection for use with Euler aerodynamics. In addition to the aeroelastic and aerothermoelastic results presented, three time domain flutter identification techniques are compared, namely the moving block approach, the least squares curve fitting method, and a system identification technique using an Auto-Regressive model of the aeroelastic system. In general, the three methods agree well. The system identification technique, however, provided quick damping and frequency estimations with minimal response record length, and therefore o ers significant reductions in computational cost. In the present case, the computational cost was reduced by 75%. The aeroelastic and aerothermoelastic results presented illustrate the applicability of the CFL3D code for the hypersonic flight regime.

Trajectory Optimization and Conceptual Study of Small Test Vehicles for Hypersonic Engine Using High-Altitude Balloon

Science.gov (United States)

Tsuchiya, Takeshi; Takenaka, Youichi; Taguchi, Hideyuki; Sawai, Shujiro

Japan Aerospace Exploration Agency, JAXA announced a long-term vision recently. In the vision, JAXA aims to develop hypersonic aircrafts. A pre-cooled turbojet engine has great potential as one of newly developed hypersonic air-breathing engines. We also expect the engine to be installed in space transportation vehicles in future. For combustion test in real flight condition of the engines, JAXA has an experimental plan with a small test vehicle falling from a high-altitude balloon. This paper applies numerical analysis and optimization techniques to conceptual designs of the test vehicle in order to obtain the best configuration and trajectory that can achieve the flight test. The results show helpful knowledge when we design prototype vehicles.

Cryogenic Propellant Storage and Transfer

Data.gov (United States)

National Aeronautics and Space Administration — Space Flight Demonstration development has been canceled in favor of a ground test bed development for of passive/active cryogenic propellant storage, transfer, and...

Concept and performance study of turbocharged solid propellant ramjet

Science.gov (United States)

Li, Jiang; Liu, Kai; Liu, Yang; Liu, Shichang

2018-06-01

This study proposes a turbocharged solid propellant ramjet (TSPR) propulsion system that integrates a turbocharged system consisting of a solid propellant (SP) air turbo rocket (ATR) and the fuel-rich gas generator of a solid propellant ramjet (SPR). First, a suitable propellant scheme was determined for the TSPR. A solid hydrocarbon propellant is used to generate gas for driving the turbine, and a boron-based fuel-rich propellant is used to provide fuel-rich gas to the afterburner. An appropriate TSPR structure was also determined. The TSPR's thermodynamic cycle was analysed to prove its theoretical feasibility. The results showed that the TSPR's specific cycle power was larger than those of SP-ATR and SPR and thermal efficiency was slightly less than that of SP-ATR. Overall, TSPR showed optimal performance in a wide flight envelope. The specific impulses and specific thrusts of TSPR, SP-ATR, and SPR in the flight envelope were calculated and compared. TSPR's flight envelope roughly overlapped that of SP-ATR, its specific impulse was larger than that of SP-ATR, and its specific thrust was larger than those of SP-ATR and SPR. Attempts to improve the TSPR off-design performance prompted our proposal of a control plan for off-design codes in which both the turbocharger corrected speed and combustor excess gas coefficient are kept constant. An off-design performance model was established by analysing the TSPR working process. We concluded that TSPR with a constant corrected speed had wider flight envelope, higher thrust, and higher specific impulse than TSPR with a constant physical speed determined by calculating the performance of off-design TSPR codes under different control plans. The results of this study can provide a reference for further studies on TSPRs.

Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Technology Development Overview

Science.gov (United States)

Hughes, Stephen J.; Cheatwood, F. McNeil; Calomino, Anthony M.; Wright, Henry S.; Wusk, Mary E.; Hughes, Monica F.

2013-01-01

The successful flight of the Inflatable Reentry Vehicle Experiment (IRVE)-3 has further demonstrated the potential value of Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology. This technology development effort is funded by NASA's Space Technology Mission Directorate (STMD) Game Changing Development Program (GCDP). This paper provides an overview of a multi-year HIAD technology development effort, detailing the projects completed to date and the additional testing planned for the future.

A Low Cost Approach to the Design of Autopilot for Hypersonic Glider

Science.gov (United States)

Liang, Wang; Weihua, Zhang; Ke, Peng; Donghui, Wang

2017-12-01

This paper proposes a novel integrated guidance and control (IGC) approach to improve the autopilot design with low cost for hypersonic glider in dive and pull-up phase. The main objective is robust and adaptive tracking of flight path angle (FPA) under severe flight scenarios. Firstly, the nonlinear IGC model is developed with a second order actuator dynamics. Then the adaptive command filtered back-stepping control is implemented to deal with the large aerodynamics coefficient uncertainties, control surface uncertainties and unmatched time-varying disturbances. For the autopilot, a back-stepping sliding mode control is designed to track the control surface deflection, and a nonlinear differentiator is used to avoid direct differentiating the control input. Through a series of 6-DOF numerical simulations, it’s shown that the proposed scheme successfully cancels out the large uncertainties and disturbances in tracking different kinds of FPA trajectory. The contribution of this paper lies in the application and determination of nonlinear integrated design of guidance and control system for hypersonic glider.

A matching approach to communicate through the plasma sheath surrounding a hypersonic vehicle

International Nuclear Information System (INIS)

Gao, Xiaotian; Jiang, Binhao

2015-01-01

In order to overcome the communication blackout problem suffered by hypersonic vehicles, a matching approach has been proposed for the first time in this paper. It utilizes a double-positive (DPS) material layer surrounding a hypersonic vehicle antenna to match with the plasma sheath enclosing the vehicle. Analytical analysis and numerical results indicate a resonance between the matched layer and the plasma sheath will be formed to mitigate the blackout problem in some conditions. The calculated results present a perfect radiated performance of the antenna, when the match is exactly built between these two layers. The effects of the parameters of the plasma sheath have been researched by numerical methods. Based on these results, the proposed approach is easier to realize and more flexible to the varying radiated conditions in hypersonic flight comparing with other methods

Hyper-X Research Vehicle - Artist Concept in Flight

Science.gov (United States)

1997-01-01

An artist's conception of the X-43A Hypersonic Experimental Vehicle, or 'Hyper-X' in flight. The X-43A was developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will

Advanced Design and Manufacture of Cryogenic Propellant Tanks for Air Launched Liquid Rockets

Data.gov (United States)

National Aeronautics and Space Administration — Generation Orbit (GO) is developing a sub-orbital system to enable rapid and inexpensive hypersonic flight regime test capabilities. To keep the cost of their launch...

Detailed modeling of electron emission for transpiration cooling of hypersonic vehicles

Science.gov (United States)

Hanquist, Kyle M.; Hara, Kentaro; Boyd, Iain D.

2017-02-01

Electron transpiration cooling (ETC) is a recently proposed approach to manage the high heating loads experienced at the sharp leading edges of hypersonic vehicles. Computational fluid dynamics (CFD) can be used to investigate the feasibility of ETC in a hypersonic environment. A modeling approach is presented for ETC, which includes developing the boundary conditions for electron emission from the surface, accounting for the space-charge limit effects of the near-wall plasma sheath. The space-charge limit models are assessed using 1D direct-kinetic plasma sheath simulations, taking into account the thermionically emitted electrons from the surface. The simulations agree well with the space-charge limit theory proposed by Takamura et al. for emitted electrons with a finite temperature, especially at low values of wall bias, which validates the use of the theoretical model for the hypersonic CFD code. The CFD code with the analytical sheath models is then used for a test case typical of a leading edge radius in a hypersonic flight environment. The CFD results show that ETC can lower the surface temperature of sharp leading edges of hypersonic vehicles, especially at higher velocities, due to the increase in ionized species enabling higher electron heat extraction from the surface. The CFD results also show that space-charge limit effects can limit the ETC reduction of surface temperatures, in comparison to thermionic emission assuming no effects of the electric field within the sheath.

Online Recorded Data-Based Composite Neural Control of Strict-Feedback Systems With Application to Hypersonic Flight Dynamics.

Science.gov (United States)

Xu, Bin; Yang, Daipeng; Shi, Zhongke; Pan, Yongping; Chen, Badong; Sun, Fuchun

2017-09-25

This paper investigates the online recorded data-based composite neural control of uncertain strict-feedback systems using the backstepping framework. In each step of the virtual control design, neural network (NN) is employed for uncertainty approximation. In previous works, most designs are directly toward system stability ignoring the fact how the NN is working as an approximator. In this paper, to enhance the learning ability, a novel prediction error signal is constructed to provide additional correction information for NN weight update using online recorded data. In this way, the neural approximation precision is highly improved, and the convergence speed can be faster. Furthermore, the sliding mode differentiator is employed to approximate the derivative of the virtual control signal, and thus, the complex analysis of the backstepping design can be avoided. The closed-loop stability is rigorously established, and the boundedness of the tracking error can be guaranteed. Through simulation of hypersonic flight dynamics, the proposed approach exhibits better tracking performance.

Robust Adaptive Flight Control Design of Air-breathing Hypersonic Vehicles

Science.gov (United States)

2016-12-07

advantages over rocket - based systems for space access vehicles. The major advantage of using air-breathing engine is that the extra oxidizer is not...sideslip angle (β) is calculated as Vt = p u2 + v2 +w2, α= t an−1 ( wu ), β= si n−1 ( vVt ) The rotational dynamic equations of AHV are given as Ṗ = c1QR...inverse controller for hypersonic vehicle. In 2010 International Conference on Information, Networking and Automation (ICINA), volume 2, pages V2 –240

Trajectory Optimization and Conceptual Study of Small Test Vehicles for a Hypersonic Engine Using a High-Altitude Balloon

Science.gov (United States)

Tsuchiya, Takeshi; Takenaka, Youichi; Taguchi, Hideyuki; Sawai, Shujiro

The Japan Aerospace Exploration Agency, JAXA, announced a long-term vision recently. In the vision, JAXA aims to develop hypersonic aircrafts. A pre-cooled turbojet engine has great potential as one of newly developed hypersonic airbreathing engines. We also expect the engine to be installed in space transportation vehicles in the future. For combustion test in the real flight conditions of the engines, JAXA has an experimental plan where a small test vehicle is released from a high-altitude balloon. This paper applies numerical analysis and optimization techniques to conceptual designs of the test vehicle in order to obtain the best configuration and trajectory for the flight test. The results show helpful knowledge for designing prototype vehicles.

The PROPEL Electrodynamic Tether Demonstration Mission

Science.gov (United States)

Bilen, Sven G.; Johnson, C. Les; Wiegmann, Bruce M.; Alexander, Leslie; Gilchrist, Brian E.; Hoyt, Robert P.; Elder, Craig H.; Fuhrhop, Keith P.; Scadera, Michael

2012-01-01

The PROPEL ("Propulsion using Electrodynamics") mission will demonstrate the operation of an electrodynamic tether propulsion system in low Earth orbit and advance its technology readiness level for multiple applications. The PROPEL mission has two primary objectives: first, to demonstrate the capability of electrodynamic tether technology to provide robust and safe, near-propellantless propulsion for orbit-raising, de-orbit, plane change, and station keeping, as well as to perform orbital power harvesting and formation flight; and, second, to fully characterize and validate the performance of an integrated electrodynamic tether propulsion system, qualifying it for infusion into future multiple satellite platforms and missions with minimal modification. This paper provides an overview of the PROPEL system and design reference missions; mission goals and required measurements; and ongoing PROPEL mission design efforts.

Hypersonic phononic crystals.

Science.gov (United States)

Gorishnyy, T; Ullal, C K; Maldovan, M; Fytas, G; Thomas, E L

2005-03-25

In this Letter we propose the use of hypersonic phononic crystals to control the emission and propagation of high frequency phonons. We report the fabrication of high quality, single crystalline hypersonic crystals using interference lithography and show that direct measurement of their phononic band structure is possible with Brillouin light scattering. Numerical calculations are employed to explain the nature of the observed propagation modes. This work lays the foundation for experimental studies of hypersonic crystals and, more generally, phonon-dependent processes in nanostructures.

Cryogenic Propellant Storage and Transfer Engineering Development Unit Hydrogen Tank

Science.gov (United States)

Werkheiser, Arthur

2015-01-01

The Cryogenic Propellant Storage and Transfer (CPST) project has been a long-running program in the Space Technology Mission Directorate to enhance the knowledge and technology related to handling cryogenic propellants, specifically liquid hydrogen. This particular effort, the CPST engineering development unit (EDU), was a proof of manufacturability effort in support of a flight article. The EDU was built to find and overcome issues related to manufacturability and collect data to anchor the thermal models for use on the flight design.

Hypersonic MHD Propulsion System Integration for the Mercury Lightcraft

International Nuclear Information System (INIS)

Myrabo, L.N.; Rosa, R.J.

2004-01-01

Introduced herein are the design, systems integration, and performance analysis of an exotic magnetohydrodynamic (MHD) slipstream accelerator engine for a single-occupant 'Mercury' lightcraft. This ultra-energetic, laser-boosted vehicle is designed to ride a 'tractor beam' into space, transmitted from a future orbital network of satellite solar power stations. The lightcraft's airbreathing combined-cycle engine employs a rotary pulsed detonation thruster mode for lift-off and landing, and an MHD slipstream accelerator mode at hypersonic speeds. The latter engine transforms the transatmospheric acceleration path into a virtual electromagnetic 'mass-driver' channel; the hypersonic momentum exchange process (with the atmosphere) enables engine specific impulses in the range of 6000 to 16,000 seconds, and propellant mass fractions as low as 10%. The single-stage-to-orbit, highly reusable lightcraft can accelerate at 3 Gs into low Earth orbit with its throttle just barely beyond 'idle' power, or virtually 'disappear' at 30 G's and beyond. The objective of this advanced lightcraft design is to lay the technological foundations for a safe, very low cost (e.g., 1000X below chemical rockets) air and space transportation for human life in the mid-21st Century - a system that will be completely 'green' and independent of Earth's limited fossil fuel reserves

Hypersonic MHD Propulsion System Integration for the Mercury Lightcraft

Science.gov (United States)

Myrabo, L. N.; Rosa, R. J.

2004-03-01

Introduced herein are the design, systems integration, and performance analysis of an exotic magnetohydrodynamic (MHD) slipstream accelerator engine for a single-occupant ``Mercury'' lightcraft. This ultra-energetic, laser-boosted vehicle is designed to ride a `tractor beam' into space, transmitted from a future orbital network of satellite solar power stations. The lightcraft's airbreathing combined-cycle engine employs a rotary pulsed detonation thruster mode for lift-off & landing, and an MHD slipstream accelerator mode at hypersonic speeds. The latter engine transforms the transatmospheric acceleration path into a virtual electromagnetic `mass-driver' channel; the hypersonic momentum exchange process (with the atmosphere) enables engine specific impulses in the range of 6000 to 16,000 seconds, and propellant mass fractions as low as 10%. The single-stage-to-orbit, highly reusable lightcraft can accelerate at 3 Gs into low Earth orbit with its throttle just barely beyond `idle' power, or virtually `disappear' at 30 G's and beyond. The objective of this advanced lightcraft design is to lay the technological foundations for a safe, very low cost (e.g., 1000X below chemical rockets) air and space transportation for human life in the mid-21st Century - a system that will be completely `green' and independent of Earth's limited fossil fuel reserves.

Airbreathing Hypersonic Systems Focus at NASA Langley Research Center

Science.gov (United States)

Hunt, James L.; Rausch, Vincent L.

1998-01-01

This paper presents the status of the airbreathing hypersonic airplane and space-access vehicle design matrix, reflects on the synergies and issues, and indicates the thrust of the effort to resolve the design matrix and to focus/advance systems technology maturation. Priority is given to the design of the vision operational vehicles followed by flow-down requirements to flight demonstrator vehicles and their design for eventual consideration in the Future-X Program.

Transition Experiments on Large Bluntness Cones with Distributed Roughness in Hypersonic Flight

Science.gov (United States)

Reda, Daniel. C.; Wilder, Michael C.; Prabhu, Dinesh K.

2012-01-01

Large bluntness cones with smooth nosetips and roughened frusta were flown in the NASA Ames hypersonic ballistic range at a Mach number of 10 through quiescent air environments. Global surface intensity (temperature) distributions were optically measured and analyzed to determine transition onset and progression over the roughened surface. Real-gas Navier-Stokes calculations of model flowfields, including laminar boundary layer development in these flowfields, were conducted to predict values of key dimensionless parameters used to correlate transition on such configurations in hypersonic flow. For these large bluntness cases, predicted axial distributions of the roughness Reynolds number showed (for each specified freestream pressure) that this parameter was a maximum at the physical beginning of the roughened zone and decreased with increasing run length along the roughened surface. Roughness-induced transition occurred downstream of this maximum roughness Reynolds number location, and progressed upstream towards the beginning of the roughened zone as freestream pressure was systematically increased. Roughness elements encountered at the upstream edge of the roughened frusta thus acted like a finite-extent trip array, consistent with published results concerning the tripping effectiveness of roughness bands placed on otherwise smooth surfaces.

Feasibility Study and Demonstration of an Aluminum and Ice Solid Propellant

Directory of Open Access Journals (Sweden)

Timothee L. Pourpoint

2012-01-01

Full Text Available Aluminum-water reactions have been proposed and studied for several decades for underwater propulsion systems and applications requiring hydrogen generation. Aluminum and water have also been proposed as a frozen propellant, and there have been proposals for other refrigerated propellants that could be mixed, frozen in situ, and used as solid propellants. However, little work has been done to determine the feasibility of these concepts. With the recent availability of nanoscale aluminum, a simple binary formulation with water is now feasible. Nanosized aluminum has a lower ignition temperature than micron-sized aluminum particles, partly due to its high surface area, and burning times are much faster than micron aluminum. Frozen nanoscale aluminum and water mixtures are stable, as well as insensitive to electrostatic discharge, impact, and shock. Here we report a study of the feasibility of an nAl-ice propellant in small-scale rocket experiments. The focus here is not to develop an optimized propellant; however improved formulations are possible. Several static motor experiments have been conducted, including using a flight-weight casing. The flight weight casing was used in the first sounding rocket test of an aluminum-ice propellant, establishing a proof of concept for simple propellant mixtures making use of nanoscale particles.

The screw propeller

Science.gov (United States)

Larrabee, E. E.

1980-07-01

Marine and air screw propellers are considered in terms of theoretical hydrodynamics as developed by Joukowsky, Prandtl, and Betz. Attention is given to the flow around wings of finite span where spanwise flow exists and where lift and the bound vorticity must all go smoothly to zero at the wing tips. The concept of a trailing vortex sheet made up of infinitesimal line vortexes roughly aligned with the direction of flight is discussed in this regard. Also considered is induced velocity, which tends to convect the sheet downward at every stage in the roll-up process, the vortex theory of propellers and the Betz-Prandtl circulation distribution. The performance of the Gossamer Albatross and of a pedal-driven biplane called the Chrysalis are also discussed.

Longitudinal aerodynamic characteristics of light, twin-engine, propeller-driven airplanes

Science.gov (United States)

Wolowicz, C. H.; Yancey, R. B.

1972-01-01

Representative state-of-the-art analytical procedures and design data for predicting the longitudinal static and dynamic stability and control characteristics of light, propeller-driven airplanes are presented. Procedures for predicting drag characteristics are also included. The procedures are applied to a twin-engine, propeller-driven airplane in the clean configuration from zero lift to stall conditions. The calculated characteristics are compared with wind-tunnel and flight data. Included in the comparisons are level-flight trim characteristics, period and damping of the short-period oscillatory mode, and windup-turn characteristics. All calculations are documented.

Pegasus(Registered trademark) Wing-Glove Experiment to Document Hypersonic Crossflow Transition: Measurement System and Selected Flight Results

Science.gov (United States)

Bertelrud, Arild; delaTova, Geva; Hamory, Philip J.; Young, Ronald; Noffz, Gregory K.; Dodson, Michael; Graves, Sharon S.; Diamond, John K.; Bartlett, James E.; Noack, Robert;

Validation of engineering methods for predicting hypersonic vehicle controls forces and moments

Science.gov (United States)

Maughmer, M.; Straussfogel, D.; Long, L.; Ozoroski, L.

1991-01-01

This work examines the ability of the aerodynamic analysis methods contained in an industry standard conceptual design code, the Aerodynamic Preliminary Analysis System (APAS II), to estimate the forces and moments generated through control surface deflections from low subsonic to high hypersonic speeds. Predicted control forces and moments generated by various control effectors are compared with previously published wind-tunnel and flight-test data for three vehicles: the North American X-15, a hypersonic research airplane concept, and the Space Shuttle Orbiter. Qualitative summaries of the results are given for each force and moment coefficient and each control derivative in the various speed ranges. Results show that all predictions of longitudinal stability and control derivatives are acceptable for use at the conceptual design stage.

Hypersonic rarefied-flow aerodynamics inferred from Shuttle Orbiter acceleration measurements

Science.gov (United States)

Blanchard, R. C.; Hinson, E. W.

1989-01-01

Data obtained from multiple flights of sensitive accelerometers on the Space Shuttle Orbiter during reentry have been used to develop an improved aerodynamic model for the Orbiter normal- and axial-force coefficients in hypersonic rarefied flow. The lack of simultaneous atmospheric density measurements was overcome in part by using the ratio of normal-to-axial acceleration, in which density cancels, as a constraint. Differences between the preflight model and the flight-acceleration-derived model in the continuum regime are attributed primarily to real gas effects. New insights are gained into the variation of the force coefficients in the transition between the continuum regime and free molecule flow.

Minimum weight passive insulation requirements for hypersonic cruise vehicles.

Science.gov (United States)

Ardema, M. D.

1972-01-01

Analytical solutions are derived for two representative cases of the transient heat conduction equation to determine the minimum weight requirements for passive insulation systems of hypersonic cruise vehicles. The cases discussed are the wet wall case with the interior wall temperature held to that of the boiling point of the fuel throughout the flight, and the dry wall case where the heat transferred through the insulation is absorbed by the interior structure whose temperature is allowed to rise.

Hypersonic ground test capabilities for T and E testing above mach 8 ''a case where S and T meets T and E''

International Nuclear Information System (INIS)

Constantino, M; Miles, R; Brown, G; Laster, M; Nelson, G

1999-01-01

Simulation of hypersonic flight in ground test and evaluation (T and E) facilities is a challenging and formidable task, especially to fully duplicate the flight environment above approximately Mach 8 for most all hypersonic flight systems that have been developed, conceived, or envisioned. Basically, and for many years, the enabling technology to build such a ground test wind tunnel facility has been severely limited in the area of high-temperature, high-strength materials and thermal protection approaches. To circumvent the problems, various approaches have been used, including partial simulation and use of similarity laws and reduced test time. These approaches often are not satisfactory, i.e. operability and durability testing for air-breathing propulsion development and thermal protection development of many flight systems. Thus, there is a strong need for science and technology (S and T) community involvement in technology development to address these problems. This paper discusses a specific case where this need exists and where significant S and T involvement has made and continues to make significant contributions. The case discussed will be an Air Force research program currently underway to develop enabling technologies for a Mach 8-15 hypersonic true temperature wind tunnel with relatively long run time. The research is based on a concept proposed by princeton University using radiant or beamed energy into the supersonic nozzle flow

High-Lift Propeller System Configuration Selection for NASA's SCEPTOR Distributed Electric Propulsion Flight Demonstrator

Science.gov (United States)

Patterson, Michael D.; Derlaga, Joseph M.; Borer, Nicholas K.

2016-01-01

Although the primary function of propellers is typically to produce thrust, aircraft equipped with distributed electric propulsion (DEP) may utilize propellers whose main purpose is to act as a form of high-lift device. These \\high-lift propellers" can be placed upstream of wing such that, when the higher-velocity ow in the propellers' slipstreams interacts with the wing, the lift is increased. This technique is a main design feature of a new NASA advanced design project called Scalable Convergent Electric Propulsion Technology Operations Research (SCEPTOR). The goal of the SCEPTOR project is design, build, and y a DEP aircraft to demonstrate that such an aircraft can be much more ecient than conventional designs. This paper provides details into the high-lift propeller system con guration selection for the SCEPTOR ight demonstrator. The methods used in the high-lift propeller system conceptual design and the tradeo s considered in selecting the number of propellers are discussed.

Affordable Flight Demonstration of the GTX Air-Breathing SSTO Vehicle Concept

Science.gov (United States)

Krivanek, Thomas M.; Roche, Joseph M.; Riehl, John P.; Kosareo, Daniel N.

2003-01-01

The rocket based combined cycle (RBCC) powered single-stage-to-orbit (SSTO) reusable launch vehicle has the potential to significantly reduce the total cost per pound for orbital payload missions. To validate overall system performance, a flight demonstration must be performed. This paper presents an overview of the first phase of a flight demonstration program for the GTX SSTO vehicle concept. Phase 1 will validate the propulsion performance of the vehicle configuration over the supersonic and hypersonic air- breathing portions of the trajectory. The focus and goal of Phase 1 is to demonstrate the integration and performance of the propulsion system flowpath with the vehicle aerodynamics over the air-breathing trajectory. This demonstrator vehicle will have dual mode ramjetkcramjets, which include the inlet, combustor, and nozzle with geometrically scaled aerodynamic surface outer mold lines (OML) defining the forebody, boundary layer diverter, wings, and tail. The primary objective of this study is to demon- strate propulsion system performance and operability including the ram to scram transition, as well as to validate vehicle aerodynamics and propulsion airframe integration. To minimize overall risk and develop ment cost the effort will incorporate proven materials, use existing turbomachinery in the propellant delivery systems, launch from an existing unmanned remote launch facility, and use basic vehicle recovery techniques to minimize control and landing requirements. A second phase would demonstrate propulsion performance across all critical portions of a space launch trajectory (lift off through transition to all-rocket) integrated with flight-like vehicle systems.

CFD for hypersonic airbreathing aircraft

Science.gov (United States)

Kumar, Ajay

1989-01-01

A general discussion is given on the use of advanced computational fluid dynamics (CFD) in analyzing the hypersonic flow field around an airbreathing aircraft. Unique features of the hypersonic flow physics are presented and an assessment is given of the current algorithms in terms of their capability to model hypersonic flows. Several examples of advanced CFD applications are then presented.

Aurora Flight Sciences' Perseus B Remotely Piloted Aircraft in Flight

Science.gov (United States)

1998-01-01

A long, slender wing and a pusher propeller at the rear characterize the Perseus B remotely piloted research aircraft, seen here during a test flight in June 1998. Perseus B is a remotely piloted aircraft developed as a design-performance testbed under NASA's Environmental Research Aircraft and Sensor Technology (ERAST) project. Perseus is one of several flight vehicles involved in the ERAST project. A piston engine, propeller-powered aircraft, Perseus was designed and built by Aurora Flight Sciences Corporation, Manassas, Virginia. The objectives of Perseus B's ERAST flight tests have been to reach and maintain horizontal flight above altitudes of 60,000 feet and demonstrate the capability to fly missions lasting from 8 to 24 hours, depending on payload and altitude requirements. The Perseus B aircraft established an unofficial altitude record for a single-engine, propeller-driven, remotely piloted aircraft on June 27, 1998. It reached an altitude of 60,280 feet. In 1999, several modifications were made to the Perseus aircraft including engine, avionics, and flight-control-system improvements. These improvements were evaluated in a series of operational readiness and test missions at the Dryden Flight Research Center, Edwards, California. Perseus is a high-wing monoplane with a conventional tail design. Its narrow, straight, high-aspect-ratio wing is mounted atop the fuselage. The aircraft is pusher-designed with the propeller mounted in the rear. This design allows for interchangeable scientific-instrument payloads to be placed in the forward fuselage. The design also allows for unobstructed airflow to the sensors and other devices mounted in the payload compartment. The Perseus B that underwent test and development in 1999 was the third generation of the Perseus design, which began with the Perseus Proof-Of-Concept aircraft. Perseus was initially developed as part of NASA's Small High-Altitude Science Aircraft (SHASA) program, which later evolved into the ERAST

Modelling of Influence of Hypersonic Conditions on Gyroscopic Inertial Navigation Sensor Suspension

Directory of Open Access Journals (Sweden)

Korobiichuk Igor

2017-06-01

Full Text Available The upcoming hypersonic technologies pose a difficult task for air navigation systems. The article presents a designed model of elastic interaction of penetrating acoustic radiation with flat isotropic suspension elements of an inertial navigation sensor in the operational conditions of hypersonic flight. It has been shown that the acoustic transparency effect in the form of a spatial-frequency resonance becomes possible with simultaneous manifestation of the wave coincidence condition in the acoustic field and equality of the natural oscillation frequency of a finite-size plate and a forced oscillation frequency of an infinite plate. The effect can lead to additional measurement errors of the navigation system. Using the model, the worst and best case suspension oscillation frequencies can be determined, which will help during the design of a navigation system.

Adaptive Command Filtered Integrated Guidance and Control for Hypersonic Vehicle with Magnitude, Rate and Bandwidth Constraints

Directory of Open Access Journals (Sweden)

Wang Liang

2018-01-01

Full Text Available This paper proposes a novel integrated guidance and control (IGC method for hypersonic vehicle in terminal phase. Firstly, the system model is developed with a second order actuator dynamics. Then the back-stepping controller is designed hierarchically with command filters, where the first order command filters are implemented to construct the virtual control input with ideal states predicted by an adaptive estimator, and the nonlinear command filter is designed to produce magnitude, rate and bandwidth limited control surface deflection finally tracked by a terminal sliding mode controller with finite convergence time. Through a series of 6-DOF numerical simulations, it’s indicated that the proposed method successfully cancels out the large aerodynamics coefficient uncertainties and disturbances in hypersonic flight under limited control surface deflection. The contribution of this paper lies in the application and determination of nonlinear integrated design of guidance and control system for hypersonic vehicle.

Modeling, Measurements, and Fundamental Database Development for Nonequilibrium Hypersonic Aerothermodynamics

Science.gov (United States)

Bose, Deepak

2012-01-01

The design of entry vehicles requires predictions of aerothermal environment during the hypersonic phase of their flight trajectories. These predictions are made using computational fluid dynamics (CFD) codes that often rely on physics and chemistry models of nonequilibrium processes. The primary processes of interest are gas phase chemistry, internal energy relaxation, electronic excitation, nonequilibrium emission and absorption of radiation, and gas-surface interaction leading to surface recession and catalytic recombination. NASAs Hypersonics Project is advancing the state-of-the-art in modeling of nonequilibrium phenomena by making detailed spectroscopic measurements in shock tube and arcjets, using ab-initio quantum mechanical techniques develop fundamental chemistry and spectroscopic databases, making fundamental measurements of finite-rate gas surface interactions, implementing of detailed mechanisms in the state-of-the-art CFD codes, The development of new models is based on validation with relevant experiments. We will present the latest developments and a roadmap for the technical areas mentioned above

Combining MHD Airbreathing and Fusion Rocket Propulsion for Earth-to-Orbit Flight

International Nuclear Information System (INIS)

Froning, H. D. Jr; Yang, Yang; Momota, H.; Burton, E.; Miley, G. H.; Luo, Nie

2005-01-01

Previous studies have shown that Single-State-to-Orbit (SSTO) vehicle propellant can be reduced by Magnets-Hydro-Dynamic (MHD) processes that minimize airbreathing propulsion losses and propellant consumption during atmospheric flight. Similarly additional reduction in SSTO propellant is enabled by Inertial Electrostatic Confinement (IEC) fusion, whose more energetic reactions reduce rocket propellant needs. MHD airbreathing propulsion during an SSTO vehicle's initial atmospheric flight phase and IEC fusion propulsion during its final exo-atmospheric flight phase is therefore being explored. Accomplished work is not yet sufficient for claiming such a vehicle's feasibility. But takeoff and propellant mass for an MHD airbreathing and IEC fusion vehicle could be as much as 25 and 40 percent less than one with ordinary airbreathing and IEC fusion; and as much as 50 and 70 percent less than SSTO takeoff and propellant mass with MHD airbreathing and chemical rocket propulsion. Thus this unusual combined cycle engine shows great promise for performance gains beyond contemporary combined-cycle airbreathing engines

Experimental research on air propellers

Science.gov (United States)

Durand, William F

1918-01-01

The purposes of the experimental investigation on the performance of air propellers described in this report are as follows: (1) the development of a series of design factors and coefficients drawn from model forms distributed with some regularity over the field of air-propeller design and intended to furnish a basis of check with similar work done in other aerodynamic laboratories, and as a point of departure for the further study of special or individual types and forms; (2) the establishment of a series of experimental values derived from models and intended for later use as a basis for comparison with similar results drawn from certain selected full-sized forms and tested in free flight.

HIFiRE-1 Turbulent Shock Boundary Layer Interaction - Flight Data and Computations

Science.gov (United States)

Kimmel, Roger L.; Prabhu, Dinesh

2015-01-01

The Hypersonic International Flight Research Experimentation (HIFiRE) program is a hypersonic flight test program executed by the Air Force Research Laboratory (AFRL) and Australian Defence Science and Technology Organisation (DSTO). This flight contained a cylinder-flare induced shock boundary layer interaction (SBLI). Computations of the interaction were conducted for a number of times during the ascent. The DPLR code used for predictions was calibrated against ground test data prior to exercising the code at flight conditions. Generally, the computations predicted the upstream influence and interaction pressures very well. Plateau pressures on the cylinder were predicted well at all conditions. Although the experimental heat transfer showed a large amount of scatter, especially at low heating levels, the measured heat transfer agreed well with computations. The primary discrepancy between the experiment and computation occurred in the pressures measured on the flare during second stage burn. Measured pressures exhibited large overshoots late in the second stage burn, the mechanism of which is unknown. The good agreement between flight measurements and CFD helps validate the philosophy of calibrating CFD against ground test, prior to exercising it at flight conditions.

Recombination Catalysts for Hypersonic Fuels

Science.gov (United States)

Chinitz, W.

1998-01-01

The goal of commercially-viable access to space will require technologies that reduce propulsion system weight and complexity, while extracting maximum energy from the products of combustion. This work is directed toward developing effective nozzle recombination catalysts for the supersonic and hypersonic aeropropulsion engines used to provide such access to space. Effective nozzle recombination will significantly reduce rk=le length (hence, propulsion system weight) and reduce fuel requirements, further decreasing the vehicle's gross lift-off weight. Two such catalysts have been identified in this work, barium and antimony compounds, by developing chemical kinetic reaction mechanisms for these materials and determining the engine performance enhancement for a typical flight trajectory. Significant performance improvements are indicated, using only 2% (mole or mass) of these compounds in the combustor product gas.

Space Shuttle and Hypersonic Entry

Science.gov (United States)

Campbell, Charles H.; Gerstenmaier, William H.

2014-01-01

Fifty years of human spaceflight have been characterized by the aerospace operations of the Soyuz, of the Space Shuttle and, more recently, of the Shenzhou. The lessons learned of this past half decade are important and very significant. Particularly interesting is the scenario that is downstream from the retiring of the Space Shuttle. A number of initiatives are, in fact, emerging from in the aftermath of the decision to terminate the Shuttle program. What is more and more evident is that a new era is approaching: the era of the commercial usage and of the commercial exploitation of space. It is probably fair to say, that this is the likely one of the new frontiers of expansion of the world economy. To make a comparison, in the last 30 years our economies have been characterized by the digital technologies, with examples ranging from computers, to cellular phones, to the satellites themselves. Similarly, the next 30 years are likely to be characterized by an exponential increase of usage of extra atmospheric resources, as a result of more economic and efficient way to access space, with aerospace transportation becoming accessible to commercial investments. We are witnessing the first steps of the transportation of future generation that will drastically decrease travel time on our Planet, and significantly enlarge travel envelope including at least the low Earth orbits. The Steve Jobs or the Bill Gates of the past few decades are being replaced by the aggressive and enthusiastic energy of new entrepreneurs. It is also interesting to note that we are now focusing on the aerospace band, that lies on top of the aeronautical shell, and below the low Earth orbits. It would be a mistake to consider this as a known envelope based on the evidences of the flights of Soyuz, Shuttle and Shenzhou. Actually, our comprehension of the possible hypersonic flight regimes is bounded within really limited envelopes. The achievement of a full understanding of the hypersonic flight

Nonlinear control of linear parameter varying systems with applications to hypersonic vehicles

Science.gov (United States)

Wilcox, Zachary Donald

The focus of this dissertation is to design a controller for linear parameter varying (LPV) systems, apply it specifically to air-breathing hypersonic vehicles, and examine the interplay between control performance and the structural dynamics design. Specifically a Lyapunov-based continuous robust controller is developed that yields exponential tracking of a reference model, despite the presence of bounded, nonvanishing disturbances. The hypersonic vehicle has time varying parameters, specifically temperature profiles, and its dynamics can be reduced to an LPV system with additive disturbances. Since the HSV can be modeled as an LPV system the proposed control design is directly applicable. The control performance is directly examined through simulations. A wide variety of applications exist that can be effectively modeled as LPV systems. In particular, flight systems have historically been modeled as LPV systems and associated control tools have been applied such as gain-scheduling, linear matrix inequalities (LMIs), linear fractional transformations (LFT), and mu-types. However, as the type of flight environments and trajectories become more demanding, the traditional LPV controllers may no longer be sufficient. In particular, hypersonic flight vehicles (HSVs) present an inherently difficult problem because of the nonlinear aerothermoelastic coupling effects in the dynamics. HSV flight conditions produce temperature variations that can alter both the structural dynamics and flight dynamics. Starting with the full nonlinear dynamics, the aerothermoelastic effects are modeled by a temperature dependent, parameter varying state-space representation with added disturbances. The model includes an uncertain parameter varying state matrix, an uncertain parameter varying non-square (column deficient) input matrix, and an additive bounded disturbance. In this dissertation, a robust dynamic controller is formulated for a uncertain and disturbed LPV system. The developed

Air-Breathing Hypersonic Vehicle Tracking Control Based on Adaptive Dynamic Programming.

Science.gov (United States)

Mu, Chaoxu; Ni, Zhen; Sun, Changyin; He, Haibo

2017-03-01

In this paper, we propose a data-driven supplementary control approach with adaptive learning capability for air-breathing hypersonic vehicle tracking control based on action-dependent heuristic dynamic programming (ADHDP). The control action is generated by the combination of sliding mode control (SMC) and the ADHDP controller to track the desired velocity and the desired altitude. In particular, the ADHDP controller observes the differences between the actual velocity/altitude and the desired velocity/altitude, and then provides a supplementary control action accordingly. The ADHDP controller does not rely on the accurate mathematical model function and is data driven. Meanwhile, it is capable to adjust its parameters online over time under various working conditions, which is very suitable for hypersonic vehicle system with parameter uncertainties and disturbances. We verify the adaptive supplementary control approach versus the traditional SMC in the cruising flight, and provide three simulation studies to illustrate the improved performance with the proposed approach.

Adaptive terminal sliding mode control for hypersonic flight vehicles with strictly lower convex function based nonlinear disturbance observer.

Science.gov (United States)

Wu, Yun-Jie; Zuo, Jing-Xing; Sun, Liang-Hua

2017-11-01

In this paper, the altitude and velocity tracking control of a generic hypersonic flight vehicle (HFV) is considered. A novel adaptive terminal sliding mode controller (ATSMC) with strictly lower convex function based nonlinear disturbance observer (SDOB) is proposed for the longitudinal dynamics of HFV in presence of both parametric uncertainties and external disturbances. First, for the sake of enhancing the anti-interference capability, SDOB is presented to estimate and compensate the equivalent disturbances by introducing a strictly lower convex function. Next, the SDOB based ATSMC (SDOB-ATSMC) is proposed to guarantee the system outputs track the reference trajectory. Then, stability of the proposed control scheme is analyzed by the Lyapunov function method. Compared with other HFV control approaches, key novelties of SDOB-ATSMC are that a novel SDOB is proposed and drawn into the (virtual) control laws to compensate the disturbances and that several adaptive laws are used to deal with the differential explosion problem. Finally, it is illustrated by the simulation results that the new method exhibits an excellent robustness and a better disturbance rejection performance than the convention approach. Copyright © 2017 ISA. Published by Elsevier Ltd. All rights reserved.

Experimental And Numerical Investigation Of Aerothermal Characteristics Of The IXV Hypersonic Vehicle

Science.gov (United States)

Paris, S.; Charbonnier, D.; Tran, D.

2011-05-01

The main results of the aerothermodynamic hypersonic characterization of Intermediate eXperimental Vehicle (IXV), by means of both CFD simulations and wind tunnel measurements, have been reported and analyzed. In the framework of ESA FLPP Program, the VKI (Von Karman Institute) was in charge of an experimental test campaign for the consolidation of the aerothermal database in cold hypersonic regime. The tests campaign has been carried out at VKI Free Piston Longshot wind tunnel at mach 14. The numerical simulations have been performed for VKI wind tunnel conditions by CFSE with the in-house NSMB flow solver (Navier-Stokes Multi-Blocks 3D), the goal being to support the procedure of extrapolation-to-flight of the measurements and the general aerothermal characterization. Laminar, transitional and fully turbulent flows have been computed, with air considered as an ideal gas, for the wind tunnel tests numerical rebuilding. A detailed comparison of all measured and predicted hypersonic relevant phenomena and parameters (surface pressure and heat flux) is reported in the paper, together with a detailed description of configuration, freestream conditions, model attitude effects and flap deflection effect. The detailed analyze of the experimental and numerical data gives information on the nature of the flow on the body and on the flaps for the most critical configuration

Influence of different propellant systems on ablation of EPDM insulators in overload state

Science.gov (United States)

Guan, Yiwen; Li, Jiang; Liu, Yang; Xu, Tuanwei

2018-04-01

This study examines the propellants used in full-scale solid rocket motors (SRM) and investigates how insulator ablation is affected by two propellant formulations (A and B) during flight overload conditions. An experimental study, theoretical analysis, and numerical simulations were performed to discover the intrinsic causes of insulator ablation rates from the perspective of lab-scaled ground-firing tests, the decoupling of thermochemical ablation, and particle erosion. In addition, the difference in propellant composition, and the insulator charring layer microstructure were analyzed. Results reveal that the degree of insulator ablation is positively correlated with the propellant burn rate, particle velocity, and aggregate concentrations during the condensed phase. A lower ratio of energetic additive material in the AP oxidizer of the propellant is promising for the reduction in particle size and increase in the burn rate and pressure index. However, the overall higher velocity of a two-phase flow causes severe erosion of the insulation material. While the higher ratio of energetic additive to the AP oxidizer imparts a smaller ablation rate to the insulator (under lab-scale test conditions), the slag deposition problem in the combustion chamber may cause catastrophic consequences for future large full-scale SRM flight experiments.

New Hypersonic Shock Tunnel at the Laboratory of Aerothermodynamics and Hypersonics Prof. Henry T. Nagamatsu

International Nuclear Information System (INIS)

Toro, P. G. P.; Minucci, M. A. S.; Chanes, J. B. Jr; Oliveira, A. C.; Gomes, F. A. A.; Myrabo, L. N.; Nagamatsu, Henry T.

2008-01-01

The new 0.60-m. nozzle exit diameter hypersonic shock tunnel was designed to study advanced air-breathing propulsion system such as supersonic combustion and/or laser technologies. In addition, it may be used for hypersonic flow studies and investigations of the electromagnetic (laser) energy addition for flow control. This new hypersonic shock tunnel was designed and installed at the Laboratory for of Aerothermodynamics and Hypersonics Prof. Henry T. Nagamatsu, IEAv-CTA, Brazil. The design of the tunnel enables relatively long test times, 2-10 milliseconds, suitable for the experiments performed at the laboratory. Free stream Mach numbers ranging from 6 to 25 can be produced and stagnation pressures and temperatures up to 360 atm. and up to 9,000 K, respectively, can be generated. Shadowgraph and schlieren optical techniques will be used for flow visualization

Prediction of shock-layer ultraviolet radiation for hypersonic vehicles in near space

Directory of Open Access Journals (Sweden)

Niu Qinglin

2016-10-01

Full Text Available A systemic and validated model was developed to predict ultraviolet spectra features from the shock layer of near-space hypersonic vehicles in the “solar blind” band region. Computational procedures were performed with 7-species thermal non-equilibrium fluid mechanics, finite rate chemistry, and radiation calculations. The thermal non-equilibrium flow field was calculated with a two-temperature model by the finite volume technique and verified against the bow-shock ultra-violet (BSUV flight experiments. The absorption coefficient of the mixture gases was evaluated with a line-by-line method and validated through laboratory shock tube measurements. Using the line of sight (LOS method, radiation was calculated from three BSUV flights at altitudes of 38, 53.5 and 71 km. The investigation focused on the level and structure of ultraviolet spectra radiated from a NO band system in wavelengths of 200–400 nm. Results predicted by the current model show qualitative spatial agreement with the measured data. At a velocity of 3.5 km/s (about Mach 11, the peak absolute intensity at an altitude of 38 km is two orders of magnitude higher than that at 53.5 km. Under the same flight conditions, the spectra structures have quite a similar distribution at different viewing angles. The present computational model performs well in the prediction of the ultraviolet spectra emitted from the shock layer and will contribute to the investigation and analysis of radiative features of hypersonic vehicles in near space.

Perseus Post-flight

Science.gov (United States)

1991-01-01

Crew members check out the Perseus proof-of-concept vehicle on Rogers Dry Lake, adjacent to the Dryden Flight Research Center, Edwards, California, after a test flight in 1991. Perseus B is a remotely piloted aircraft developed as a design-performance testbed under NASA's Environmental Research Aircraft and Sensor Technology (ERAST) project. Perseus is one of several flight vehicles involved in the ERAST project. A piston engine, propeller-powered aircraft, Perseus was designed and built by Aurora Flight Sciences Corporation, Manassas, Virginia. The objectives of Perseus B's ERAST flight tests have been to reach and maintain horizontal flight above altitudes of 60,000 feet and demonstrate the capability to fly missions lasting from 8 to 24 hours, depending on payload and altitude requirements. The Perseus B aircraft established an unofficial altitude record for a single-engine, propeller-driven, remotely piloted aircraft on June 27, 1998. It reached an altitude of 60,280 feet. In 1999, several modifications were made to the Perseus aircraft including engine, avionics, and flight-control-system improvements. These improvements were evaluated in a series of operational readiness and test missions at the Dryden Flight Research Center, Edwards, California. Perseus is a high-wing monoplane with a conventional tail design. Its narrow, straight, high-aspect-ratio wing is mounted atop the fuselage. The aircraft is pusher-designed with the propeller mounted in the rear. This design allows for interchangeable scientific-instrument payloads to be placed in the forward fuselage. The design also allows for unobstructed airflow to the sensors and other devices mounted in the payload compartment. The Perseus B that underwent test and development in 1999 was the third generation of the Perseus design, which began with the Perseus Proof-Of-Concept aircraft. Perseus was initially developed as part of NASA's Small High-Altitude Science Aircraft (SHASA) program, which later evolved

Lateral-directional aerodynamic characteristics of light, twin-engine, propeller driven airplanes

Science.gov (United States)

Wolowicz, C. H.; Yancey, R. B.

1972-01-01

Analytical procedures and design data for predicting the lateral-directional static and dynamic stability and control characteristics of light, twin engine, propeller driven airplanes for propeller-off and power-on conditions are reported. Although the consideration of power effects is limited to twin engine airplanes, the propeller-off considerations are applicable to single engine airplanes as well. The procedures are applied to a twin engine, propeller driven, semi-low-wing airplane in the clean configuration through the linear lift range. The calculated derivative characteristics are compared with wind tunnel and flight data. Included in the calculated characteristics are the spiral mode, roll mode, and Dutch roll mode over the speed range of the airplane.

Effects of acoustic treatment on the interior noise levels of a twin-engine propeller aircraft - Experimental flight results and theoretical predictions

Science.gov (United States)

Beyer, T. B.; Powell, C. A.; Daniels, E. F.; Pope, L. D.

1984-01-01

In-flight noise level measurements were made within two cabin configurations of a general aviation business aircraft. The Fairchild Merlin IVC twin-engine aircraft was tested with bare walls and fiberglass insulation and in an executive trim configuration. Narrow-band and octave format data were subjected to analyses which permitted identification of the blade passage harmonics (BPH). Cabin noise level reductions (insertion losses) due to added insulation varied with position in the cabin, the BPH number, cabin pressure, and engine torque. The measurements were closely predicted using the propeller aircraft interior noise (PAIN) mode.

Perancangan Propeler Self-Propelled Barge

Directory of Open Access Journals (Sweden)

Billy Teguh kurniawan

2013-03-01

Full Text Available Makalah ini menyampaikan suatu penelitian tentang perancangan propeler yang optimal beserta pemilihan daya mesin yang efisien pada self-propelled barge dengan memperhitungkan besarnya nilai tahanan dari barge tersebut. Dengan penambahan sistem propulsi, diharapkan barge dapat beroperasi dengan lebih efisien dibandingkan saat barge beroperasi menggunakan sistem towing atau ditarik tug boat. Perhitungan tahanan barge dilakukan menggunakan metode Holtrop dan Guldhammer-Harvald sehingga dapat diperhi-tungkan geometri dan jenis propeler yang optimal beserta daya mesin yang efisien untuk barge. Propeler yang dianalisis adalah propeler tipe B-Troost Series, sedangkan variasi yang dilakukan untuk perencanaan propeler pada kajian ini adalah variasi putaran propeler pada rentang antara 310-800 rpm, serta variasi jumlah daun pada rentang tiga, empat, lima, dan enam. Besarnya nilai tahanan self-propelled barge untuk metode Holtrop adalah 105.91 kilonewton, sedangkan hasil per-hitungan dari metode Guldhammer-Harvald didapatkan nilai sebesar 109.14 kilonewton. Tipe propeler yang dipilih setelah dilakukan uji kavitasi adalah tipe Troost Series B4-40, dengan diameter sebesar 2.1 m, efisiensi sebesar 0.421, pitch ratio se-besar 0.591, dengan putaran propeler 400 rpm. Daya mesin yg dibutuhkan barge pada kondisi maksimum (BHPMCR sebesar 1669.5 HP. Dengan mempertimbangkan daya tersebut, maka dipilih mesin jenis Caterpillar tipe Marine 3516B yang mem-punyai daya maksimum sebesar 1285 kilowatt atau 1722.5 horsepower dengan putaran mesin sebesar 1200 rpm

Perseus in Flight

Science.gov (United States)

1991-01-01

The Perseus proof-of-concept vehicle flies over Rogers Dry Lake at the Dryden Flight Research Center, Edwards, California, to test basic design concepts for the remotely-piloted, high-altitude vehicle. Perseus B is a remotely piloted aircraft developed as a design-performance testbed under NASA's Environmental Research Aircraft and Sensor Technology (ERAST) project. Perseus is one of several flight vehicles involved in the ERAST project. A piston engine, propeller-powered aircraft, Perseus was designed and built by Aurora Flight Sciences Corporation, Manassas, Virginia. The objectives of Perseus B's ERAST flight tests have been to reach and maintain horizontal flight above altitudes of 60,000 feet and demonstrate the capability to fly missions lasting from 8 to 24 hours, depending on payload and altitude requirements. The Perseus B aircraft established an unofficial altitude record for a single-engine, propeller-driven, remotely piloted aircraft on June 27, 1998. It reached an altitude of 60,280 feet. In 1999, several modifications were made to the Perseus aircraft including engine, avionics, and flight-control-system improvements. These improvements were evaluated in a series of operational readiness and test missions at the Dryden Flight Research Center, Edwards, California. Perseus is a high-wing monoplane with a conventional tail design. Its narrow, straight, high-aspect-ratio wing is mounted atop the fuselage. The aircraft is pusher-designed with the propeller mounted in the rear. This design allows for interchangeable scientific-instrument payloads to be placed in the forward fuselage. The design also allows for unobstructed airflow to the sensors and other devices mounted in the payload compartment. The Perseus B that underwent test and development in 1999 was the third generation of the Perseus design, which began with the Perseus Proof-Of-Concept aircraft. Perseus was initially developed as part of NASA's Small High-Altitude Science Aircraft (SHASA

A Flight Demonstration of Plasma Rocket Propulsion

Science.gov (United States)

Petro, Andrew

1999-01-01

The Advanced Space Propulsion Laboratory at the Johnson Space Center has been engaged in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One advantageous feature of this rocket concept is the ability to vary its specific impulse so that it can be operated in a mode which maximizes propellant efficiency or a mode which maximizes thrust. This presentation will describe a proposed flight experiment in which a simple version of the rocket will be tested in space. In addition to the plasma rocket, the flight experiment will also demonstrate the use of a superconducting electromagnet, extensive use of heat pipes, and possibly the transfer of cryogenic propellant in space.

Hypersonic Materials and Structures

Science.gov (United States)

Glass, David E.

2016-01-01

Thermal protection systems (TPS) and hot structures are required for a range of hypersonic vehicles ranging from ballistic reentry to hypersonic cruise vehicles, both within Earth's atmosphere and non-Earth atmospheres. The focus of this presentation is on air breathing hypersonic vehicles in the Earth's atmosphere. This includes single-stage to orbit (SSTO), two-stage to orbit (TSTO) accelerators, access to space vehicles, and hypersonic cruise vehicles. This paper will start out with a brief discussion of aerodynamic heating and thermal management techniques to address the high heating, followed by an overview of TPS for rocket-launched and air-breathing vehicles. The argument is presented that as we move from rocket-based vehicles to air-breathing vehicles, we need to move away from the insulated airplane approach used on the Space Shuttle Orbiter to a wide range of TPS and hot structure approaches. The primary portion of the paper will discuss issues and design options for CMC TPS and hot structure components, including leading edges, acreage TPS, and control surfaces. The current state-of-the-art will be briefly discussed for some of the components.

Green Propellant Infusion Mission Program Development and Technology Maturation

Science.gov (United States)

McLean, Christopher H.; Deininger, William D.; Joniatis, John; Aggarwal, Pravin K.; Spores, Ronald A.; Deans, Matthew; Yim, John T.; Bury, Kristen; Martinez, Jonathan; Cardiff, Eric H.;

A Simple Method for High-Lift Propeller Conceptual Design

Science.gov (United States)

Patterson, Michael; Borer, Nick; German, Brian

2016-01-01

In this paper, we present a simple method for designing propellers that are placed upstream of the leading edge of a wing in order to augment lift. Because the primary purpose of these "high-lift propellers" is to increase lift rather than produce thrust, these props are best viewed as a form of high-lift device; consequently, they should be designed differently than traditional propellers. We present a theory that describes how these props can be designed to provide a relatively uniform axial velocity increase, which is hypothesized to be advantageous for lift augmentation based on a literature survey. Computational modeling indicates that such propellers can generate the same average induced axial velocity while consuming less power and producing less thrust than conventional propeller designs. For an example problem based on specifications for NASA's Scalable Convergent Electric Propulsion Technology and Operations Research (SCEPTOR) flight demonstrator, a propeller designed with the new method requires approximately 15% less power and produces approximately 11% less thrust than one designed for minimum induced loss. Higher-order modeling and/or wind tunnel testing are needed to verify the predicted performance.

Multi-Objective Trajectory Optimization of a Hypersonic Reconnaissance Vehicle with Temperature Constraints

Science.gov (United States)

Masternak, Tadeusz J.

This research determines temperature-constrained optimal trajectories for a scramjet-based hypersonic reconnaissance vehicle by developing an optimal control formulation and solving it using a variable order Gauss-Radau quadrature collocation method with a Non-Linear Programming (NLP) solver. The vehicle is assumed to be an air-breathing reconnaissance aircraft that has specified takeoff/landing locations, airborne refueling constraints, specified no-fly zones, and specified targets for sensor data collections. A three degree of freedom scramjet aircraft model is adapted from previous work and includes flight dynamics, aerodynamics, and thermal constraints. Vehicle control is accomplished by controlling angle of attack, roll angle, and propellant mass flow rate. This model is incorporated into an optimal control formulation that includes constraints on both the vehicle and mission parameters, such as avoidance of no-fly zones and coverage of high-value targets. To solve the optimal control formulation, a MATLAB-based package called General Pseudospectral Optimal Control Software (GPOPS-II) is used, which transcribes continuous time optimal control problems into an NLP problem. In addition, since a mission profile can have varying vehicle dynamics and en-route imposed constraints, the optimal control problem formulation can be broken up into several "phases" with differing dynamics and/or varying initial/final constraints. Optimal trajectories are developed using several different performance costs in the optimal control formulation: minimum time, minimum time with control penalties, and maximum range. The resulting analysis demonstrates that optimal trajectories that meet specified mission parameters and constraints can be quickly determined and used for larger-scale operational and campaign planning and execution.

A QMU approach for characterizing the operability limits of air-breathing hypersonic vehicles

International Nuclear Information System (INIS)

Iaccarino, Gianluca; Pecnik, Rene; Glimm, James; Sharp, David

2011-01-01

The operability limits of a supersonic combustion engine for an air-breathing hypersonic vehicle are characterized using numerical simulations and an uncertainty quantification methodology. The time-dependent compressible flow equations with heat release are solved in a simplified configuration. Verification, calibration and validation are carried out to assess the ability of the model to reproduce the flow/thermal interactions that occur when the engine unstarts due to thermal choking. quantification of margins and uncertainty (QMU) is used to determine the safe operation region for a range of fuel flow rates and combustor geometries. - Highlights: → In this work we introduce a method to study the operability limits of hypersonic scramjet engines. → The method is based on a calibrated heat release model. → It accounts explicitly for uncertainties due to flight conditions and model correlations. → We examine changes due to the combustor geometry and fuel injection.

Assessment of Turbulent CFD Against STS-128 Hypersonic Flight Data

Science.gov (United States)

Wood, William A.; Kleb, William L.; Hyatt, Andrew J.

2010-01-01

Turbulent CFD simulations are compared against surface temperature measurements of the space shuttle orbiter windward tiles at reentry flight conditions. Algebraic turbulence models are used within both the LAURA and DPLR CFD codes. The flight data are from temperature measurements obtained by seven thermocouples during the STS-128 mission (September 2009). The flight data indicate boundary layer transition onset over the Mach number range 13.5{15.5, depending upon the location on the vehicle. But the boundary layer flow appeared to be transitional down through Mach 12, based upon the flight data and CFD trends. At Mach 9 the simulations match the flight data on average within 20 F/11 C, where typical surface temperatures were approximately 1600 F/870 C.

Hypersonic sliding target tracking in near space

Directory of Open Access Journals (Sweden)

Xiang-yu Zhang

2015-12-01

Full Text Available To improve the tracking accuracy of hypersonic sliding target in near space, the influence of target hypersonic movement on radar detection and tracking is analyzed, and an IMM tracking algorithm is proposed based on radial velocity compensating and cancellation processing of high dynamic biases under the earth centered earth fixed (ECEF coordinate. Based on the analysis of effect of target hypersonic movement, a measurement model is constructed to reduce the filter divergence which is caused by the model mismatch. The high dynamic biases due to the target hypersonic movement are approximately compensated through radial velocity estimation to achieve the hypersonic target tracking at low systematic biases in near space. The high dynamic biases are further eliminated by the cancellation processing of different radars, in which the track association problem can be solved when the dynamic biases are low. An IMM algorithm based on constant acceleration (CA, constant turning (CT and Singer models is used to achieve the hypersonic sliding target tracking in near space. Simulation results show that the target tracking in near space can be achieved more effectively by using the proposed algorithm.

Vent System Analysis for the Cryogenic Propellant Storage Transfer Ground Test Article

Science.gov (United States)

Hedayat, A

2013-01-01

To test and validate key capabilities and technologies required for future exploration elements such as large cryogenic propulsion stages and propellant depots, NASA is leading the efforts to develop and design the Cryogenic Propellant Storage and Transfer (CPST) Cryogenic Fluid Management (CFM) payload. The primary objectives of CPST payload are to demonstrate: 1) in-space storage of cryogenic propellants for long duration applications; and 2) in-space transfer of cryogenic propellants. The Ground Test Article (GTA) is a technology development version of the CPST payload. The GTA consists of flight-sized and flight-like storage and transfer tanks, liquid acquisition devices, transfer, and pressurization systems with all of the CPST functionality. The GTA is designed to perform integrated passive and active thermal storage and transfer performance testing with liquid hydrogen (LH2) in a vacuum environment. The GTA storage tank is designed to store liquid hydrogen and the transfer tank is designed to be 5% of the storage tank volume. The LH2 transfer subsystem is designed to transfer propellant from one tank to the other utilizing pressure or a pump. The LH2 vent subsystem is designed to prevent over-pressurization of the storage and transfer tanks. An in-house general-purpose computer program was utilized to model and simulate the vent subsystem operation. The modeling, analysis, and the results will be presented in the final paper.

Photogrammetry of a Hypersonic Inflatable Aerodynamic Decelerator

Science.gov (United States)

Kushner, Laura Kathryn; Littell, Justin D.; Cassell, Alan M.

2013-01-01

In 2012, two large-scale models of a Hypersonic Inflatable Aerodynamic decelerator were tested in the National Full-Scale Aerodynamic Complex at NASA Ames Research Center. One of the objectives of this test was to measure model deflections under aerodynamic loading that approximated expected flight conditions. The measurements were acquired using stereo photogrammetry. Four pairs of stereo cameras were mounted inside the NFAC test section, each imaging a particular section of the HIAD. The views were then stitched together post-test to create a surface deformation profile. The data from the photogram- metry system will largely be used for comparisons to and refinement of Fluid Structure Interaction models. This paper describes how a commercial photogrammetry system was adapted to make the measurements and presents some preliminary results.

Anisotropic power spectrum of refractive-index fluctuation in hypersonic turbulence.

Science.gov (United States)

Li, Jiangting; Yang, Shaofei; Guo, Lixin; Cheng, Mingjian

2016-11-10

An anisotropic power spectrum of the refractive-index fluctuation in hypersonic turbulence was obtained by processing the experimental image of the hypersonic plasma sheath and transforming the generalized anisotropic von Kármán spectrum. The power spectrum suggested here can provide as good a fit to measured spectrum data for hypersonic turbulence as that recorded from the nano-planar laser scattering image. Based on the newfound anisotropic hypersonic turbulence power spectrum, Rytov approximation was employed to establish the wave structure function and the spatial coherence radius model of electromagnetic beam propagation in hypersonic turbulence. Enhancing the anisotropy characteristics of the hypersonic turbulence led to a significant improvement in the propagation performance of electromagnetic beams in hypersonic plasma sheath. The influence of hypersonic turbulence on electromagnetic beams increases with the increase of variance of the refractive-index fluctuation and the decrease of turbulence outer scale and anisotropy parameters. The spatial coherence radius was much smaller than that in atmospheric turbulence. These results are fundamental to understanding electromagnetic wave propagation in hypersonic turbulence.

Numerical Modeling of Pressurization of Cryogenic Propellant Tank for Integrated Vehicle Fluid System

Science.gov (United States)

Majumdar, Alok K.; LeClair, Andre C.; Hedayat, Ali

2016-01-01

This paper presents a numerical model of pressurization of a cryogenic propellant tank for the Integrated Vehicle Fluid (IVF) system using the Generalized Fluid System Simulation Program (GFSSP). The IVF propulsion system, being developed by United Launch Alliance, uses boiloff propellants to drive thrusters for the reaction control system as well as to run internal combustion engines to develop power and drive compressors to pressurize propellant tanks. NASA Marshall Space Flight Center (MSFC) has been running tests to verify the functioning of the IVF system using a flight tank. GFSSP, a finite volume based flow network analysis software developed at MSFC, has been used to develop an integrated model of the tank and the pressurization system. This paper presents an iterative algorithm for converging the interface boundary conditions between different component models of a large system model. The model results have been compared with test data.

Fluid dynamic modeling and numerical simulation of low-density hypersonic flow

Science.gov (United States)

Cheng, H. K.; Wong, Eric Y.

1988-01-01

The concept of a viscous shock-layer and several related versions of continuum theories/methods are examined for their adequacy as a viable framework to study flow physics and aerothermodynamics of relevance to sustained hypersonic flights. Considering the flat plate at angle of attack, or the wedge, as a generic example for the major aerodynamic component of a hypersonic vehicle, the relative importance of the molecular-transport effects behind the shock (in the form of the 'shock slip') and the wall-slip effects are studied. In the flow regime where the shock-transition-zone thickness remains small compared to the shock radius of curvature, a quasi-one-dimensional shock structure under the Burnett/thirteen-moment approximation, as well as particulate/collisional models, can be consistently developed. The fully viscous version of the shock-layer model is shown to provide the crucial boundary condition downstream the shock in this case. The gas-kinetic basis of the continuum description for the flow behind the bow shock, and certain features affecting the non-equilibrium flow chemistry, are also discussed.

Mini-Sniffer II in Flight

Science.gov (United States)

1976-01-01

This photograph shows the second Mini-Sniffer undergoing flight testing over Rogers Dry Lake in Edwards, California. This version of the Mini-Sniffer lacked the canard of the original version and had wing tips and tail booms added. The Mini-Sniffer was a remotely controlled, propeller-driven vehicle developed at the NASA Flight Research Center (which became the Dryden Flight Research Center, Edwards, California, in 1976) as a potential platform to sample the upper atmosphere for pollution. The vehicle, flown from 1975 to 1977, was one of the earliest attempts by NASA to develop an aircraft that could sense turbulence and measure natural and human-produced atmospheric pollutants at altitudes above 80,000 feet with a variable-load propeller that was never flight-tested. Three Mini-Sniffer vehicles were built. The number 1 Mini-Sniffer vehicle had swept wings with a span of 18 feet and canards on the nose. It flew 12 flights with the gas-powered engine at low altitudes of around 2,500 feet. The number 1 vehicle was then modified into version number 2 by removing the canards and wing rudders and adding wing tips and tail booms. Twenty flights were made with this version, up to altitudes of 20,000 feet. The number 3 vehicle had a longer fuselage, was lighter in weight, and was powered by the non-air-breathing hydrazine engine designed by NASA's Johnson Space Center in Houston, Texas. This version was designed to fly a 25-pound payload to an altitude of 70,000 feet for one hour or to climb to 90,000 feet and glide back. The number 3 Mini-Sniffer made one flight to 20,000 feet and was not flown again because of a hydrazine leak problem. All three versions used a pusher propeller to free the nose area for an atmospheric-sampling payload. At various times the Mini-Sniffer has been considered for exploration in the carbon dioxide atmosphere of the planet Mars, where the gravity (38 percent of that on Earth) would reduce the horsepower needed for flight.

Hyper-X Research Vehicle - Artist Concept in Flight with Scramjet Engine Firing

Science.gov (United States)

1997-01-01

This is an artist's depiction of a Hyper-X research vehicle under scramjet power in free-flight following separation from its booster rocket. The X-43A was developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need

The Effect of Slipstream Obstructions on Air Propellers

Science.gov (United States)

Lesley, E P; Woods, B M

1924-01-01

The screw propeller on airplanes is usually placed near other objects, and hence its performance may be modified by them. Results of tests on propellers free from slip stream obstructions, both fore and aft, are therefore subject to correction, for the effect of such obstructions and the purpose of the investigation was to determine the effect upon the thrust and torque coefficients and efficiency, for previously tested air propellers, of obstructions placed in the slip stream, it being realized that such previous tests had been conducted under somewhat ideal conditions that are impracticable of realization in flight. Simple geometrical forms were used for the initial investigation. Such forms offered the advantage of easy, exact reproduction at another time or in other laboratories, and it was believed that the effects of obstructions usually encountered might be deduced or surmise from those chosen.

Terminal Sliding Mode Control with Unidirectional Auxiliary Surfaces for Hypersonic Vehicles Based on Adaptive Disturbance Observer

Directory of Open Access Journals (Sweden)

Naibao He

2015-01-01

Full Text Available A novel flight control scheme is proposed using the terminal sliding mode technique, unidirectional auxiliary surfaces and the disturbance observer model. These proposed dynamic attitude control systems can improve control performance of hypersonic vehicles despite uncertainties and external disturbances. The terminal attractor is employed to improve the convergence rate associated with the critical damping characteristics problem noted in short-period motions of hypersonic vehicles. The proposed robust attitude control scheme uses a dynamic terminal sliding mode with unidirectional auxiliary surfaces. The nonlinear disturbance observer is designed to estimate system uncertainties and external disturbances. The output of the disturbance observer aids the robust adaptive control scheme and improves robust attitude control performance. Finally, simulation results are presented to illustrate the effectiveness of the proposed terminal sliding mode with unidirectional auxiliary surfaces.

Hypersonic Navier-Stokes Comparisons to Orbiter Flight Data

Science.gov (United States)

Candler, Graham V.; Campbell, Charles H.

2010-01-01

During the STS-119 flight of Space Shuttle Discovery, two sets of surface temperature measurements were made. Under the HYTHIRM program3 quantitative thermal images of the windward side of the Orbiter with a were taken. In addition, the Boundary Layer Transition Flight Experiment 4 made thermocouple measurements at discrete locations on the Orbiter wind side. Most of these measurements were made downstream of a surface protuberance designed to trip the boundary layer to turbulent flow. In this paper, we use the US3D computational fluid dynamics code to simulate the Orbiter flow field at conditions corresponding to the STS-119 re-entry. We employ a standard two-temperature, five-species finite-rate model for high-temperature air, and the surface catalysis model of Stewart.1 This work is similar to the analysis of Wood et al . 2 except that we use a different approach for modeling turbulent flow. We use the one-equation Spalart-Allmaras turbulence model8 with compressibility corrections 9 and an approach for tripping the boundary layer at discrete locations. In general, the comparison between the simulations and flight data is remarkably good

Hypersonic Engine Leading Edge Experiments in a High Heat Flux, Supersonic Flow Environment

Science.gov (United States)

Gladden, Herbert J.; Melis, Matthew E.

1994-01-01

A major concern in advancing the state-of-the-art technologies for hypersonic vehicles is the development of an aeropropulsion system capable of withstanding the sustained high thermal loads expected during hypersonic flight. Three aerothermal load related concerns are the boundary layer transition from laminar to turbulent flow, articulating panel seals in high temperature environments, and strut (or cowl) leading edges with shock-on-shock interactions. A multidisciplinary approach is required to address these technical concerns. A hydrogen/oxygen rocket engine heat source has been developed at the NASA Lewis Research Center as one element in a series of facilities at national laboratories designed to experimentally evaluate the heat transfer and structural response of the strut (or cowl) leading edge. A recent experimental program conducted in this facility is discussed and related to cooling technology capability. The specific objective of the experiment discussed is to evaluate the erosion and oxidation characteristics of a coating on a cowl leading edge (or strut leading edge) in a supersonic, high heat flux environment. Heat transfer analyses of a similar leading edge concept cooled with gaseous hydrogen is included to demonstrate the complexity of the problem resulting from plastic deformation of the structures. Macro-photographic data from a coated leading edge model show progressive degradation over several thermal cycles at aerothermal conditions representative of high Mach number flight.

Theseus in Flight

Science.gov (United States)

1996-01-01

The twin pusher propeller-driven engines of the Theseus research aircraft can be clearly seen in this photo, taken during a 1996 research flight at NASA's Dryden Flight Research Center, Edwards, California. The Theseus aircraft, built and operated by Aurora Flight Sciences Corporation, Manassas, Virginia, was a unique aircraft flown at NASA's Dryden Flight Research Center, Edwards, California, under a cooperative agreement between NASA and Aurora. Dryden hosted the Theseus program, providing hangar space and range safety for flight testing. Aurora Flight Sciences was responsible for the actual flight testing, vehicle flight safety, and operation of the aircraft. The Theseus remotely piloted aircraft flew its maiden flight on May 24, 1996, at Dryden. During its sixth flight on November 12, 1996, Theseus experienced an in-flight structural failure that resulted in the loss of the aircraft. As of the beginning of the year 2000, Aurora had not rebuilt the aircraft. Theseus was built for NASA under an innovative, $4.9 million fixed-price contract by Aurora Flight Sciences Corporation and its partners, West Virginia University, Morgantown, West Virginia, and Fairmont State College, Fairmont, West Virginia. The twin-engine, unpiloted vehicle had a 140-foot wingspan, and was constructed largely of composite materials. Powered by two 80-horsepower, turbocharged piston engines that drove twin 9-foot-diameter propellers, Theseus was designed to fly autonomously at high altitudes, with takeoff and landing under the active control of a ground-based pilot in a ground control station 'cockpit.' With the potential ability to carry 700 pounds of science instruments to altitudes above 60,000 feet for durations of greater than 24 hours, Theseus was intended to support research in areas such as stratospheric ozone depletion and the atmospheric effects of future high-speed civil transport aircraft engines. Instruments carried aboard Theseus also would be able to validate satellite

Further analysis of MHD acceleration for a hypersonic wind tunnel

International Nuclear Information System (INIS)

Christiansen, M.J.; Schmidt, H.J.; Chapman, J.N.

1995-01-01

A previously completed MHD study of the use of an MHD accelerator with seeded air from a state-of-the-art arc heater, was generally hailed as showing that the system studied has some promise of meeting the most critical hypersonic testing requirements. However, some concerns existed about certain aspects of the results. This paper discusses some of these problems and presents analysis of potential solutions. Specifically the problems addressed are; reducing the amount of seed in the flow, reducing test chamber temperatures, and reducing the oxygen dissociation. Modeling techniques are used to study three design variables of the MHD accelerator. The accelerator channel inlet Mach number, the accelerator channel divergence angle, and the magnetic field strength are all studied. These variables are all optimized to meet the goals for seed, temperature, and dissociated oxygen reduction. The results of this paper are encouraging, showing that all three goals can be met. General relationships are observed as to how the design variables affect the performance of the MHD accelerator facility. This paper expands on the results presented in the UTSI report and further supports the feasibility of MHD acceleration as a means to provide hypersonic flight simulation

The Additional Error of Inertial Sensors Induced by Hypersonic Flight Conditions.

Science.gov (United States)

Karachun, Volodimir; Mel'nick, Viktorij; Korobiichuk, Igor; Nowicki, Michał; Szewczyk, Roman; Kobzar, Svitlana

2016-02-26

The emergence of hypersonic technology pose a new challenge for inertial navigation sensors, widely used in aerospace industry. The main problems are: extremely high temperatures, vibration of the fuselage, penetrating acoustic radiation and shock N-waves. The nature of the additional errors of the gyroscopic inertial sensor with hydrostatic suspension components under operating conditions generated by forced precession of the movable part of the suspension due to diffraction phenomena in acoustic fields is explained. The cause of the disturbing moments in the form of the Coriolis inertia forces during the transition of the suspension surface into the category of impedance is revealed. The boundaries of occurrence of the features on the resonance wave match are described. The values of the "false" angular velocity as a result of the elastic-stress state of suspension in the acoustic fields are determined.

Comparison of CFD Predictions with Shuttle Global Flight Thermal Imagery and Discrete Surface Measurements

Science.gov (United States)

Wood, William A.; Kleb, William L.; Tang, chun Y.; Palmer, Grant E.; Hyatt, Andrew J.; Wise, Adam J.; McCloud, Peter L.

2010-01-01

Surface temperature measurements from the STS-119 boundary-layer transition experiment on the space shuttle orbiter Discovery provide a rare opportunity to assess turbulent CFD models at hypersonic flight conditions. This flight data was acquired by on-board thermocouples and by infrared images taken off-board by the Hypersonic Thermodynamic Infrared Measurements (HYTHIRM) team, and is suitable for hypersonic CFD turbulence assessment between Mach 6 and 14. The primary assessment is for the Baldwin-Lomax and Cebeci-Smith algebraic turbulence models in the DPLR and LAURA CFD codes, respectively. A secondary assessment is made of the Shear-Stress Transport (SST) two-equation turbulence model in the DPLR code. Based upon surface temperature comparisons at eleven thermocouple locations, the algebraic-model turbulent CFD results average 4% lower than the measurements for Mach numbers less than 11. For Mach numbers greater than 11, the algebraic-model turbulent CFD results average 5% higher than the three available thermocouple measurements. Surface temperature predictions from the two SST cases were consistently 3 4% higher than the algebraic-model results. The thermocouple temperatures exhibit a change in trend with Mach number at about Mach 11; this trend is not reflected in the CFD results. Because the temperature trends from the turbulent CFD simulations and the flight data diverge above Mach 11, extrapolation of the turbulent CFD accuracy to higher Mach numbers is not recommended.

Iodine Hall Thruster Propellant Feed System for a CubeSat

Science.gov (United States)

Polzin, Kurt A.

2014-01-01

There has been significant work recently in the development of iodine-fed Hall thrusters for in-space propulsion applications.1 The use of iodine as a propellant provides many advantages over present xenon-gas-fed Hall thruster systems. Iodine is a solid at ambient temperature (no pressurization required) and has no special handling requirements, making it safe for secondary flight opportunities. It has exceptionally high ?I sp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing system level advantages over mid-term high power electric propulsion options. Iodine provides thrust and efficiency that are comparable to xenonfed Hall thrusters while operating in the same discharge current and voltage regime, making it possible to leverage the development of flight-qualified xenon Hall thruster power processing units for the iodine application. Work at MSFC is presently aimed at designing, integrating, and demonstrating a flight-like iodine feed system suitable for the Hall thruster application. This effort represents a significant advancement in state-of-the-art. Though Iodine thrusters have demonstrated high performance with mission enabling potential, a flight-like feed system has never been demonstrated and iodine compatible components do not yet exist. Presented in this paper is the end-to-end integrated feed system demonstration. The system includes a propellant tank with active feedback-control heating, fill and drain interfaces, latching and proportional flow control valves (PFCV), flow resistors, and flight-like CubeSat power and control electronics. Hardware is integrated into a CubeSat-sized structure, calibrated and tested under vacuum conditions, and operated under under hot-fire conditions using a Busek BHT-200 thruster designed for iodine. Performance of the system is evaluated thorugh accurate measurement of thrust and a calibrated of mass flow rate measurement, which is a function of

Aerothermodynamics of expert ballistic vehicle at hypersonic speeds

Science.gov (United States)

Kharitonov, A. M.; Adamov, N. P.; Chirkashenko, V. F.; Mazhul, I. I.; Shpak, S. I.; Shiplyuk, A. N.; Vasenyov, L. G.; Zvegintsev, V. I.; Muylaert, J. M.

2012-01-01

The European EXPErimental Re-entry Test bed (EXPERT) vehicle is intended for studying various basic phenomena, such as the boundary-layer transition on blunted bodies, real gas effects during shock wave/boundary layer interaction, and effect of surface catalycity. Another task is to develop methods for recalculating the results of windtunnel experiments to flight conditions. The EXPERT program implies large-scale preflight research, in particular, various calculations with the use of advanced numerical methods, experimental studies of the models in various wind tunnels, and comparative analysis of data obtained for possible extrapolation of data to in-flight conditions. The experimental studies are performed in various aerodynamic centers of Europe and Russia under contracts with ESA-ESTEC. In particular, extensive experiments are performed at the Von Karman Institute for Fluid Dynamics (VKI, Belgium) and also at the DLR aerospace center in Germany. At ITAM SB RAS, the experimental studies of the EXPERT model characteristic were performed under ISTC Projects 2109, 3151, and 3550, in the T-313 supersonic wind tunnel and AT-303 hypersonic wind tunnel.

Efficient Numerical Simulation of Aerothermoelastic Hypersonic Vehicles

Science.gov (United States)

Klock, Ryan J.

Hypersonic vehicles operate in a high-energy flight environment characterized by high dynamic pressures, high thermal loads, and non-equilibrium flow dynamics. This environment induces strong fluid, thermal, and structural dynamics interactions that are unique to this flight regime. If these vehicles are to be effectively designed and controlled, then a robust and intuitive understanding of each of these disciplines must be developed not only in isolation, but also when coupled. Limitations on scaling and the availability of adequate test facilities mean that physical investigation is infeasible. Ever growing computational power offers the ability to perform elaborate numerical simulations, but also has its own limitations. The state of the art in numerical simulation is either to create ever more high-fidelity physics models that do not couple well and require too much processing power to consider more than a few seconds of flight, or to use low-fidelity analytical models that can be tightly coupled and processed quickly, but do not represent realistic systems due to their simplifying assumptions. Reduced-order models offer a middle ground by distilling the dominant trends of high-fidelity training solutions into a form that can be quickly processed and more tightly coupled. This thesis presents a variably coupled, variable-fidelity, aerothermoelastic framework for the simulation and analysis of high-speed vehicle systems using analytical, reduced-order, and surrogate modeling techniques. Full launch-to-landing flights of complete vehicles are considered and used to define flight envelopes with aeroelastic, aerothermal, and thermoelastic limits, tune in-the-loop flight controllers, and inform future design considerations. A partitioned approach to vehicle simulation is considered in which regions dominated by particular combinations of processes are made separate from the overall solution and simulated by a specialized set of models to improve overall processing

Implementation Options for the PROPEL Electrodynamic Tether Demonstration Mission

Science.gov (United States)

Bilen, Sven G.; Johnson, C. Les; Gilchrist, Brian E.; Hoyt, Robert P.; Elder, Craig H.; Fuhrhop, Keith P.; Scadera, Michael; Stone, Nobie

2014-01-01

The PROPEL ("Propulsion using Electrodynamics") flight demonstration mission concept will demonstrate the use of an electrodynamic tether (EDT) for generating thrust, which will allow the propulsion system to overcome the limitations of the rocket equation. The mission concept has been developed by a team of government, industry, and academia partners led by NASA Marshall Space Flight Center (MSFC). PROPEL is being designed for versatility of the EDT system with multiple end users in mind and to be flexible with respect to platform. Previously, we reported on a comprehensive mission design for PROPEL with a mission duration of six months or longer with multiple mission goals including demonstration of significant boost, deboost, inclination change, and drag make-up activities. To explore a range of possible configurations, primarily driven by cost considerations, other mission concept designs have been pursued. In partnership with the NASA's Office of Chief Technologist (OCT) Game Changing Program, NASA MSFC Leadership, and the MSFC Advanced Concepts Office, a mission concept design was developed for a near-term EDT propulsion flight validation mission. The Electrodynamic Tether Propulsion Study (ETPS) defined an EDT propulsion system capable of very large delta-V for use on future missions developed by NASA, DoD, and commercial customers. To demonstrate the feasibility of an ETPS, the study focused on a space demonstration mission concept design with configuration of a pair of tethered satellite busses, one of which is the Japanese H-II Transfer Vehicle (HTV). The HTV would fly its standard ISS resupply mission. When resupply mission is complete, the ISS reconfigures and releases the HTV to perform the EDT experiment at safe orbital altitudes below the ISS. Though the focus of this particular mission concept design addresses a scenario involving the HTV or a similar vehicle, the propulsion system's capability is relevant to a number of applications, as noted above

Comparison of Analysis with Test for Static Loading of Two Hypersonic Inflatable Aerodynamic Decelerator Concepts

Science.gov (United States)

Lyle, Karen H.

2015-01-01

Acceptance of new spacecraft structural architectures and concepts requires validated design methods to minimize the expense involved with technology demonstration via flight-testing. Hypersonic Inflatable Aerodynamic Decelerator (HIAD) architectures are attractive for spacecraft deceleration because they are lightweight, store compactly, and utilize the atmosphere to decelerate a spacecraft during entry. However, designers are hesitant to include these inflatable approaches for large payloads or spacecraft because of the lack of flight validation. This publication summarizes results comparing analytical results with test data for two concepts subjected to representative entry, static loading. The level of agreement and ability to predict the load distribution is considered sufficient to enable analytical predictions to be used in the design process.

Robust Stability Clearance of Flight Control Law Based on Global Sensitivity Analysis

OpenAIRE

Ou, Liuli; Liu, Lei; Dong, Shuai; Wang, Yongji

2014-01-01

To validate the robust stability of the flight control system of hypersonic flight vehicle, which suffers from a large number of parametrical uncertainties, a new clearance framework based on structural singular value ( $\\mu $ ) theory and global uncertainty sensitivity analysis (SA) is proposed. In this framework, SA serves as the preprocess of uncertain model to be analysed to help engineers to determine which uncertainties affect the stability of the closed loop system more slightly. By ig...

Overview of hypersonic CFD code calibration studies

Science.gov (United States)

Miller, Charles G.

1987-01-01

The topics are presented in viewgraph form and include the following: definitions of computational fluid dynamics (CFD) code validation; climate in hypersonics and LaRC when first 'designed' CFD code calibration studied was initiated; methodology from the experimentalist's perspective; hypersonic facilities; measurement techniques; and CFD code calibration studies.

Preliminary Assessment of Using Gelled and Hybrid Propellant Propulsion for VTOL/SSTO Launch Systems

Science.gov (United States)

Palaszewski, Bryan; OLeary, Robert; Pelaccio, Dennis G.

1998-01-01

A novel, reusable, Vertical-Takeoff-and-Vertical-Takeoff-and-Landing, Single-Stage-to-Orbit (VTOL/SSTO) launch system concept, named AUGMENT-SSTO, is presented in this paper to help quantify the advantages of employing gelled and hybrid propellant propulsion system options for such applications. The launch vehicle system concept considered uses a highly coupled, main high performance liquid oxygen/liquid hydrogen (LO2/LH2) propulsion system, that is used only for launch, while a gelled or hybrid propellant propulsion system auxiliary propulsion system is used during final orbit insertion, major orbit maneuvering, and landing propulsive burn phases of flight. Using a gelled or hybrid propellant propulsion system for major orbit maneuver burns and landing has many advantages over conventional VTOL/SSTO concepts that use LO2/LH2 propulsion system(s) burns for all phases of flight. The applicability of three gelled propellant systems, O2/H2/Al, O2/RP-1/Al, and NTO/MMH/Al, and a state-of-the-art (SOA) hybrid propulsion system are examined in this study. Additionally, this paper addresses the applicability of a high performance gelled O2/H2 propulsion system to perform the primary, as well as the auxiliary propulsion system functions of the vehicle.

The Role of Guidance, Navigation, and Control in Hypersonic Vehicle Multidisciplinary Design and Optimization

Science.gov (United States)

Ouzts, Peter J.; Soloway, Donald I.; Moerder, Daniel D.; Wolpert, David H.; Benavides, Jose Victor

2009-01-01

Airbreathing hypersonic systems offer distinct performance advantages over rocket-based systems for space access vehicles. However, these performance advantages are dependent upon advances in current state-of-the-art technologies in many areas such as ram/scramjet propulsion integration, high temperature materials, aero-elastic structures, thermal protection systems, transition to hypersonics and hypersonic control elements within the framework of complex physics and new design methods. The complex interactions between elements of an airbreathing hypersonic vehicle represent a new paradigm in vehicle design to achieve the optimal performance necessary to meet space access mission objectives. In the past, guidance, navigation, and control (GNC) analysis often follows completion of the vehicle conceptual design process. Individual component groups design subsystems which are then integrated into a vehicle configuration. GNC is presented the task of developing control approaches to meet vehicle performance objectives given that configuration. This approach may be sufficient for vehicles where significant performance margins exist. However, for higher performance vehicles engaging the GNC discipline too late in the design cycle has been costly. For example, the X-29 experimental flight vehicle was built as a technology demonstrator. One of the many technologies to be demonstrated was the use of light-weight material composites for structural components. The use of light-weight materials increased the flexibility of the X- 29 beyond that of conventional metal alloy constructed aircraft. This effect was not considered when the vehicle control system was designed and built. The impact of this is that the control system did not have enough control authority to compensate for the effects of the first fundamental structural mode of the vehicle. As a result, the resulting pitch rate response of the vehicle was below specification and no post-design changes could recover the

Research on the Frequency Aliasing of Resistance Acceleration Guidance for Reentry Flight

Directory of Open Access Journals (Sweden)

Han Pengxin

2017-01-01

Full Text Available According to the special response of resistance acceleration during hypersonic reentry flight, different guidance frequency will result to very different flight and control response. The analysis model for the response of resistance acceleration to the attack angle and dynamic press is put forward respectively in this paper. And the frequency aliasing phenomenon of guidance is revealed. The simulation results to the same vehicle sufficiently substantiate the frequency aliasing of resistance acceleration during reentry guidance.

Liquid oxygen (LO2) propellant conditioning concept testing

Science.gov (United States)

Perry, Gretchen L. E.; Orth, Michael S.; Mehta, Gopal K.

1993-01-01

Testing of a simplified LO2 propellant conditioning concept for future expendable launch vehicles is discussed. Four different concepts are being investigated: no-bleed, low-bleed, use of a recirculation line, and He bubbling. A full-scale test article, which is a facsimile of a propellant feed duct with an attached section to simulate heat input from an LO2 turbopump, is to be tested at the Cold Flow Facility of the Marshall Space Flight Center West Test Area. Work to date includes: design and fabrication of the test article, design of the test facility and initial fabrication, development of a test matrix and test procedures, initial predictions of test output, and heat leak calibration and heat exchanger tests on the test articles.

Perseus A in Flight with Moon

Science.gov (United States)

1994-01-01

The Perseus A, a remotely-piloted, high-altitude research aircraft, is seen here framed against the moon and sky during a research mission at the Dryden Flight Research Center, Edwards, California in August 1994. Perseus B is a remotely piloted aircraft developed as a design-performance testbed under NASA's Environmental Research Aircraft and Sensor Technology (ERAST) project. Perseus is one of several flight vehicles involved in the ERAST project. A piston engine, propeller-powered aircraft, Perseus was designed and built by Aurora Flight Sciences Corporation, Manassas, Virginia. The objectives of Perseus B's ERAST flight tests have been to reach and maintain horizontal flight above altitudes of 60,000 feet and demonstrate the capability to fly missions lasting from 8 to 24 hours, depending on payload and altitude requirements. The Perseus B aircraft established an unofficial altitude record for a single-engine, propeller-driven, remotely piloted aircraft on June 27, 1998. It reached an altitude of 60,280 feet. In 1999, several modifications were made to the Perseus aircraft including engine, avionics, and flight-control-system improvements. These improvements were evaluated in a series of operational readiness and test missions at the Dryden Flight Research Center, Edwards, California. Perseus is a high-wing monoplane with a conventional tail design. Its narrow, straight, high-aspect-ratio wing is mounted atop the fuselage. The aircraft is pusher-designed with the propeller mounted in the rear. This design allows for interchangeable scientific-instrument payloads to be placed in the forward fuselage. The design also allows for unobstructed airflow to the sensors and other devices mounted in the payload compartment. The Perseus B that underwent test and development in 1999 was the third generation of the Perseus design, which began with the Perseus Proof-Of-Concept aircraft. Perseus was initially developed as part of NASA's Small High-Altitude Science Aircraft

Analysis of Aerodynamic Load of LSU-03 (LAPAN Surveillance UAV-03) Propeller

Science.gov (United States)

Rahmadi Nuranto, Awang; Jamaludin Fitroh, Ahmad; Syamsudin, Hendri

2018-04-01

The existing propeller of the LSU-03 aircraft is made of wood. To improve structural strength and obtain better mechanical properties, the propeller will be redesigned usingcomposite materials. It is necessary to simulate and analyze the design load. This research paper explainsthe simulation and analysis of aerodynamic load prior to structural design phase of composite propeller. Aerodynamic load calculations are performed using both the Blade Element Theory(BET) and the Computational Fluid Dynamic (CFD)simulation. The result of both methods show a close agreement, the different thrust forces is only 1.2 and 4.1% for two type mesh. Thus the distribution of aerodynamic loads along the surface of the propeller blades of the 3-D CFD simulation results are considered valid and ready to design the composite structure. TheCFD results is directly imported to the structure model using the Direct Import CFD / One-Way Fluid Structure Interaction (FSI) method. Design load of propeller is chosen at the flight condition at speed of 20 km/h at 7000 rpm.

Unstart Coupling Mechanism Analysis of Multiple-Modules Hypersonic Inlet

Directory of Open Access Journals (Sweden)

Jichao Hu

2013-01-01

Full Text Available The combination of multiplemodules in parallel manner is an important way to achieve the much higher thrust of scramjet engine. For the multiple-modules scramjet engine, when inlet unstarted oscillatory flow appears in a single-module engine due to high backpressure, how to interact with each module by massflow spillage, and whether inlet unstart occurs in other modules are important issues. The unstarted flowfield and coupling characteristic for a three-module hypersonic inlet caused by center module II and side module III were, conducted respectively. The results indicate that the other two hypersonic inlets are forced into unstarted flow when unstarted phenomenon appears on a single-module hypersonic inlet due to high backpressure, and the reversed flow in the isolator dominates the formation, expansion, shrinkage, and disappearance of the vortexes, and thus, it is the major factor of unstart coupling of multiple-modules hypersonic inlet. The coupling effect among multiple modules makes hypersonic inlet be more likely unstarted.

Unstart coupling mechanism analysis of multiple-modules hypersonic inlet.

Science.gov (United States)

Hu, Jichao; Chang, Juntao; Wang, Lei; Cao, Shibin; Bao, Wen

2013-01-01

The combination of multiplemodules in parallel manner is an important way to achieve the much higher thrust of scramjet engine. For the multiple-modules scramjet engine, when inlet unstarted oscillatory flow appears in a single-module engine due to high backpressure, how to interact with each module by massflow spillage, and whether inlet unstart occurs in other modules are important issues. The unstarted flowfield and coupling characteristic for a three-module hypersonic inlet caused by center module II and side module III were, conducted respectively. The results indicate that the other two hypersonic inlets are forced into unstarted flow when unstarted phenomenon appears on a single-module hypersonic inlet due to high backpressure, and the reversed flow in the isolator dominates the formation, expansion, shrinkage, and disappearance of the vortexes, and thus, it is the major factor of unstart coupling of multiple-modules hypersonic inlet. The coupling effect among multiple modules makes hypersonic inlet be more likely unstarted.

Flight Determination of the Longitudinal Stability Characteristics of a 0.133-Scale Rocket-Powered Model of the Consolidated Vultee XFY-1 Airplane without Propellers at Mach Numbers from 0.73 to 1.19, TED No. NACA DE 369

Science.gov (United States)

Hastings, Earl E., Jr.; Mitcham, Grady L.

1954-01-01

A flight test has been conducted to determine the longitudinal stability and control,characteristics of a 0.133-scale model of the Consolidated Vultee XFY-1 airplane without propellers for the Mach number range between 0.73 and 1.19.

Heat removing under hypersonic conditions

Directory of Open Access Journals (Sweden)

Semenov Mikhail E.

2016-01-01

Full Text Available In this paper we consider the heat transfer properties of the axially symmetric body with parabolic shape at hypersonic speeds (with a Mach number M > 5. We use the numerical methods based on the implicit difference scheme (Fedorenko method with direct method based on LU-decomposition and iterative method based on the Gauss-Seigel method. Our numerical results show that the heat removing process should be performed in accordance with the nonlinear law of heat distribution over the surface taking into account the hypersonic conditions of motion.

Cryogenic Propellant Storage and Transfer Technology Demonstration For Long Duration In-Space Missions

Science.gov (United States)

Meyer, Michael L.; Motil, Susan M.; Kortes, Trudy F.; Taylor, William J.; McRight, Patrick S.

2012-01-01

The high specific impulse of cryogenic propellants can provide a significant performance advantage for in-space transfer vehicles. The upper stages of the Saturn V and various commercial expendable launch vehicles have used liquid oxygen and liquid hydrogen propellants; however, the application of cryogenic propellants has been limited to relatively short duration missions due to the propensity of cryogens to absorb environmental heat resulting in fluid losses. Utilizing advanced cryogenic propellant technologies can enable the efficient use of high performance propellants for long duration missions. Crewed mission architectures for beyond low Earth orbit exploration can significantly benefit from this capability by developing realistic launch spacing for multiple launch missions, by prepositioning stages and by staging propellants at an in-space depot. The National Aeronautics and Space Administration through the Office of the Chief Technologist is formulating a Cryogenic Propellant Storage and Transfer Technology Demonstration Mission to mitigate the technical and programmatic risks of infusing these advanced technologies into the development of future cryogenic propellant stages or in-space propellant depots. NASA is seeking an innovative path for human space exploration, which strengthens the capability to extend human and robotic presence throughout the solar system. This mission will test and validate key cryogenic technological capabilities and has the objectives of demonstrating advanced thermal control technologies to minimize propellant loss during loiter, demonstrating robust operation in a microgravity environment, and demonstrating efficient propellant transfer on orbit. The status of the demonstration mission concept development, technology demonstration planning and technology maturation activities in preparation for flight system development are described.

Aerothermodynamic shape optimization of hypersonic blunt bodies

Science.gov (United States)

Eyi, Sinan; Yumuşak, Mine

2015-07-01

The aim of this study is to develop a reliable and efficient design tool that can be used in hypersonic flows. The flow analysis is based on the axisymmetric Euler/Navier-Stokes and finite-rate chemical reaction equations. The equations are coupled simultaneously and solved implicitly using Newton's method. The Jacobian matrix is evaluated analytically. A gradient-based numerical optimization is used. The adjoint method is utilized for sensitivity calculations. The objective of the design is to generate a hypersonic blunt geometry that produces the minimum drag with low aerodynamic heating. Bezier curves are used for geometry parameterization. The performances of the design optimization method are demonstrated for different hypersonic flow conditions.

DNS Studies of Transitional Hypersonic Reacting Flows Over 3-D Hypersonic Vehicles

National Research Council Canada - National Science Library

Zhong, Xiaolin

2003-01-01

The objectives of this research project are to develop CFD techniques and to conduct DNS studies of fundamental flow physics leading to boundary-layer instability and transition in hypersonic flows...

Evaluation of icing drag coefficient correlations applied to iced propeller performance prediction

Science.gov (United States)

Miller, Thomas L.; Shaw, R. J.; Korkan, K. D.

1987-01-01

Evaluation of three empirical icing drag coefficient correlations is accomplished through application to a set of propeller icing data. The various correlations represent the best means currently available for relating drag rise to various flight and atmospheric conditions for both fixed-wing and rotating airfoils, and the work presented here ilustrates and evaluates one such application of the latter case. The origins of each of the correlations are discussed, and their apparent capabilities and limitations are summarized. These correlations have been made to be an integral part of a computer code, ICEPERF, which has been designed to calculate iced propeller performance. Comparison with experimental propeller icing data shows generally good agreement, with the quality of the predicted results seen to be directly related to the radial icing extent of each case. The code's capability to properly predict thrust coefficient, power coefficient, and propeller efficiency is shown to be strongly dependent on the choice of correlation selected, as well as upon proper specificatioon of radial icing extent.

Experimental Investigation of Hypersonic Flow and Plasma Aerodynamic Actuation Interaction

International Nuclear Information System (INIS)

Sun Quan; Cheng Bangqin; Li Yinghong; Cui Wei; Yu Yonggui; Jie Junhun

2013-01-01

For hypersonic flow, it was found that the most effective plasma actuator is derived from an electromagnetic perturbation. An experimental study was performed between hypersonic flow and plasma aerodynamic actuation interaction in a hypersonic shock tunnel, in which a Mach number of 7 was reached. The plasma discharging characteristic was acquired in static flows. In a hypersonic flow, the flow field can affect the plasma discharging characteristics. DC discharging without magnetic force is unstable, and the discharge channel cannot be maintained. When there is a magnetic field, the energy consumption of the plasma source is approximately three to four times larger than that without a magnetic field, and at the same time plasma discharge can also affect the hypersonic flow field. Through schlieren pictures and pressure measurement, it was found that plasma discharging could induce shockwaves and change the total pressure and wall pressure of the flow field

A Comprehensive Analysis of the X-15 Flight 3-65 Accident

Science.gov (United States)

Dennehy, Cornelius J.; Orr, Jeb S.; Barshi, Immanuel; Statler, Irving C.

2014-01-01

The November 15, 1967, loss of X-15 Flight 3-65-97 (hereafter referred to as Flight 3-65) was a unique incident in that it was the first and only aerospace flight accident involving loss of crew on a vehicle with an adaptive flight control system (AFCS). In addition, Flight 3-65 remains the only incidence of a single-pilot departure from controlled flight of a manned entry vehicle in a hypersonic flight regime. To mitigate risk to emerging aerospace systems, the NASA Engineering and Safety Center (NESC) proposed a comprehensive review of this accident. The goal of the assessment was to resolve lingering questions regarding the failure modes of the aircraft systems (including the AFCS) and thoroughly analyze the interactions among the human agents and autonomous systems that contributed to the loss of the pilot and aircraft. This document contains the outcome of the accident review.

An improved flux-split algorithm applied to hypersonic flows in chemical equilibrium

Science.gov (United States)

Palmer, Grant

1988-01-01

An explicit, finite-difference, shock-capturing numerical algorithm is presented and applied to hypersonic flows assumed to be in thermochemical equilibrium. Real-gas chemistry is either loosely coupled to the gasdynamics by way of a Gibbs free energy minimization package or fully coupled using species mass conservation equations with finite-rate chemical reactions. A scheme is developed that maintains stability in the explicit, finite-rate formulation while allowing relatively high time steps. The codes use flux vector splitting to difference the inviscid fluxes and employ real-gas corrections to viscosity and thermal conductivity. Numerical results are compared against existing ballistic range and flight data. Flows about complex geometries are also computed.

Improved Dutch Roll Approximation for Hypersonic Vehicle

Directory of Open Access Journals (Sweden)

Liang-Liang Yin

2014-06-01

Full Text Available An improved dutch roll approximation for hypersonic vehicle is presented. From the new approximations, the dutch roll frequency is shown to be a function of the stability axis yaw stability and the dutch roll damping is mainly effected by the roll damping ratio. In additional, an important parameter called roll-to-yaw ratio is obtained to describe the dutch roll mode. Solution shows that large-roll-to-yaw ratio is the generate character of hypersonic vehicle, which results the large error for the practical approximation. Predictions from the literal approximations derived in this paper are compared with actual numerical values for s example hypersonic vehicle, results show the approximations work well and the error is below 10 %.

Hypersonic vehicle control law development using H(infinity) and micron-synthesis

Science.gov (United States)

Gregory, Irene M.; Mcminn, John D.; Shaughnessy, John D.; Chowdhry, Rajiv S.

1993-01-01

Hypersonic vehicle control law development using H(infinity) and mu-synthesis is discussed. Airbreathing SSTO vehicles has a mutli-faceted mission that includes orbital operations, as well as re-entry and descent culminating in horizontal landing. However, the most challenging part of the operations is the ascent to orbit. The airbreathing propulsion requires lengthy atmospheric flight that may last as long as 30 minutes and take the vehicle half way around the globe. The vehicles's ascent is characterized by tight payload to orbit margins which translate into minimum fuel orbit as the performance criteria. Issues discussed include: SSTO airbreathing vehicle issues; control system performance requirements; robust control law framework; H(infinity) controller frequency analysis; and mu controller frequency analysis.

A Vision of Quantitative Imaging Technology for Validation of Advanced Flight Technologies

Science.gov (United States)

Horvath, Thomas J.; Kerns, Robert V.; Jones, Kenneth M.; Grinstead, Jay H.; Schwartz, Richard J.; Gibson, David M.; Taylor, Jeff C.; Tack, Steve; Dantowitz, Ronald F.

2011-01-01

Flight-testing is traditionally an expensive but critical element in the development and ultimate validation and certification of technologies destined for future operational capabilities. Measurements obtained in relevant flight environments also provide unique opportunities to observe flow phenomenon that are often beyond the capabilities of ground testing facilities and computational tools to simulate or duplicate. However, the challenges of minimizing vehicle weight and internal complexity as well as instrumentation bandwidth limitations often restrict the ability to make high-density, in-situ measurements with discrete sensors. Remote imaging offers a potential opportunity to noninvasively obtain such flight data in a complementary fashion. The NASA Hypersonic Thermodynamic Infrared Measurements Project has demonstrated such a capability to obtain calibrated thermal imagery on a hypersonic vehicle in flight. Through the application of existing and accessible technologies, the acreage surface temperature of the Shuttle lower surface was measured during reentry. Future hypersonic cruise vehicles, launcher configurations and reentry vehicles will, however, challenge current remote imaging capability. As NASA embarks on the design and deployment of a new Space Launch System architecture for access beyond earth orbit (and the commercial sector focused on low earth orbit), an opportunity exists to implement an imagery system and its supporting infrastructure that provides sufficient flexibility to incorporate changing technology to address the future needs of the flight test community. A long term vision is offered that supports the application of advanced multi-waveband sensing technology to aid in the development of future aerospace systems and critical technologies to enable highly responsive vehicle operations across the aerospace continuum, spanning launch, reusable space access and global reach. Motivations for development of an Agency level imagery

Perseus B over Edwards AFB on a Development Flight

Science.gov (United States)

1998-01-01

A long, slender wing and a pusher propeller at the rear characterize the Perseus B remotely-piloted research aircraft, seen here during a test flight in April1998. Perseus B is a remotely piloted aircraft developed as a design-performance testbed under NASA's Environmental Research Aircraft and Sensor Technology (ERAST) project. Perseus is one of several flight vehicles involved in the ERAST project. A piston engine, propeller-powered aircraft, Perseus was designed and built by Aurora Flight Sciences Corporation, Manassas, Virginia. The objectives of Perseus B's ERAST flight tests have been to reach and maintain horizontal flight above altitudes of 60,000 feet and demonstrate the capability to fly missions lasting from 8 to 24 hours, depending on payload and altitude requirements. The Perseus B aircraft established an unofficial altitude record for a single-engine, propeller-driven, remotely piloted aircraft on June 27, 1998. It reached an altitude of 60,280 feet. In 1999, several modifications were made to the Perseus aircraft including engine, avionics, and flight-control-system improvements. These improvements were evaluated in a series of operational readiness and test missions at the Dryden Flight Research Center, Edwards, California. Perseus is a high-wing monoplane with a conventional tail design. Its narrow, straight, high-aspect-ratio wing is mounted atop the fuselage. The aircraft is pusher-designed with the propeller mounted in the rear. This design allows for interchangeable scientific-instrument payloads to be placed in the forward fuselage. The design also allows for unobstructed airflow to the sensors and other devices mounted in the payload compartment. The Perseus B that underwent test and development in 1999 was the third generation of the Perseus design, which began with the Perseus Proof-Of-Concept aircraft. Perseus was initially developed as part of NASA's Small High-Altitude Science Aircraft (SHASA) program, which later evolved into the ERAST

An Entry Flight Controls Analysis for a Reusable Launch Vehicle

Science.gov (United States)

Calhoun, Philip

2000-01-01

The NASA Langley Research Center has been performing studies to address the feasibility of various single-stage to orbit concepts for use by NASA and the commercial launch industry to provide a lower cost access to space. Some work on the conceptual design of a typical lifting body concept vehicle, designated VentureStar(sup TM) has been conducted in cooperation with the Lockheed Martin Skunk Works. This paper will address the results of a preliminary flight controls assessment of this vehicle concept during the atmospheric entry phase of flight. The work includes control analysis from hypersonic flight at the atmospheric entry through supersonic speeds to final approach and landing at subsonic conditions. The requirements of the flight control effectors are determined over the full range of entry vehicle Mach number conditions. The analysis was performed for a typical maximum crossrange entry trajectory utilizing angle of attack to limit entry heating and providing for energy management, and bank angle to modulation of the lift vector to provide downrange and crossrange capability to fly the vehicle to a specified landing site. Sensitivity of the vehicle open and closed loop characteristics to CG location, control surface mixing strategy and wind gusts are included in the results. An alternative control surface mixing strategy utilizing a reverse aileron technique demonstrated a significant reduction in RCS torque and fuel required to perform bank maneuvers during entry. The results of the control analysis revealed challenges for an early vehicle configuration in the areas of hypersonic pitch trim and subsonic longitudinal controllability.

Conjugate Heat Transfer Study in Hypersonic Flows

Science.gov (United States)

Sahoo, Niranjan; Kulkarni, Vinayak; Peetala, Ravi Kumar

2018-04-01

Coupled and decoupled conjugate heat transfer (CHT) studies are carried out to imitate experimental studies for heat transfer measurement in hypersonic flow regime. The finite volume based solvers are used for analyzing the heat interaction between fluid and solid domains. Temperature and surface heat flux signals are predicted by both coupled and decoupled CHT analysis techniques for hypersonic Mach numbers. These two methodologies are also used to study the effect of different wall materials on surface parameters. Effectiveness of these CHT solvers has been verified for the inverse problem of wall heat flux recovery using various techniques reported in the literature. Both coupled and decoupled CHT techniques are seen to be equally useful for prediction of local temperature and heat flux signals prior to the experiments in hypersonic flows.

CFD on hypersonic flow geometries with aeroheating

Science.gov (United States)

Sohail, Muhammad Amjad; Chao, Yan; Hui, Zhang Hui; Ullah, Rizwan

2012-11-01

The hypersonic flowfield around a blunted cone and cone-flare exhibits some of the major features of the flows around space vehicles, e.g. a detached bow shock in the stagnation region and the oblique shock wave/boundary layer interaction at the cone-flare junction. The shock wave/boundary layer interaction can produce a region of separated flow. This phenomenon may occur, for example, at the upstream-facing corner formed by a deflected control surface on a hypersonic entry vehicle, where the length of separation has implications for control effectiveness. Computational fluid-dynamics results are presented to show the flowfield around a blunted cone and cone-flare configurations in hypersonic flow with separation. This problem is of particular interest since it features most of the aspects of the hypersonic flow around planetary entry vehicles. The region between the cone and the flare is particularly critical with respect to the evaluation of the surface pressure and heat flux with aeroheating. Indeed, flow separation is induced by the shock wave boundary layer interaction, with subsequent flow reattachment, that can dramatically enhance the surface heat transfer. The exact determination of the extension of the recirculation zone is a particularly delicate task for numerical codes. Laminar flow and turbulent computations have been carried out using a full Navier-Stokes solver, with freestream conditions provided by the experimental data obtained at Mach 6, 8, and 16.34 wind tunnel. The numerical results are compared with the measured pressure and surface heat flux distributions in the wind tunnel and a good agreement is found, especially on the length of the recirculation region and location of shock waves. The critical physics of entropy layer, boundary layers, boundary layers and shock wave interaction and flow behind shock are properly captured and elaborated.. Hypersonic flows are characterized by high Mach number and high total enthalpy. An elevated

Hypersonic Tunnel Facility (HTF)

Data.gov (United States)

Federal Laboratory Consortium — The Hypersonic Tunnel Facility (HTF) is a blow-down, non-vitiated (clean air) free-jet wind tunnel capable of testing large-scale, propulsion systems at Mach 5, 6,...

Space Transportation Infrastructure Supported By Propellant Depots

Science.gov (United States)

Smitherman, David; Woodcock, Gordon

2012-01-01

A space transportation infrastructure is described that utilizes propellant depot servicing platforms to support all foreseeable missions in the Earth-Moon vicinity and deep space out to Mars. The infrastructure utilizes current expendable launch vehicle (ELV) systems such as the Delta IV Heavy, Atlas V, and Falcon 9, for all crew, cargo, and propellant launches to orbit. Propellant launches are made to Low-Earth-Orbit (LEO) Depot and an Earth-Moon Lagrange Point 1 (L1) Depot to support a new reusable in-space transportation vehicles. The LEO Depot supports missions to Geosynchronous Earth Orbit (GEO) for satellite servicing and to L1 for L1 Depot missions. The L1 Depot supports Lunar, Earth-Sun L2 (ESL2), Asteroid and Mars Missions. New vehicle design concepts are presented that can be launched on current 5 meter diameter ELV systems. These new reusable vehicle concepts include a Crew Transfer Vehicle (CTV) for crew transportation between the LEO Depot, L1 Depot and missions beyond L1; a new reusable lunar lander for crew transportation between the L1 Depot and the lunar surface; and Mars orbital Depot are based on International Space Station (ISS) heritage hardware. Data provided includes the number of launches required for each mission utilizing current ELV systems (Delta IV Heavy or equivalent) and the approximate vehicle masses and propellant requirements. Also included is a discussion on affordability with ideas on technologies that could reduce the number of launches required and thoughts on how this infrastructure include competitive bidding for ELV flights and propellant services, developments of new reusable in-space vehicles and development of a multiuse infrastructure that can support many government and commercial missions simultaneously.

Combustion Efficiency, Flameout Operability Limits and General Design Optimization for Integrated Ramjet-Scramjet Hypersonic Vehicles

Science.gov (United States)

Mbagwu, Chukwuka Chijindu

High speed, air-breathing hypersonic vehicles encounter a varied range of engine and operating conditions traveling along cruise/ascent missions at high altitudes and dynamic pressures. Variations of ambient pressure, temperature, Mach number, and dynamic pressure can affect the combustion conditions in conflicting ways. Computations were performed to understand propulsion tradeoffs that occur when a hypersonic vehicle travels along an ascent trajectory. Proper Orthogonal Decomposition methods were applied for the reduction of flamelet chemistry data in an improved combustor model. Two operability limits are set by requirements that combustion efficiency exceed selected minima and flameout be avoided. A method for flameout prediction based on empirical Damkohler number measurements is presented. Operability limits are plotted that define allowable flight corridors on an altitude versus flight Mach number performance map; fixed-acceleration ascent trajectories were considered for this study. Several design rules are also presented for a hypersonic waverider with a dual-mode scramjet engine. Focus is placed on ''vehicle integration" design, differing from previous ''propulsion-oriented" design optimization. The well-designed waverider falls between that of an aircraft (high lift-to-drag ratio) and a rocket (high thrust-to-drag ratio). 84 variations of an X-43-like vehicle were run using the MASIV scramjet reduced order model to examine performance tradeoffs. Informed by the vehicle design study, variable-acceleration trajectory optimization was performed for three constant dynamic pressures ascents. Computed flameout operability limits were implemented as additional constraints to the optimization problem. The Michigan-AFRL Scramjet In-Vehicle (MASIV) waverider model includes finite-rate chemistry, applied scaling laws for 3-D turbulent mixing, ram-scram transition and an empirical value of the flameout Damkohler number. A reduced-order modeling approach is justified

Nonlinear robust control of hypersonic aircrafts with interactions between flight dynamics and propulsion systems.

Science.gov (United States)

Li, Zhaoying; Zhou, Wenjie; Liu, Hao

2016-09-01

This paper addresses the nonlinear robust tracking controller design problem for hypersonic vehicles. This problem is challenging due to strong coupling between the aerodynamics and the propulsion system, and the uncertainties involved in the vehicle dynamics including parametric uncertainties, unmodeled model uncertainties, and external disturbances. By utilizing the feedback linearization technique, a linear tracking error system is established with prescribed references. For the linear model, a robust controller is proposed based on the signal compensation theory to guarantee that the tracking error dynamics is robustly stable. Numerical simulation results are given to show the advantages of the proposed nonlinear robust control method, compared to the robust loop-shaping control approach. Copyright © 2016 ISA. Published by Elsevier Ltd. All rights reserved.

Hypersonic modes in nanophononic semiconductors.

Science.gov (United States)

Hepplestone, S P; Srivastava, G P

2008-09-05

Frequency gaps and negative group velocities of hypersonic phonon modes in periodically arranged composite semiconductors are presented. Trends and criteria for phononic gaps are discussed using a variety of atomic-level theoretical approaches. From our calculations, the possibility of achieving semiconductor-based one-dimensional phononic structures is established. We present results of the location and size of gaps, as well as negative group velocities of phonon modes in such structures. In addition to reproducing the results of recent measurements of the locations of the band gaps in the nanosized Si/Si{0.4}Ge{0.6} superlattice, we show that such a system is a true one-dimensional hypersonic phononic crystal.

Optimal Control of Hypersonic Planning Maneuvers Based on Pontryagin’s Maximum Principle

Directory of Open Access Journals (Sweden)

A. Yu. Melnikov

2015-01-01

Full Text Available The work objective is the synthesis of simple analytical formula of the optimal roll angle of hypersonic gliding vehicles for conditions of quasi-horizontal motion, allowing its practical implementation in onboard control algorithms.The introduction justifies relevance, formulates basic control tasks, and describes a history of scientific research and achievements in the field concerned. The author reveals a common disadvantage of the other authors’ methods, i.e. the problem of practical implementation in onboard control algorithms.The similar tasks of hypersonic maneuvers are systemized according to the type of maneuver, control parameters and limitations.In the statement of the problem the glider launched horizontally with a suborbital speed glides passive in the static Atmosphere on a spherical surface of constant radius in the Central field of gravitation.The work specifies a system of equations of motion in the inertial spherical coordinate system, sets the limits on the roll angle and optimization criteria at the end of the flight: high speed or azimuth and the minimum distances to the specified geocentric points.The solution.1 A system of equations of motion is transformed by replacing the time argument with another independent argument – the normal equilibrium overload. The Hamiltonian and the equations of mated parameters are obtained using the Pontryagin’s maximum principle. The number of equations of motion and mated vector is reduced.2 The mated parameters were expressed by formulas using current movement parameters. The formulas are proved through differentiation and substitution in the equations of motion.3 The Formula of optimal roll-position control by condition of maximum is obtained. After substitution of mated parameters, the insertion of constants, and trigonometric transformations the Formula of the optimal roll angle is obtained as functions of the current parameters of motion.The roll angle is expressed as the ratio

Effect of Propellant Composition to the Temperature Sensitivity of Composite Propellant

International Nuclear Information System (INIS)

Aziz, Amir; Mamat, Rizalman; Amin, Makeen; Wan Ali, Wan Khairuddin

2012-01-01

The propellant composition is one of several parameter that influencing the temperature sensitivity of composite propellant. In this paper, experimental investigation of temperature sensitivity in burning rate of composite propellant was conducted. Four sets of different propellant compositions had been prepared with the combination of ammonium perchlorate (AP) as an oxidizer, aluminum (Al) as fuel and hydroxy-terminated polybutadiene (HTPB) as fuel and binder. For each mixture, HTPB binder was fixed at 15% and cured with isophorone diisocyanate (IPDI). By varying AP and Al, the effect of oxidizer- fuel mixture ratio (O/F) on the whole propellant can be determined. The propellant strands were manufactured using compression molded method and burnt in a strand burner using wire technique over a range of pressure from 1 atm to 31 atm. The results obtained shows that the temperature sensitivity, a, increases with increasing O/F. Propellant p80 which has O/F ratio of 80/20 gives the highest value of temperature sensitivity which is 1.687. The results shows that the propellant composition has significant effect on the temperature sensitivity of composite propellant

Hypersonic Transition Analysis for HIFiRE Experiments

Science.gov (United States)

Li, Fei; Choudhari, Meelan; Chang, Chau-Lyan; Kimmel, Roger; Adamczak, David; Smith, Mark

2012-01-01

The HIFiRE-1 flight experiment provided a valuable database pertaining to boundary layer transition over a 7-degree half-angle, circular cone model from supersonic to hypersonic Mach numbers, and a range of Reynolds numbers and angles of attack. This paper reports selected findings from the ongoing computational analysis of the measured in-flight transition behavior. Transition during the ascent phase at nearly zero degree angle of attack is dominated by second mode instabilities except in the vicinity of the cone meridian where a roughness element was placed midway along the length of the cone. The growth of first mode instabilities is found to be weak at all trajectory points analyzed from the ascent phase. For times less than approximately 18.5 seconds into the flight, the peak amplification ratio for second mode disturbances is sufficiently small because of the lower Mach numbers at earlier times, so that the transition behavior inferred from the measurements is attributed to an unknown physical mechanism, potentially related to step discontinuities in surface height near the locations of a change in the surface material. Based on the time histories of temperature and/or heat flux at transducer locations within the aft portion of the cone, the onset of transition correlated with a linear N-factor, based on parabolized stability equations, of approximately 13. Due to the large angles of attack during the re-entry phase, crossflow instability may play a significant role in transition. Computations also indicate the presence of pronounced crossflow separation over a significant portion of the trajectory segment that is relevant to transition analysis.

Perseus A, Part of the ERAST Program, in Flight

Science.gov (United States)

1993-01-01

The Perseus A remotely-piloted research vehicle flies low over Rogers Dry Lake on its maiden voyage Dec. 21, 1993, at the Dryden Flight Research Center, Edwards, California. The Perseus, designed and built by Aurora Flight Sciences Corp., was towed into the air by a ground vehicle. At about 700 ft. the aircraft was released and the engine turned the propeller to take the plane to its desired altitude. Perseus B is a remotely piloted aircraft developed as a design-performance testbed under NASA's Environmental Research Aircraft and Sensor Technology (ERAST) project. Perseus is one of several flight vehicles involved in the ERAST project. A piston engine, propeller-powered aircraft, Perseus was designed and built by Aurora Flight Sciences Corporation, Manassas, Virginia. The objectives of Perseus B's ERAST flight tests have been to reach and maintain horizontal flight above altitudes of 60,000 feet and demonstrate the capability to fly missions lasting from 8 to 24 hours, depending on payload and altitude requirements. The Perseus B aircraft established an unofficial altitude record for a single-engine, propeller-driven, remotely piloted aircraft on June 27, 1998. It reached an altitude of 60,280 feet. In 1999, several modifications were made to the Perseus aircraft including engine, avionics, and flight-control-system improvements. These improvements were evaluated in a series of operational readiness and test missions at the Dryden Flight Research Center, Edwards, California. Perseus is a high-wing monoplane with a conventional tail design. Its narrow, straight, high-aspect-ratio wing is mounted atop the fuselage. The aircraft is pusher-designed with the propeller mounted in the rear. This design allows for interchangeable scientific-instrument payloads to be placed in the forward fuselage. The design also allows for unobstructed airflow to the sensors and other devices mounted in the payload compartment. The Perseus B that underwent test and development in 1999

Liquid Oxygen Propellant Densification Unit Ground Tested With a Large-Scale Flight-Weight Tank for the X-33 Reusable Launch Vehicle

Science.gov (United States)

Tomsik, Thomas M.

2002-01-01

Propellant densification has been identified as a critical technology in the development of single-stage-to-orbit reusable launch vehicles. Technology to create supercooled high-density liquid oxygen (LO2) and liquid hydrogen (LH2) is a key means to lowering launch vehicle costs. The densification of cryogenic propellants through subcooling allows 8 to 10 percent more propellant mass to be stored in a given unit volume, thereby improving the launch vehicle's overall performance. This allows for higher propellant mass fractions than would be possible with conventional normal boiling point cryogenic propellants, considering the normal boiling point of LO2 and LH2.

Heat Transfer Analysis of Thermal Protection Structures for Hypersonic Vehicles

Science.gov (United States)

Zhou, Chen; Wang, Zhijin; Hou, Tianjiao

2017-11-01

This research aims to develop an analytical approach to study the heat transfer problem of thermal protection systems (TPS) for hypersonic vehicles. Laplace transform and integral method are used to describe the temperature distribution through the TPS subject to aerodynamic heating during flight. Time-dependent incident heat flux is also taken into account. Two different cases with heat flux and radiation boundary conditions are studied and discussed. The results are compared with those obtained by finite element analyses and show a good agreement. Although temperature profiles of such problems can be readily accessed via numerical simulations, analytical solutions give a greater insight into the physical essence of the heat transfer problem. Furthermore, with the analytical approach, rapid thermal analyses and even thermal optimization can be achieved during the preliminary TPS design.

Improved State Feedback H∞ Control for Flexible Air-Breathing Hypersonic Vehicles on LMI Approach

Directory of Open Access Journals (Sweden)

Zhang Xue

2017-01-01

Full Text Available Focusing on a nonlinear longitudinal dynamical model for Air-breathing Hypersonic Flight Vehicles (AHFV, a linearized model on a nominal trim condition is proposed. To stabilize the flight of an AHFV in the presence of external disturbances and actuator uncertainties, a state feedback H∞ control is designed. With bounds on the uncertainties, a feedback stabilization problem is converted to an optimal control problem and the cost function is minimized by solving a set of linear matrix inequalities. Since uncertainties in the design of AHFV are inevitable, to make a comparison, a general H∞ robust controller is constructed by only considering the disturbances firstly. Then the results are extended by incorporating the actual existing uncertainties as well as the external disturbances in the AHFV system. Numerical simulation shows that the controller, which takes both disturbances and uncertainties into account, can effectively stabilize the AHFV system.

Development of an aerodynamic measurement system for hypersonic rarefied flows.

Science.gov (United States)

Ozawa, T; Fujita, K; Suzuki, T

2015-01-01

A hypersonic rarefied wind tunnel (HRWT) has lately been developed at Japan Aerospace Exploration Agency in order to improve the prediction of rarefied aerodynamics. Flow characteristics of hypersonic rarefied flows have been investigated experimentally and numerically. By conducting dynamic pressure measurements with pendulous models and pitot pressure measurements, we have probed flow characteristics in the test section. We have also improved understandings of hypersonic rarefied flows by integrating a numerical approach with the HRWT measurement. The development of the integration scheme between HRWT and numerical approach enables us to estimate the hypersonic rarefied flow characteristics as well as the direct measurement of rarefied aerodynamics. Consequently, this wind tunnel is capable of generating 25 mm-core flows with the free stream Mach number greater than 10 and Knudsen number greater than 0.1.

Experimental Flight Characterization of a Canard-Controlled, Subsonic Missile

Science.gov (United States)

2017-08-01

Frankfort mount was used to place the gun between stations 5 and 6 in the spark range building. The propelling charge was contained within a plastic ...additive manufacturing housing and used 83 g of M38 propellant with about 2 g of black powder wrapped around an electric match. A shot-start link was...ARL-TR-8086 ● AUG 2017 US Army Research Laboratory Experimental Flight Characterization of a Canard-Controlled, Subsonic Missile

Design of adaptive switching control for hypersonic aircraft

Directory of Open Access Journals (Sweden)

Xin Jiao

2015-10-01

Full Text Available This article proposes a novel adaptive switching control of hypersonic aircraft based on type-2 Takagi–Sugeno–Kang fuzzy sliding mode control and focuses on the problem of stability and smoothness in the switching process. This method uses full-state feedback to linearize the nonlinear model of hypersonic aircraft. Combining the interval type-2 Takagi–Sugeno–Kang fuzzy approach with sliding mode control keeps the adaptive switching process stable and smooth. For rapid stabilization of the system, the adaptive laws use a direct constructive Lyapunov analysis together with an established type-2 Takagi–Sugeno–Kang fuzzy logic system. Simulation results indicate that the proposed control scheme can maintain the stability and smoothness of switching process for the hypersonic aircraft.

14 CFR 121.643 - Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Supplemental operations.

Science.gov (United States)

2010-01-01

... operating nontransport category airplanes type certificated after December 31, 1964, to fly for 30 minutes...-powered airplanes: Supplemental operations. 121.643 Section 121.643 Aeronautics and Space FEDERAL AVIATION... Flight Release Rules § 121.643 Fuel supply: Nonturbine and turbo-propeller-powered airplanes...

Fuels and Space Propellants for Reusable Launch Vehicles: A Small Business Innovation Research Topic and Its Commercial Vision

Science.gov (United States)

Palaszewski, Bryan A.

1997-01-01

Under its Small Business Innovation Research (SBIR) program (and with NASA Headquarters support), the NASA Lewis Research Center has initiated a topic entitled "Fuels and Space Propellants for Reusable Launch Vehicles." The aim of this project would be to assist in demonstrating and then commercializing new rocket propellants that are safer and more environmentally sound and that make space operations easier. Soon it will be possible to commercialize many new propellants and their related component technologies because of the large investments being made throughout the Government in rocket propellants and the technologies for using them. This article discusses the commercial vision for these fuels and propellants, the potential for these propellants to reduce space access costs, the options for commercial development, and the benefits to nonaerospace industries. This SBIR topic is designed to foster the development of propellants that provide improved safety, less environmental impact, higher density, higher I(sub sp), and simpler vehicle operations. In the development of aeronautics and space technology, there have been limits to vehicle performance imposed by traditionally used propellants and fuels. Increases in performance are possible with either increased propellant specific impulse, increased density, or both. Flight system safety will also be increased by the use of denser, more viscous propellants and fuels.

Hyper-X Mach 7 Scramjet Design, Ground Test and Flight Results

Science.gov (United States)

Ferlemann, Shelly M.; McClinton, Charles R.; Rock, Ken E.; Voland, Randy T.

2005-01-01

The successful Mach 7 flight test of the Hyper-X (X-43) research vehicle has provided the major, essential demonstration of the capability of the airframe integrated scramjet engine. This flight was a crucial first step toward realizing the potential for airbreathing hypersonic propulsion for application to space launch vehicles. However, it is not sufficient to have just achieved a successful flight. The more useful knowledge gained from the flight is how well the prediction methods matched the actual test results in order to have confidence that these methods can be applied to the design of other scramjet engines and powered vehicles. The propulsion predictions for the Mach 7 flight test were calculated using the computer code, SRGULL, with input from computational fluid dynamics (CFD) and wind tunnel tests. This paper will discuss the evolution of the Mach 7 Hyper-X engine, ground wind tunnel experiments, propulsion prediction methodology, flight results and validation of design methods.

Issues Associated with a Hypersonic Maglev Sled

Science.gov (United States)

Haney, Joseph W.; Lenzo, J.

1996-01-01

Magnetic levitation has been explored for application from motors to transportation. All of these applications have been at velocities where the physics of the air or operating fluids are fairly well known. Application of Maglev to hypersonic velocities (Mach greater than 5) presents many opportunities, but also issues that require understanding and resolution. Use of Maglev to upgrade the High Speed Test Track at Holloman Air Force Base in Alamogordo New Mexico is an actual hypersonic application that provides the opportunity to improve test capabilities. However, there are several design issues that require investigation. This paper presents an overview of the application of Maglev to the test track and the issues associated with developing a hypersonic Maglev sled. The focus of this paper is to address the issues with the Maglev sled design, rather than the issues with the development of superconducting magnets of the sled system.

Electromagnetic Formation Flight (EMFF) for Sparse Aperture Arrays

Science.gov (United States)

Kwon, Daniel W.; Miller, David W.; Sedwick, Raymond J.

2004-01-01

Traditional methods of actuating spacecraft in sparse aperture arrays use propellant as a reaction mass. For formation flying systems, propellant becomes a critical consumable which can be quickly exhausted while maintaining relative orientation. Additional problems posed by propellant include optical contamination, plume impingement, thermal emission, and vibration excitation. For these missions where control of relative degrees of freedom is important, we consider using a system of electromagnets, in concert with reaction wheels, to replace the consumables. Electromagnetic Formation Flight sparse apertures, powered by solar energy, are designed differently from traditional propulsion systems, which are based on V. This paper investigates the design of sparse apertures both inside and outside the Earth's gravity field.

Hypersonic Inflatable Aerodynamic Decelerator (HIAD)

Data.gov (United States)

National Aeronautics and Space Administration — Develop an entry and descent technology to enhance and enable robotic and scientific missions to destinations with atmospheres.The Hypersonic Inflatable Aerodynamic...

Experimental Study Of SHEFEX II Hypersonic Aerodynamics And Canard Efficiency In H2K

Science.gov (United States)

Neeb, D.; Gulhan, A.

2011-05-01

One main objective of the DLR SHEFEX programme is to prove that sharp edged vehicles are capable of performing a re-entry into earth atmosphere by using a simple thermal protection system consisting of flat ceramic tiles. In comparison to blunt nose configurations like the Space shuttle, which are normally used for re-entry configurations, the SHEFEX TPS design is able to significantly reduce the costs and complexity of TPS structures and simultaneously increase the aerodynamic performance of the flight vehicle [1], [2]. To study its characteristics and perform several defined in-flight experiments during re-entry, the vehicle’s attitude will be controlled actively by canards [3]. In the framework of the SHEFEX II project an experimental investigation has been conducted in the hypersonic wind tunnel H2K to characterize the aerodynamic performance of the vehicle in hypersonic flow regime. The model has a modular design to enable the study of a variety of different influencing parameters. Its 4 circumferential canards have been made independently adjustable to account for the simulation of different manoeuvre conditions. To study the control behaviour of the vehicle and validate CFD data, a variation of canard deflections, angle of attack and angle of sideslip have been applied. Tests have been carried out at Mach 7 and 8.7 with a Reynolds number sensitivity study at the lower Mach number. The model was equipped with a six component internal balance to realize accurate coefficient measurements. The flow topology has been analyzed using Schlieren images. Beside general aerodynamic performance and canard efficiencies, flow phenomena like shock impingement on the canards could be determined by Schlieren images as well as by the derived coefficients.

The X-43A Hyper-X Mach 7 Flight 2 Guidance, Navigation, and Control Overview and Flight Test Results

Science.gov (United States)

Bahm, Catherine; Baumann, Ethan; Martin, John; Bose, David; Beck, Roger E.; Strovers, Brian

2005-01-01

The objective of the Hyper-X program was to flight demonstrate an airframe-integrated hypersonic vehicle. On March 27, 2004, the Hyper-X program team successfully conducted flight 2 and achieved all of the research objectives. The Hyper-X research vehicle successfully separated from the Hyper-X launch vehicle and achieved the desired engine test conditions before the experiment began. The research vehicle rejected the disturbances caused by the cowl door opening and the fuel turning on and off and maintained the engine test conditions throughout the experiment. After the engine test was complete, the vehicle recovered and descended along a trajectory while performing research maneuvers. The last data acquired showed that the vehicle maintained control to the water. This report will provide an overview of the research vehicle guidance and control systems and the performance of the vehicle during the separation event and engine test. The research maneuvers were performed to collect data for aerodynamics and flight controls research. This report also will provide an overview of the flight controls related research and results.

Tip-modified Propellers

DEFF Research Database (Denmark)

Andersen, Poul

1999-01-01

The paper deals with tip-modified propellers and the methods which, over a period of two decades, have been applied to develop such propellers. The development is driven by the urge to increase the efficiency of propellers and can be seen as analogous to fitting end plates and winglets to aircraft...... propeller, have efficiency increases of a reasonable magnitude in both open-water and behind-ship conditions....

Progress in modeling hypersonic turbulent boundary layers

Science.gov (United States)

Zeman, Otto

1993-01-01

A good knowledge of the turbulence structure, wall heat transfer, and friction in turbulent boundary layers (TBL) at high speeds is required for the design of hypersonic air breathing airplanes and reentry space vehicles. This work reports on recent progress in the modeling of high speed TBL flows. The specific research goal described here is the development of a second order closure model for zero pressure gradient TBL's for the range of Mach numbers up to hypersonic speeds with arbitrary wall cooling requirements.

Design and Execution of the Hypersonic Inflatable Aerodynamic Decelerator Large-Article Wind Tunnel Experiment

Science.gov (United States)

Cassell, Alan M.

2013-01-01

The testing of 3- and 6-meter diameter Hypersonic Inflatable Aerodynamic Decelerator (HIAD) test articles was completed in the National Full-Scale Aerodynamics Complex 40 ft x 80 ft Wind Tunnel test section. Both models were stacked tori, constructed as 60 degree half-angle sphere cones. The 3-meter HIAD was tested in two configurations. The first 3-meter configuration utilized an instrumented flexible aerodynamic skin covering the inflatable aeroshell surface, while the second configuration employed a flight-like flexible thermal protection system. The 6-meter HIAD was tested in two structural configurations (with and without an aft-mounted stiffening torus near the shoulder), both utilizing an instrumented aerodynamic skin.

Flight Performance of the Inflatable Reentry Vehicle Experiment 3

Science.gov (United States)

Dillman, Robert; DiNonno, John; Bodkin, Richard; Gsell, Valerie; Miller, Nathanael; Olds, Aaron; Bruce, Walter

2013-01-01

The Inflatable Reentry Vehicle Experiment 3 (IRVE-3) launched July 23, 2012, from NASA Wallops Flight Facility (WFF) on a Black Brant XI suborbital sounding rocket and successfully performed its mission, demonstrating the survivability of a hypersonic inflatable aerodynamic decelerator (HIAD) in the reentry heating environment and also illustrating the effect of an offset center of gravity on the HIAD's lift-to-drag ratio. IRVE-3 was a follow-on to 2009's IRVE-II mission, which demonstrated exo-atmospheric inflation, reentry survivability - without significant heating - and the aerodynamic stability of a HIAD down to subsonic flight conditions. NASA Langley Research Center is leading the development of HIAD technology for use on future interplanetary and Earth reentry missions.

Hypersonic Inflatable Aerodynamic Decelerator Ground Test Development

Science.gov (United States)

Del Corso, Jospeh A.; Hughes, Stephen; Cheatwood, Neil; Johnson, Keith; Calomino, Anthony

2015-01-01

Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology readiness levels have been incrementally matured by NASA over the last thirteen years, with most recent support from NASA's Space Technology Mission Directorate (STMD) Game Changing Development Program (GCDP). Recently STMD GCDP has authorized funding and support through fiscal year 2015 (FY15) for continued HIAD ground developments which support a Mars Entry, Descent, and Landing (EDL) study. The Mars study will assess the viability of various EDL architectures to enable a Mars human architecture pathfinder mission planned for mid-2020. At its conclusion in November 2014, NASA's first HIAD ground development effort had demonstrated success with fabricating a 50 W/cm2 modular thermal protection system, a 400 C capable inflatable structure, a 10-meter scale aeroshell manufacturing capability, together with calibrated thermal and structural models. Despite the unquestionable success of the first HIAD ground development effort, it was recognized that additional investment was needed in order to realize the full potential of the HIAD technology capability to enable future flight opportunities. The second HIAD ground development effort will focus on extending performance capability in key technology areas that include thermal protection system, lifting-body structures, inflation systems, flight control, stage transitions, and 15-meter aeroshell scalability. This paper presents an overview of the accomplishments under the baseline HIAD development effort and current plans for a follow-on development effort focused on extending those critical technologies needed to enable a Mars Pathfinder mission.

Mars Science Laboratory Heatshield Flight Data Analysis

Science.gov (United States)

Mahzari, Milad; White, Todd

2017-01-01

NASA Mars Science Laboratory (MSL), which landed the Curiosity rover on the surface of Mars on August 5th, 2012, was the largest and heaviest Mars entry vehicle representing a significant advancement in planetary entry, descent and landing capability. Hypersonic flight performance data was collected using MSLs on-board sensors called Mars Entry, Descent and Landing Instrumentation (MEDLI). This talk will give an overview of MSL entry and a description of MEDLI sensors. Observations from flight data will be examined followed by a discussion of analysis efforts to reconstruct surface heating from heatshields in-depth temperature measurements. Finally, a brief overview of MEDLI2 instrumentation, which will fly on NASAs Mars2020 mission, will be presented with a discussion on how lessons learned from MEDLI data affected the design of MEDLI2 instrumentation.

High quality ceramic coatings sprayed by high efficiency hypersonic plasma spraying gun

International Nuclear Information System (INIS)

Zhu Sheng; Xu Binshi; Yao JiuKun

2005-01-01

This paper introduced the structure of the high efficiency hypersonic plasma spraying gun and the effects of hypersonic plasma jet on the sprayed particles. The optimised spraying process parameters for several ceramic powders such as Al 2 O 3 , Cr 2 O 3 , ZrO 2 , Cr 3 C 2 and Co-WC were listed. The properties and microstructure of the sprayed ceramic coatings were investigated. Nano Al 2 O 3 -TiO 2 ceramic coating sprayed by using the high efficiency hypersonic plasma spraying was also studied. Compared with the conventional air plasma spraying, high efficiency hypersonic plasma spraying improves greatly the ceramic coatings quality but at low cost. (orig.)

Prediction and Validation of Mars Pathfinder Hypersonic Aerodynamic Data Base

Science.gov (United States)

Gnoffo, Peter A.; Braun, Robert D.; Weilmuenster, K. James; Mitcheltree, Robert A.; Engelund, Walter C.; Powell, Richard W.

1998-01-01

Postflight analysis of the Mars Pathfinder hypersonic, continuum aerodynamic data base is presented. Measured data include accelerations along the body axis and axis normal directions. Comparisons of preflight simulation and measurements show good agreement. The prediction of two static instabilities associated with movement of the sonic line from the shoulder to the nose and back was confirmed by measured normal accelerations. Reconstruction of atmospheric density during entry has an uncertainty directly proportional to the uncertainty in the predicted axial coefficient. The sensitivity of the moment coefficient to freestream density, kinetic models and center-of-gravity location are examined to provide additional consistency checks of the simulation with flight data. The atmospheric density as derived from axial coefficient and measured axial accelerations falls within the range required for sonic line shift and static stability transition as independently determined from normal accelerations.

Pitot pressure analyses in CO2 condensing rarefied hypersonic flows

Science.gov (United States)

Ozawa, T.; Suzuki, T.; Fujita, K.

2016-11-01

In order to improve the accuracy of rarefied aerodynamic prediction, a hypersonic rarefied wind tunnel (HRWT) was developed at Japan Aerospace Exploration Agency. While this wind tunnel has been limited to inert gases, such as nitrogen or argon, we recently extended the capability of HRWT to CO2 hypersonic flows for several Mars missions. Compared to our previous N2 cases, the condensation effect may not be negligible for CO2 rarefied aerodynamic measurements. Thus, in this work, we have utilized both experimental and numerical approaches to investigate the condensation and rarefaction effects in CO2 hypersonic nozzle flows.

Hypersonic drift-tearing magnetic islands in tokamak plasmas

International Nuclear Information System (INIS)

Fitzpatrick, R.; Waelbroeck, F. L.

2007-01-01

A two-fluid theory of long wavelength, hypersonic, drift-tearing magnetic islands in low-collisionality, low-β plasmas possessing relatively weak magnetic shear is developed. The model assumes both slab geometry and cold ions, and neglects electron temperature and equilibrium current gradient effects. The problem is solved in three asymptotically matched regions. The 'inner region' contains the island. However, the island emits electrostatic drift-acoustic waves that propagate into the surrounding 'intermediate region', where they are absorbed by the plasma. Since the waves carry momentum, the inner region exerts a net force on the intermediate region, and vice versa, giving rise to strong velocity shear in the region immediately surrounding the island. The intermediate region is matched to the surrounding 'outer region', in which ideal magnetohydrodynamic holds. Isolated hypersonic islands propagate with a velocity that lies between those of the unperturbed local ion and electron fluids, but is much closer to the latter. The ion polarization current is stabilizing, and increases with increasing island width. Finally, the hypersonic branch of isolated island solutions ceases to exist above a certain critical island width. Hypersonic islands whose widths exceed the critical width are hypothesized to bifurcate to the so-called 'sonic' solution branch

Assessment of CFD Hypersonic Turbulent Heating Rates for Space Shuttle Orbiter

Science.gov (United States)

Wood, William A.; Oliver, A. Brandon

2011-01-01

Turbulent CFD codes are assessed for the prediction of convective heat transfer rates at turbulent, hypersonic conditions. Algebraic turbulence models are used within the DPLR and LAURA CFD codes. The benchmark heat transfer rates are derived from thermocouple measurements of the Space Shuttle orbiter Discovery windward tiles during the STS-119 and STS-128 entries. The thermocouples were located underneath the reaction-cured glass coating on the thermal protection tiles. Boundary layer transition flight experiments conducted during both of those entries promoted turbulent flow at unusually high Mach numbers, with the present analysis considering Mach 10{15. Similar prior comparisons of CFD predictions directly to the flight temperature measurements were unsatisfactory, showing diverging trends between prediction and measurement for Mach numbers greater than 11. In the prior work, surface temperatures and convective heat transfer rates had been assumed to be in radiative equilibrium. The present work employs a one-dimensional time-accurate conduction analysis to relate measured temperatures to surface heat transfer rates, removing heat soak lag from the flight data, in order to better assess the predictive accuracy of the numerical models. The turbulent CFD shows good agreement for turbulent fuselage flow up to Mach 13. But on the wing in the wake of the boundary layer trip, the inclusion of tile conduction effects does not explain the prior observed discrepancy in trends between simulation and experiment; the flight heat transfer measurements are roughly constant over Mach 11-15, versus an increasing trend with Mach number from the CFD.

Control Relevant Modeling and Design of Scramjet-Powered Hypersonic Vehicles

Science.gov (United States)

Dickeson, Jeffrey James

This report provides an overview of scramjet-powered hypersonic vehicle modeling and control challenges. Such vehicles are characterized by unstable non-minimum phase dynamics with significant coupling and low thrust margins. Recent trends in hypersonic vehicle research are summarized. To illustrate control relevant design issues and tradeoffs, a generic nonlinear 3DOF longitudinal dynamics model capturing aero-elastic-propulsive interactions for wedge-shaped vehicle is used. Limitations of the model are discussed and numerous modifications have been made to address control relevant needs. Two different baseline configurations are examined over a two-stage to orbit ascent trajectory. The report highlights how vehicle level-flight static (trim) and dynamic properties change over the trajectory. Thermal choking constraints are imposed on control system design as a direct consequence of having a finite FER margin. The implication of this state-dependent nonlinear FER margin constraint, the right half plane (RHP) zero, and lightly damped flexible modes, on control system bandwidth (BW) and FPA tracking has been discussed. A control methodology has been proposed that addresses the above dynamics while providing some robustness to modeling uncertainty. Vehicle closure (the ability to fly a trajectory segment subject to constraints) is provided through a proposed vehicle design methodology. The design method attempts to use open loop metrics whenever possible to design the vehicle. The design method is applied to a vehicle/control law closed loop nonlinear simulation for validation. The 3DOF longitudinal modeling results are validated against a newly released NASA 6DOF code.

CFD for hypersonic propulsion

Science.gov (United States)

Povinelli, Louis A.

1991-01-01

An overview is given of research activity on the application of computational fluid dynamics (CDF) for hypersonic propulsion systems. After the initial consideration of the highly integrated nature of air-breathing hypersonic engines and airframe, attention is directed toward computations carried out for the components of the engine. A generic inlet configuration is considered in order to demonstrate the highly three dimensional viscous flow behavior occurring within rectangular inlets. Reacting flow computations for simple jet injection as well as for more complex combustion chambers are then discussed in order to show the capability of viscous finite rate chemical reaction computer simulations. Finally, the nozzle flow fields are demonstrated, showing the existence of complex shear layers and shock structure in the exhaust plume. The general issues associated with code validation as well as the specific issue associated with the use of CFD for design are discussed. A prognosis for the success of CFD in the design of future propulsion systems is offered.

Post Flight Analysis Of SHEFEX I: Shock Tunnel Testing And Related CFD Analysis

Science.gov (United States)

Schramm, Jan Martinez; Barth, Tarik; Wagner, Alexander; Hannemann, Klaus

2011-05-01

The SHarp Edge Flight EXperiment (SHEFEX) program of the German Aerospace Center (DLR) is primarily focused on the investigation of the potential to utilise improved shapes for space vehicles by considering sharp edges and facetted surfaces. One goal is to set up a sky based test facility to gain knowledge of the physics of hypersonic flow, complemented by numerical analysis and ground based testing. Further, the series of SHEFEX flight experiments is an excellent test bed for new technological concepts and flight instrumentation, and it is a source of motivation for young scientist and engineers providing an excellent school for future space-program engineers and managers. After the successful first SHEFEX flight in October 2005, a second flight is scheduled for September 2011 and additional flights are planned for 2015 ff. With the SHEFEX-I flight and the subsequent numerical and experimental post flight analysis, DLR could for the first time close the loop between the three major disciplines of aerothermodynamic research namely CFD, ground based testing and flight.

Hypersonic Control Modeling and Simulation Tool for Lifting Towed Ballutes, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Global Aerospace Corporation proposes to develop a hypersonic control modeling and simulation tool for hypersonic aeroassist vehicles. Our control and simulation...

Perseus A High Altitude Remotely Piloted Aircraft being Towed in Flight

Science.gov (United States)

1994-01-01

Perseus A, a remotely piloted, high-altitude research vehicle designed by Aurora Flight Sciences Corp., takes off from Rogers Dry Lake at the Dryden Flight Research Center, Edwards, California. The Perseus was towed into the air by a ground vehicle. At about 700 ft. the aircraft was released and the engine turned the propeller to take the plane to its desired altitude. Perseus B is a remotely piloted aircraft developed as a design-performance testbed under NASA's Environmental Research Aircraft and Sensor Technology (ERAST) project. Perseus is one of several flight vehicles involved in the ERAST project. A piston engine, propeller-powered aircraft, Perseus was designed and built by Aurora Flight Sciences Corporation, Manassas, Virginia. The objectives of Perseus B's ERAST flight tests have been to reach and maintain horizontal flight above altitudes of 60,000 feet and demonstrate the capability to fly missions lasting from 8 to 24 hours, depending on payload and altitude requirements. The Perseus B aircraft established an unofficial altitude record for a single-engine, propeller-driven, remotely piloted aircraft on June 27, 1998. It reached an altitude of 60,280 feet. In 1999, several modifications were made to the Perseus aircraft including engine, avionics, and flight-control-system improvements. These improvements were evaluated in a series of operational readiness and test missions at the Dryden Flight Research Center, Edwards, California. Perseus is a high-wing monoplane with a conventional tail design. Its narrow, straight, high-aspect-ratio wing is mounted atop the fuselage. The aircraft is pusher-designed with the propeller mounted in the rear. This design allows for interchangeable scientific-instrument payloads to be placed in the forward fuselage. The design also allows for unobstructed airflow to the sensors and other devices mounted in the payload compartment. The Perseus B that underwent test and development in 1999 was the third generation of the

Lateral control strategy for a hypersonic cruise missile

Directory of Open Access Journals (Sweden)

Yonghua Fan

2017-04-01

Full Text Available Hypersonic cruise missile always adopts the configuration of waverider body with the restraint of scramjet. As a result, the lateral motion exhibits serious coupling, and the controller design of the lateral lateral system cannot be conducted separately for yaw channel and roll channel. A multiple input and multiple output optimal control method with integrators is presented to design the lateral combined control system for hypersonic cruise missile. A hypersonic cruise missile lateral model is linearized as a multiple input and multiple output plant, which is coupled by kinematics and fin deflection between yaw and roll. In lateral combined controller, the integrators are augmented, respectively, into the loop of roll angle and lateral overload to ensure that the commands are tracked with zero steady-state error. Through simulation, the proposed controller demonstrates good performance in tracking the command of roll angle and lateral overload.

Effect of static shape deformation on aerodynamics and aerothermodynamics of hypersonic inflatable aerodynamic decelerator

Science.gov (United States)

Guo, Jinghui; Lin, Guiping; Bu, Xueqin; Fu, Shiming; Chao, Yanmeng

2017-07-01

The inflatable aerodynamic decelerator (IAD), which allows heavier and larger payloads and offers flexibility in landing site selection at higher altitudes, possesses potential superiority in next generation space transport system. However, due to the flexibilities of material and structure assembly, IAD inevitably experiences surface deformation during atmospheric entry, which in turn alters the flowfield around the vehicle and leads to the variations of aerodynamics and aerothermodynamics. In the current study, the effect of the static shape deformation on the hypersonic aerodynamics and aerothermodynamics of a stacked tori Hypersonic Inflatable Aerodynamic Decelerator (HIAD) is demonstrated and analyzed in detail by solving compressible Navier-Stokes equations with Menter's shear stress transport (SST) turbulence model. The deformed shape is obtained by structural modeling in the presence of maximum aerodynamic pressure during entry. The numerical results show that the undulating shape deformation makes significant difference to flow structure. In particular, the more curved outboard forebody surface results in local flow separations and reattachments in valleys, which consequently yields remarkable fluctuations of surface conditions with pressure rising in valleys yet dropping on crests while shear stress and heat flux falling in valleys yet rising on crests. Accordingly, compared with the initial (undeformed) shape, the corresponding differences of surface conditions get more striking outboard, with maximum augmentations of 379 pa, 2224 pa, and 19.0 W/cm2, i.e., 9.8%, 305.9%, and 101.6% for the pressure, shear stress and heat flux respectively. Moreover, it is found that, with the increase of angle of attack, the aerodynamic characters and surface heating vary and the aeroheating disparities are evident between the deformed and initial shape. For the deformable HIAD model investigated in this study, the more intense surface conditions and changed flight

Liquid Hydrogen Propellant Tank Sub-Surface Pressurization with Gaseous Helium

Science.gov (United States)

Stephens, J. R.; Cartagena, W.

2015-01-01

A series of tests were conducted to evaluate the performance of a propellant tank pressurization system with the pressurant diffuser intentionally submerged beneath the surface of the liquid. Propellant tanks and pressurization systems are typically designed with the diffuser positioned to apply pressurant gas directly into the tank ullage space when the liquid propellant is settled. Space vehicles, and potentially propellant depots, may need to conduct tank pressurization operations in micro-gravity environments where the exact location of the liquid relative to the diffuser is not well understood. If the diffuser is positioned to supply pressurant gas directly to the tank ullage space when the propellant is settled, then it may become partially or completely submerged when the liquid becomes unsettled in a microgravity environment. In such case, the pressurization system performance will be adversely affected requiring additional pressurant mass and longer pressurization times. This series of tests compares and evaluates pressurization system performance using the conventional method of supplying pressurant gas directly to the propellant tank ullage, and then supplying pressurant gas beneath the liquid surface. The pressurization tests were conducted on the Engineering Development Unit (EDU) located at Test Stand 300 at NASA Marshall Space Flight Center (MSFC). EDU is a ground based Cryogenic Fluid Management (CFM) test article supported by Glenn Research Center (GRC) and MSFC. A 150 ft3 propellant tank was filled with liquid hydrogen (LH2). The pressurization system used regulated ambient helium (GHe) as a pressurant, a variable position valve to maintain flow rate, and two identical independent pressurant diffusers. The ullage diffuser was located in the forward end of the tank and was completely exposed to the tank ullage. The submerged diffuser was located in the aft end of the tank and was completely submerged when the tank liquid level was 10% or greater

KAPPEL Propeller. Development of a Marine Propeller with Non-planar Lifting Surfaces

DEFF Research Database (Denmark)

Kappel, J.; Andersen, Poul

2002-01-01

The principle of non-planar lifting surfaces is applied to the design of modern aircraft wings to obtain better lift to drag ratios. Whereas a pronounced fin or "winglet" at the wingtip has been developed for aircraft, the application of the non-planar principle to marine propellers, dealt...... with in this paper, has led to the KAPPEL propeller with blades curved towards the suction side integrating the fin or winglet into the propeller blade. The combined theoretical, experimental and practical approach to develop and design marine propellers with non-planar lifting surfaces has resulted in propellers...

Nonlinear Adaptive Rotational Speed Control Design and Experiment of the Propeller of an Electric Micro Air Vehicle

Directory of Open Access Journals (Sweden)

Shouzhao Sheng

2016-01-01

Full Text Available Micro Air Vehicles (MAVs driven by electric propellers are of interest for military and civilian applications. The rotational speed control of such electric propellers is an important factor for improving the flight performance of the vehicles, such as their positioning accuracy and stability. Therefore, this paper presents a nonlinear adaptive control scheme for the electric propulsion system of a certain MAV, which can not only speed up the convergence rates of adjustable parameters, but can also ensure the overall stability of the adjustable parameters. The significant improvement of the dynamic tracking accuracy of the rotational speed can be easily achieved through the combination of the proposed control algorithm and linear control methods. The experimental test results have also demonstrated the positive effect of the nonlinear adaptive control scheme on the flight performance of the MAV.

Hypersonic expansion of the Fokker--Planck equation

International Nuclear Information System (INIS)

Fernandez-Feria, R.

1989-01-01

A systematic study of the hypersonic limit of a heavy species diluted in a much lighter gas is made via the Fokker--Planck equation governing its velocity distribution function. In particular, two different hypersonic expansions of the Fokker--Planck equation are considered, differing from each other in the momentum equation of the heavy gas used as the basis of the expansion: in the first of them, the pressure tensor is neglected in that equation while, in the second expansion, the pressure tensor term is retained. The expansions are valid when the light gas Mach number is O(1) or larger and the difference between the mean velocities of light and heavy components is small compared to the light gas thermal speed. They can be applied away from regions where the spatial gradient of the distribution function is very large, but it is not restricted with respect to the temporal derivative of the distribution function. The hydrodynamic equations corresponding to the lowest order of both expansions constitute two different hypersonic closures of the moment equations. For the subsequent orders in the expansions, closed sets of moment equations (hydrodynamic equations) are given. Special emphasis is made on the order of magnitude of the errors of the lowest-order hydrodynamic quantities. It is shown that if the heat flux vanishes initially, these errors are smaller than one might have expected from the ordinary scaling of the hypersonic closure. Also it is found that the normal solution of both expansions is a Gaussian distribution at the lowest order

Investigation of the Aerodynamic Performance of a DG808s UAS in Propeller Slipstream Using Computational Fluid Dynamics

Science.gov (United States)

Chandra, Yatish

Unmanned Aerial Systems (UASs) are relatively affordable and immediately available compared to commercial aircraft. Hence, their aerodynamics and design accuracies are often based on extrapolating from design standards and procedures widely used in the aerospace industry for commercial aircraft with most often, acceptable results. Engineering level software such as Advanced Aircraft Analysis (AAA) use general aviation aircraft data and later extrapolate them onto UASs for aerodynamic and flight dynamics modeling but are limited by their platform repository and relatively high Reynolds number evaluations. UASs however, are aircraft which fly at comparatively low speeds and low Reynolds number with close proximities between the components wherein such standards may not hold good. This thesis focuses on evaluating the accuracy and impact of such industry standards on the aerodynamics and flight dynamics of UASs. A DG808s UAS is chosen for the study which was previously modeled using the AAA software at The University of Kansas by the Flight Systems Team. Using the STAR-CCM+ code, performance data were compared and assessed with AAA. Aerodynamic simulations were carried out for two different configurations viz., aircraft with and without propeller slipstream effects. Data obtained for the non-powered simulations were found to be in good agreement with the AAA model. For the powered flight however, discrepancies between the AAA model and CFD data were observed with large values for the vertical tail side-force coefficient. A comparison with the system identification data from the flight tests was made to confirm and validate this vertical tail behavior with the help of rudder deflection inputs. A relationship between the propeller RPM and the aerodynamic model was established by simulating two different propeller speeds. Based on the STAR-CCM+ data and the resulting comparisons with AAA, updates necessary to the UAS aerodynamic and flight dynamics models currently used

Effect of Dielectric Barrier Discharge Plasma Actuators on Non-equilibrium Hypersonic Flows

Science.gov (United States)

2014-10-28

results for MIG with the US3D code devel- oped at the University of Minnesota.61 US3D is an unstruc- tured CFD code for hypersonic flow solution used...Effect of dielectric barrier discharge plasma actuators on non-equilibrium hypersonic flows Ankush Bhatia,1 Subrata Roy,1 and Ryan Gosse2 1Applied...a cylindrical body in Mach 17 hypersonic flow is presented. This application focuses on using sinusoidal dielectric barrier discharge plasma actuators

Technology Maturation in Preparation for the Cryogenic Propellant Storage and Transfer (CPST) Technology Demonstration Mission (TDM)

Science.gov (United States)

Meyer, Michael L.; Doherty, Michael P.; Moder, Jeffrey P.

2014-01-01

In support of its goal to find an innovative path for human space exploration, NASA embarked on the Cryogenic Propellant Storage and Transfer (CPST) Project, a Technology Demonstration Mission (TDM) to test and validate key cryogenic capabilities and technologies required for future exploration elements, opening up the architecture for large in-space cryogenic propulsion stages and propellant depots. Recognizing that key Cryogenic Fluid Management (CFM) technologies anticipated for on-orbit (flight) demonstration would benefit from additional maturation to a readiness level appropriate for infusion into the design of the flight demonstration, the NASA Headquarters Space Technology Mission Directorate (STMD) authorized funding for a one-year technology maturation phase of the CPST project. The strategy, proposed by the CPST Project Manager, focused on maturation through modeling, concept studies, and ground tests of the storage and fluid transfer of CFM technology sub-elements and components that were lower than a Technology Readiness Level (TRL) of 5. A technology maturation plan (TMP) was subsequently approved which described: the CFM technologies selected for maturation, the ground testing approach to be used, quantified success criteria of the technologies, hardware and data deliverables, and a deliverable to provide an assessment of the technology readiness after completion of the test, study or modeling activity. The specific technologies selected were grouped into five major categories: thick multilayer insulation, tank applied active thermal control, cryogenic fluid transfer, propellant gauging, and analytical tool development. Based on the success of the technology maturation efforts, the CPST project was approved to proceed to flight system development.

Hull-Propeller Interaction and Its Effect on Propeller Cavitation

DEFF Research Database (Denmark)

Regener, Pelle Bo

In order to predict the required propulsion power for a ship reliably and accurately, it is not sufficient to only evaluate the resistance of the hull and the propeller performance in open water alone. Interaction effects between hull and propeller can even be a decisive factor in ship powering...... prediction and design optimization. The hull-propeller interaction coefficients of effective wake fraction, thrust deduction factor, and relative rotative efficiency are traditionally determined by model tests. Self-propulsion model tests consistently show an increase in effective wake fractions when using...... velocities. This offers an opportunity for additional insight into hull-propeller interaction and the propeller’s actual operating condition behind the ship, as the actual (effective) inflow is computed. Self-propulsion simulations at model and full scale were carried out for a bulk carrier, once...

The influence of the choice of propeller design tool on propeller performance

OpenAIRE

Skåland, Edvard Knutsen

2016-01-01

In this master thesis different propeller design and analysis methods are presented and compared in terms of the accuracy and computational efficiency of their theory. These methods include lifting line, vortex lattice lifting surface and panel methods. A propeller design program based on lifting line theory was developed by the author. This program has been used together with the propeller design programs OpenProp and AKPD to make six propeller designs. The designs are based o...

Reaching High Altitudes on Mars with an Inflatable Hypersonic Drag Balloon (Ballute)

CERN Document Server

Griebel, Hannes

2010-01-01

The concept of probing the atmosphere of planet Mars by means of a hypersonic drag balloon, a device known as a “ballute”, is a novel approach to planetary science. In this concept, the probe deploys an inflatable drag body out in space and may then enter the atmosphere either once or several times until it slowly descends towards the ground, taking continuous atmospheric and other readings across a large altitude and ground range. Hannes Griebel discusses the theory behind such a mission along with experience gained during its practical implementation, such as mission design, manufacturing, packing and deployment techniques as well as ground and flight tests. The author also studies other ballute applications, specifically emergency low Earth orbit recovery and delivering payloads to high altitude landing sites on Mars.

CFD code calibration and inlet-fairing effects on a 3D hypersonic powered-simulation model

Science.gov (United States)

Huebner, Lawrence D.; Tatum, Kenneth E.

1993-01-01

A three-dimensional (3D) computational study has been performed addressing issues related to the wind tunnel testing of a hypersonic powered-simulation model. The study consisted of three objectives. The first objective was to calibrate a state-of-the-art computational fluid dynamics (CFD) code in its ability to predict hypersonic powered-simulation flows by comparing CFD solutions with experimental surface pressure dam. Aftbody lower surface pressures were well predicted, but lower surface wing pressures were less accurately predicted. The second objective was to determine the 3D effects on the aftbody created by fairing over the inlet; this was accomplished by comparing the CFD solutions of two closed-inlet powered configurations with a flowing-inlet powered configuration. Although results at four freestream Mach numbers indicate that the exhaust plume tends to isolate the aftbody surface from most forebody flowfield differences, a smooth inlet fairing provides the least aftbody force and moment variation compared to a flowing inlet. The final objective was to predict and understand the 3D characteristics of exhaust plume development at selected points on a representative flight path. Results showed a dramatic effect of plume expansion onto the wings as the freestream Mach number and corresponding nozzle pressure ratio are increased.

A weakly coupled semiconductor superlattice as a harmonic hypersonic-electrical transducer

International Nuclear Information System (INIS)

Poyser, C L; Akimov, A V; Campion, R P; Kent, A J; Balanov, A G

2015-01-01

We study experimentally and theoretically the effects of high-frequency strain pulse trains on the charge transport in a weakly coupled semiconductor superlattice. In a frequency range of the order of 100 GHz such excitation may be considered as single harmonic hypersonic excitation. While travelling along the axis of the SL, the hypersonic acoustic wavepacket affects the electron tunnelling, and thus governs the electrical current through the device. We reveal how the change of current depends on the parameters of the hypersonic excitation and on the bias applied to the superlattice. We have found that the changes in the transport properties of the superlattices caused by the acoustic excitation can be largely explained using the current–voltage relation of the unperturbed system. Our experimental measurements show multiple peaks in the dependence of the transferred charge on the repetition rate of the strain pulses in the train. We demonstrate that these resonances can be understood in terms of the spectrum of the applied acoustic perturbation after taking into account the multiple reflections in the metal film serving as a generator of hypersonic excitation. Our findings suggest an application of the semiconductor superlattice as a hypersonic-electrical transducer, which can be used in various microwave devices. (paper)

A unified flight control methodology for a compound rotorcraft in fundamental and aerobatic maneuvering flight

Science.gov (United States)

Thorsen, Adam

regime. An energy management system was developed in order to manage performance limits (namely power required) to promote carefree maneuvering and alleviate pilot workload. This system features limits on pilot commands and has additional logic for preserving control margins and limiting maximum speed in a dive. Nonlinear dynamic inversion (NLDI) is the framework of the unified controller, which incorporates primary and redundant controls. The inner loop of the NLDI controller regulates bank angle, pitch attitude, and yaw rate, while the outer loop command structure is varied (three modes). One version uses an outer loop that commands velocities in the longitudinal and vertical axes (velocity mode), another commands longitudinal acceleration and vertical speed (acceleration mode), and the third commands longitudinal acceleration and transitions from velocity to acceleration command in the vertical axis (aerobatic mode). The flight envelope is discretized into low, cruise, and high speed flight regimes. The unified outer loop primary control effectors for the longitudinal and vertical axes (collective pitch, pitch attitude, and propeller pitch) vary depending on flight regime. A weighted pseudoinverse is used to phase either the collective or propeller pitch in/out of a redundant control role. The controllers were evaluated in Penn State's Rotorcraft Flight Simulator retaining the cyclic stick for vertical and lateral axis control along with pedal inceptors for yaw axis control. A throttle inceptor was used in place of the pilot's traditional left hand inceptor for longitudinal axis control. Ultimately, a simple rigid body model of the aircraft was sufficient enough to design a controller with favorable performance and stability characteristics. This unified flight control system promoted a low enough pilot workload so that an untrained pilot (the author) was able to pilot maneuvers of varying complexity with ease. The framework of this unified system is generalized

Flight Test of the Lateral Stability of a 0.133-Scale Model of the Convair XFY-1 Airplane with Windmilling Propellers at Mach Numbers from 0.70 to 1.12 (TED No. NACA DE 369)

Science.gov (United States)

Hollinger, James A.; Mitcham, Grady L.

1955-01-01

A flight test of a rocket-propelled model of the Convair XFY-1 airplane was conducted to determine the lateral stability and control characteristics, The 0.133-scale model had windmilling propellers for this test, which covered a Mach number range of O.70 to 1.12. The center of gravity was located at 13.9 percent of the mean aerodynamic chord. The methods of analysis included both a solution by vector diagrams and simple one- and two-degree-of-freedom methods. The model was both statically and dynamically stable throughout the speed range of the testa The roll damping was good, and the slope of the side-force curve varied little with speed. The rudder was effective throughout the test speed range, although it was reduced to about 43 percent of its subsonic value at supersonic speeds.

Hypersonic evanescent waves generated with a planar spiral coil.

Science.gov (United States)

Stevenson, A C; Araya-Kleinsteuber, B; Sethi, R S; Mehta, H M; Lowe, C R

2003-09-01

A planar spiral coil has been used to induce hypersonic evanescent waves in a quartz substrate with the unique ability to focus the acoustic wave down onto the chemical recognition layer. These special sensing conditions were achieved by investigating the application of a radio frequency current to a coaxial waveguide and spiral coil, so that wideband repeating electrical resonance conditions could be established over the MHz to GHz frequency range. At a selected operating frequency of 1.09 GHz, the evanescent wave depth of a quartz crystal hypersonic resonance is reduced to 17 nm, minimising unwanted coupling to the bulk fluid. Verification of the validity of the hypersonic resonance was carried out by characterising the system electrically and acoustically: Impedance calculations of the combined coil and coaxial waveguide demonstrated an excellent fit to the measured data, although above 400 MHz a transition zone was identified where unwanted impedance is parasitic of the coil influence efficiency, so the signal-to-noise ratio is reduced from 3000 to 300. Acoustic quartz crystal resonances at intervals of precisely 13.2138 MHz spacing, from the 6.6 MHz ultrasonic range and onto the desired hypersonic range above 1 GHz, were incrementally detected. Q factor measurements demonstrated that reductions in energy lost from the resonator to the fluid interface were consistent with the anticipated shrinkage of the evanescent wave with increasing operating frequency. Amplitude and frequency reduction in contact with a glucose solution was demonstrated at 1.09 GHz. The complex physical conditions arising at the solid-liquid interface under hypersonic entrainment are discussed with respect to acceleration induced slippage, rupture, longitudinal and shear radiation and multiphase relaxation affects.

Relativistic Spacecraft Propelled by Directed Energy

Science.gov (United States)

Kulkarni, Neeraj; Lubin, Philip; Zhang, Qicheng

2018-04-01

Achieving relativistic flight to enable extrasolar exploration is one of the dreams of humanity and the long-term goal of our NASA Starlight program. We derive a relativistic solution for the motion of a spacecraft propelled by radiation pressure from a directed energy (DE) system. Depending on the system parameters, low-mass spacecraft can achieve relativistic speeds, thus enabling interstellar exploration. The diffraction of the DE system plays an important role and limits the maximum speed of the spacecraft. We consider “photon recycling” as a possible method to achieving higher speeds. We also discuss recent claims that our previous work on this topic is incorrect and show that these claims arise from an improper treatment of causality.

Innovative boron nitride-doped propellants

Directory of Open Access Journals (Sweden)

Thelma Manning

2016-04-01

Full Text Available The U.S. military has a need for more powerful propellants with balanced/stoichiometric amounts of fuel and oxidants. However, balanced and more powerful propellants lead to accelerated gun barrel erosion and markedly shortened useful barrel life. Boron nitride (BN is an interesting potential additive for propellants that could reduce gun wear effects in advanced propellants (US patent pending 2015-026P. Hexagonal boron nitride is a good lubricant that can provide wear resistance and lower flame temperatures for gun barrels. Further, boron can dope steel, which drastically improves its strength and wear resistance, and can block the formation of softer carbides. A scalable synthesis method for producing boron nitride nano-particles that can be readily dispersed into propellants has been developed. Even dispersion of the nano-particles in a double-base propellant has been demonstrated using a solvent-based processing approach. Stability of a composite propellant with the BN additive was verified. In this paper, results from propellant testing of boron nitride nano-composite propellants are presented, including closed bomb and wear and erosion testing. Detailed characterization of the erosion tester substrates before and after firing was obtained by electron microscopy, inductively coupled plasma and x-ray photoelectron spectroscopy. This promising boron nitride additive shows the ability to improve gun wear and erosion resistance without any destabilizing effects to the propellant. Potential applications could include less erosive propellants in propellant ammunition for large, medium and small diameter fire arms.

Hypersonic flow past slender bodies in dispersive hydrodynamics

International Nuclear Information System (INIS)

El, G.A.; Khodorovskii, V.V.; Tyurina, A.V.

2004-01-01

The problem of two-dimensional steady hypersonic flow past a slender body is formulated for dispersive media. It is shown that for the hypersonic flow, the original 2+0 boundary-value problem is asymptotically equivalent to the 1+1 piston problem for the fully nonlinear flow in the same physical system, which allows one to take advantage of the analytic methods developed for one-dimensional systems. This type of equivalence, well known in ideal Euler gas dynamics, has not been established for dispersive hydrodynamics so far. Two examples pertaining to collisionless plasma dynamics are considered

Multichannel High Resolution Wide Swath SAR Imaging for Hypersonic Air Vehicle with Curved Trajectory

Directory of Open Access Journals (Sweden)

Rui Zhou

2018-01-01

Full Text Available Synthetic aperture radar (SAR equipped on the hypersonic air vehicle in near space has many advantages over the conventional airborne SAR. However, its high-speed maneuvering characteristics with curved trajectory result in serious range migration, and exacerbate the contradiction between the high resolution and wide swath. To solve this problem, this paper establishes the imaging geometrical model matched with the flight trajectory of the hypersonic platform and the multichannel azimuth sampling model based on the displaced phase center antenna (DPCA technology. Furthermore, based on the multichannel signal reconstruction theory, a more efficient spectrum reconstruction model using discrete Fourier transform is proposed to obtain the azimuth uniform sampling data. Due to the high complexity of the slant range model, it is difficult to deduce the processing algorithm for SAR imaging. Thus, an approximate range model is derived based on the minimax criterion, and the optimal second-order approximate coefficients of cosine function are obtained using the two-population coevolutionary algorithm. On this basis, aiming at the problem that the traditional Omega-K algorithm cannot compensate the residual phase with the difficulty of Stolt mapping along the range frequency axis, this paper proposes an Exact Transfer Function (ETF algorithm for SAR imaging, and presents a method of range division to achieve wide swath imaging. Simulation results verify the effectiveness of the ETF imaging algorithm.

Multichannel High Resolution Wide Swath SAR Imaging for Hypersonic Air Vehicle with Curved Trajectory.

Science.gov (United States)

Zhou, Rui; Sun, Jinping; Hu, Yuxin; Qi, Yaolong

2018-01-31

Synthetic aperture radar (SAR) equipped on the hypersonic air vehicle in near space has many advantages over the conventional airborne SAR. However, its high-speed maneuvering characteristics with curved trajectory result in serious range migration, and exacerbate the contradiction between the high resolution and wide swath. To solve this problem, this paper establishes the imaging geometrical model matched with the flight trajectory of the hypersonic platform and the multichannel azimuth sampling model based on the displaced phase center antenna (DPCA) technology. Furthermore, based on the multichannel signal reconstruction theory, a more efficient spectrum reconstruction model using discrete Fourier transform is proposed to obtain the azimuth uniform sampling data. Due to the high complexity of the slant range model, it is difficult to deduce the processing algorithm for SAR imaging. Thus, an approximate range model is derived based on the minimax criterion, and the optimal second-order approximate coefficients of cosine function are obtained using the two-population coevolutionary algorithm. On this basis, aiming at the problem that the traditional Omega-K algorithm cannot compensate the residual phase with the difficulty of Stolt mapping along the range frequency axis, this paper proposes an Exact Transfer Function (ETF) algorithm for SAR imaging, and presents a method of range division to achieve wide swath imaging. Simulation results verify the effectiveness of the ETF imaging algorithm.

Computational study of collisions between O(3P) and NO(2Π) at temperatures relevant to the hypersonic flight regime

International Nuclear Information System (INIS)

Castro-Palacio, Juan Carlos; Nagy, Tibor; Meuwly, Markus; Bemish, Raymond J.

2014-01-01

Reactions involving N and O atoms dominate the energetics of the reactive air flow around spacecraft when reentering the atmosphere in the hypersonic flight regime. For this reason, the thermal rate coefficients for reactive processes involving O( 3 P) and NO( 2 Π) are relevant over a wide range of temperatures. For this purpose, a potential energy surface (PES) for the ground state of the NO 2 molecule is constructed based on high-level ab initio calculations. These ab initio energies are represented using the reproducible kernel Hilbert space method and Legendre polynomials. The global PES of NO 2 in the ground state is constructed by smoothly connecting the surfaces of the grids of various channels around the equilibrium NO 2 geometry by a distance-dependent weighting function. The rate coefficients were calculated using Monte Carlo integration. The results indicate that at high temperatures only the lowest A-symmetry PES is relevant. At the highest temperatures investigated (20 000 K), the rate coefficient for the “O1O2+N” channel becomes comparable (to within a factor of around three) to the rate coefficient of the oxygen exchange reaction. A state resolved analysis shows that the smaller the vibrational quantum number of NO in the reactants, the higher the relative translational energy required to open it and conversely with higher vibrational quantum number, less translational energy is required. This is in accordance with Polanyi's rules. However, the oxygen exchange channel (NO2+O1) is accessible at any collision energy. Finally, this work introduces an efficient computational protocol for the investigation of three-atom collisions in general

Computational study of collisions between O(3P) and NO(2Π) at temperatures relevant to the hypersonic flight regime.

Science.gov (United States)

Castro-Palacio, Juan Carlos; Nagy, Tibor; Bemish, Raymond J; Meuwly, Markus

2014-10-28

Reactions involving N and O atoms dominate the energetics of the reactive air flow around spacecraft when reentering the atmosphere in the hypersonic flight regime. For this reason, the thermal rate coefficients for reactive processes involving O((3)P) and NO((2)Π) are relevant over a wide range of temperatures. For this purpose, a potential energy surface (PES) for the ground state of the NO2 molecule is constructed based on high-level ab initio calculations. These ab initio energies are represented using the reproducible kernel Hilbert space method and Legendre polynomials. The global PES of NO2 in the ground state is constructed by smoothly connecting the surfaces of the grids of various channels around the equilibrium NO2 geometry by a distance-dependent weighting function. The rate coefficients were calculated using Monte Carlo integration. The results indicate that at high temperatures only the lowest A-symmetry PES is relevant. At the highest temperatures investigated (20,000 K), the rate coefficient for the "O1O2+N" channel becomes comparable (to within a factor of around three) to the rate coefficient of the oxygen exchange reaction. A state resolved analysis shows that the smaller the vibrational quantum number of NO in the reactants, the higher the relative translational energy required to open it and conversely with higher vibrational quantum number, less translational energy is required. This is in accordance with Polanyi's rules. However, the oxygen exchange channel (NO2+O1) is accessible at any collision energy. Finally, this work introduces an efficient computational protocol for the investigation of three-atom collisions in general.

再突入飛行体まわりの空力加熱のCFD解析と実験

OpenAIRE

Yamamoto, Yukimitsu; 山本　行光

1997-01-01

Recent comparisons of hypersonic CFD (Computational Fluid Dynamics) analysis with hypersonic wind tunnel experiments, including NAL M10 HWT (Hypersonic Wind Tunnel), ONERA S4MA HWT, Calspan's shock tunnel and Caltech T-5 high enthalpy shock tunnel, are introduced. Also, hypersonic reentry flight analysis, such as OREX (Orbital Reentry Experiment) and HYFLEX (Hypersonic Flight Experiment) are made. Through these studies, aerothermodynamic heating characteristics of reentry vehicle are investig...

Assessment of Aerothermodynamic Flight Prediction Tools through Ground and Flight Experimentation (Evaluation des outils aerothermodynamiques de prediction de vol par l’experimentation au sol et en vol)

Science.gov (United States)

2011-11-01

AVT-136 iii Table of Contents Page List of Figures vii List of Tables xii Programme Committee xiii Executive Summary and Synthèse ES-1 Chapter 1...136) Executive Summary In the upcoming decades military capabilities based on hypersonic flight systems will become increasingly important to NATO...rapportant très largement au développement des capacités hypersoniques envisagées : 1) Pointe avant et bords d’attaque ; 2) Interaction de chocs et

Windtunnel Rebuilding And Extrapolation To Flight At Transsonic Speed For ExoMars

Science.gov (United States)

Fertig, Markus; Neeb, Dominik; Gulhan, Ali

2011-05-01

The static as well as the dynamic behaviour of the EXOMARS vehicle in the transonic velocity regime has been investigated experimentally by the Supersonic and Hypersonic Technology Department of DLR in order to investigate the behaviour prior to parachute opening. Since the experimental work was performed in air, a numerical extrapolation to flight by means of CFD is necessary. At low supersonic speed this extrapolation to flight was performed by the Spacecraft Department of the Institute of Flow Technology of DLR employing the CFD code TAU. Numerical as well as experimental results for the wind tunnel test at Mach 1.2 will be compared and discussed for three different angles of attack.

Classifier utility modeling and analysis of hypersonic inlet start/unstart considering training data costs

Science.gov (United States)

Chang, Juntao; Hu, Qinghua; Yu, Daren; Bao, Wen

2011-11-01

Start/unstart detection is one of the most important issues of hypersonic inlets and is also the foundation of protection control of scramjet. The inlet start/unstart detection can be attributed to a standard pattern classification problem, and the training sample costs have to be considered for the classifier modeling as the CFD numerical simulations and wind tunnel experiments of hypersonic inlets both cost time and money. To solve this problem, the CFD simulation of inlet is studied at first step, and the simulation results could provide the training data for pattern classification of hypersonic inlet start/unstart. Then the classifier modeling technology and maximum classifier utility theories are introduced to analyze the effect of training data cost on classifier utility. In conclusion, it is useful to introduce support vector machine algorithms to acquire the classifier model of hypersonic inlet start/unstart, and the minimum total cost of hypersonic inlet start/unstart classifier can be obtained by the maximum classifier utility theories.

A computational study of inviscid hypersonic flows using energy relaxation method

International Nuclear Information System (INIS)

Nagdewe, Suryakant; Kim, H. D.; Shevare, G. R.

2008-01-01

Reasonable analysis of hypersonic flows requires a thermodynamic non-equilibrium model to properly simulate strong shock waves or high pressure and temperature states in the flow field. The energy relaxation method (ERM) has been used to model such a non-equilibrium effect which is generally expressed as a hyperbolic system of equations with a stiff relaxation source term. Relaxation time that is multiplied with source terms is responsible for nonequilibrium in the system. In the present study, a numerical analysis has been carried out with varying values of relaxation time for several hypersonic flows with AUSM (advection upstream splitting method) as a numerical scheme. Vibration modes of thermodynamic nonequilibrium effects are considered. The results obtained showed that, as the relaxation time reduces to zero, the solution marches toward equilibrium, while it shows non-equilibrium effects, as the relaxation time increases. The present computations predicted the experiment results of hypersonic flows with good accuracy. The work carried out suggests that the present energy relaxation method can be robust for analysis of hypersonic flows

Laminar and Turbulent Flow Calculations for the Hifire-5B Flight Test

Science.gov (United States)

2017-11-01

STATES AIR FORCE AFRL-RQ-WP-TP-2017-0172 LAMINAR AND TURBULENT FLOW CALCULATIONS FOR THE HIFIRE-5B FLIGHT TEST Roger L. Kimmel Hypersonic Sciences...stationary instabilities of the three-dimensional flow as the grid becomes finer. It may not be possible to obtain a strictly laminar basic state on a very...fine grid. A basic state solution was desired for the laminar flow calculations, and the oscillations observed in Fig. 3 were judged to be undesirable

High speed digital holographic interferometry for hypersonic flow visualization

Science.gov (United States)

Hegde, G. M.; Jagdeesh, G.; Reddy, K. P. J.

2013-06-01

Optical imaging techniques have played a major role in understanding the flow dynamics of varieties of fluid flows, particularly in the study of hypersonic flows. Schlieren and shadowgraph techniques have been the flow diagnostic tools for the investigation of compressible flows since more than a century. However these techniques provide only the qualitative information about the flow field. Other optical techniques such as holographic interferometry and laser induced fluorescence (LIF) have been used extensively for extracting quantitative information about the high speed flows. In this paper we present the application of digital holographic interferometry (DHI) technique integrated with short duration hypersonic shock tunnel facility having 1 ms test time, for quantitative flow visualization. Dynamics of the flow fields in hypersonic/supersonic speeds around different test models is visualized with DHI using a high-speed digital camera (0.2 million fps). These visualization results are compared with schlieren visualization and CFD simulation results. Fringe analysis is carried out to estimate the density of the flow field.

Frequencies of inaudible high-frequency sounds differentially affect brain activity: positive and negative hypersonic effects.

Directory of Open Access Journals (Sweden)

Ariko Fukushima

Full Text Available The hypersonic effect is a phenomenon in which sounds containing significant quantities of non-stationary high-frequency components (HFCs above the human audible range (max. 20 kHz activate the midbrain and diencephalon and evoke various physiological, psychological and behavioral responses. Yet important issues remain unverified, especially the relationship existing between the frequency of HFCs and the emergence of the hypersonic effect. In this study, to investigate the relationship between the hypersonic effect and HFC frequencies, we divided an HFC (above 16 kHz of recorded gamelan music into 12 band components and applied them to subjects along with an audible component (below 16 kHz to observe changes in the alpha2 frequency component (10-13 Hz of spontaneous EEGs measured from centro-parieto-occipital regions (Alpha-2 EEG, which we previously reported as an index of the hypersonic effect. Our results showed reciprocal directional changes in Alpha-2 EEGs depending on the frequency of the HFCs presented with audible low-frequency component (LFC. When an HFC above approximately 32 kHz was applied, Alpha-2 EEG increased significantly compared to when only audible sound was applied (positive hypersonic effect, while, when an HFC below approximately 32 kHz was applied, the Alpha-2 EEG decreased (negative hypersonic effect. These findings suggest that the emergence of the hypersonic effect depends on the frequencies of inaudible HFC.

Frequencies of inaudible high-frequency sounds differentially affect brain activity: positive and negative hypersonic effects.

Science.gov (United States)

Fukushima, Ariko; Yagi, Reiko; Kawai, Norie; Honda, Manabu; Nishina, Emi; Oohashi, Tsutomu

2014-01-01

The hypersonic effect is a phenomenon in which sounds containing significant quantities of non-stationary high-frequency components (HFCs) above the human audible range (max. 20 kHz) activate the midbrain and diencephalon and evoke various physiological, psychological and behavioral responses. Yet important issues remain unverified, especially the relationship existing between the frequency of HFCs and the emergence of the hypersonic effect. In this study, to investigate the relationship between the hypersonic effect and HFC frequencies, we divided an HFC (above 16 kHz) of recorded gamelan music into 12 band components and applied them to subjects along with an audible component (below 16 kHz) to observe changes in the alpha2 frequency component (10-13 Hz) of spontaneous EEGs measured from centro-parieto-occipital regions (Alpha-2 EEG), which we previously reported as an index of the hypersonic effect. Our results showed reciprocal directional changes in Alpha-2 EEGs depending on the frequency of the HFCs presented with audible low-frequency component (LFC). When an HFC above approximately 32 kHz was applied, Alpha-2 EEG increased significantly compared to when only audible sound was applied (positive hypersonic effect), while, when an HFC below approximately 32 kHz was applied, the Alpha-2 EEG decreased (negative hypersonic effect). These findings suggest that the emergence of the hypersonic effect depends on the frequencies of inaudible HFC.

Optimization of the Upper Surface of Hypersonic Vehicle Based on CFD Analysis

Science.gov (United States)

Gao, T. Y.; Cui, K.; Hu, S. C.; Wang, X. P.; Yang, G. W.

2011-09-01

For the hypersonic vehicle, the aerodynamic performance becomes more intensive. Therefore, it is a significant event to optimize the shape of the hypersonic vehicle to achieve the project demands. It is a key technology to promote the performance of the hypersonic vehicle with the method of shape optimization. Based on the existing vehicle, the optimization to the upper surface of the Simplified hypersonic vehicle was done to obtain a shape which suits the project demand. At the cruising condition, the upper surface was parameterized with the B-Spline curve method. The incremental parametric method and the reconstruction technology of the local mesh were applied here. The whole flow field was been calculated and the aerodynamic performance of the craft were obtained by the computational fluid dynamic (CFD) technology. Then the vehicle shape was optimized to achieve the maximum lift-drag ratio at attack angle 3°, 4° and 5°. The results will provide the reference for the practical design.

Hyperheat: a thermal signature model for super- and hypersonic missiles

Science.gov (United States)

van Binsbergen, S. A.; van Zelderen, B.; Veraar, R. G.; Bouquet, F.; Halswijk, W. H. C.; Schleijpen, H. M. A.

2017-10-01

In performance prediction of IR sensor systems for missile detection, apart from the sensor specifications, target signatures are essential variables. Very often, for velocities up to Mach 2-2.5, a simple model based on the aerodynamic heating of a perfect gas was used to calculate the temperatures of missile targets. This typically results in an overestimate of the target temperature with correspondingly large infrared signatures and detection ranges. Especially for even higher velocities, this approach is no longer accurate. Alternatives like CFD calculations typically require more complex sets of inputs and significantly more computing power. The MATLAB code Hyperheat was developed to calculate the time-resolved skin temperature of axisymmetric high speed missiles during flight, taking into account the behaviour of non-perfect gas and proper heat transfer to the missile surface. Allowing for variations in parameters like missile shape, altitude, atmospheric profile, angle of attack, flight duration and super- and hypersonic velocities up to Mach 30 enables more accurate calculations of the actual target temperature. The model calculates a map of the skin temperature of the missile, which is updated over the flight time of the missile. The sets of skin temperature maps are calculated within minutes, even for >100 km trajectories, and can be easily converted in thermal infrared signatures for further processing. This paper discusses the approach taken in Hyperheat. Then, the thermal signature of a set of typical missile threats is calculated using both the simple aerodynamic heating model and the Hyperheat code. The respective infrared signatures are compared, as well as the difference in the corresponding calculated detection ranges.

Toward a CFD nose-to-tail capability - Hypersonic unsteady Navier-Stokes code validation

Science.gov (United States)

Edwards, Thomas A.; Flores, Jolen

1989-01-01

Computational fluid dynamics (CFD) research for hypersonic flows presents new problems in code validation because of the added complexity of the physical models. This paper surveys code validation procedures applicable to hypersonic flow models that include real gas effects. The current status of hypersonic CFD flow analysis is assessed with the Compressible Navier-Stokes (CNS) code as a case study. The methods of code validation discussed to beyond comparison with experimental data to include comparisons with other codes and formulations, component analyses, and estimation of numerical errors. Current results indicate that predicting hypersonic flows of perfect gases and equilibrium air are well in hand. Pressure, shock location, and integrated quantities are relatively easy to predict accurately, while surface quantities such as heat transfer are more sensitive to the solution procedure. Modeling transition to turbulence needs refinement, though preliminary results are promising.

Coupled thermal, structural and vibrational analysis of a hypersonic engine for flight test

Energy Technology Data Exchange (ETDEWEB)

Sook-Ying, Ho [Defence Science and Technology Organisation, SA (Australia); Paull, A. [Queensland Univ., Dept. of Mechanical Engineering (Australia)

2006-07-15

This paper describes a relatively simple and quick method for implementing aerodynamic heating models into a finite element code for non-linear transient thermal-structural and thermal-structural-vibrational analyses of a Mach 10 generic HyShot scram-jet engine. The thermal-structural-vibrational response of the engine was studied for the descent trajectory from 60 to 26 km. Aerodynamic heating fluxes, as a function of spatial position and time for varying trajectory points, were implemented in the transient heat analysis. Additionally, the combined effect of varying dynamic pressure and thermal loads with altitude was considered. This aero-thermal-structural analysis capability was used to assess the temperature distribution, engine geometry distortion and yielding of the structural material due to aerodynamic heating during the descent trajectory, and for optimising the wall thickness, nose radius of leading edge, etc. of the engine intake. A structural vibration analysis was also performed following the aero-thermal-structural analysis to determine the changes in natural frequencies of the structural vibration modes that occur at the various temperatures associated with the descent trajectory. This analysis provides a unique and relatively simple design strategy for predicting and mitigating the thermal-structural-vibrational response of hypersonic engines. (authors)

Multi Laser Pulse Investigation of the DEAS Concept in Hypersonic Flow

International Nuclear Information System (INIS)

Minucci, M.A.S.; Toro, P.G.P.; Oliveira, A.C.; Chanes, J.B. Jr.; Ramos, A.G.; Nagamatsu, H.T.; Myrabo, L.N.

2004-01-01

The present paper presents recent experimental results on the Laser-Supported Directed Energy 'Air Spike' - DEAS in hypersonic flow achieved by the Laboratory of Aerothermodynamics and Hypersonics - LAH, Brazil. Two CO2 TEA lasers, sharing the same optical cavity, have been used in conjunction with the IEAv 0.3m Hypersonic Shock Tunnel - HST to demonstrate the Laser-Supported DEAS concept. A single and double laser pulse, generated during the tunnel useful test time, were focused through a NaCl lens upstream of a Double Apollo Disc model fitted with seven piezoelectric pressure transducers and six platinum thin film heat transfer gauges. The objective being to corroborate previous results as well as to obtain additional pressure and heat flux distributions information when two laser pulses are used

Flight performance summary for three NASA Terrier-Malemute II sounding rockets

Science.gov (United States)

Patterson, R. A.

1982-01-01

The subject of this paper is the presentation of flight data for three Terrier-Malemute II sounding rocket vehicles. The Malemute motor was modified by adding insulation and using a propellant that produced less Al2O3 agglomerate in the chamber. This modification, designated Malemute II, reduced the sensitivity of the motor to the roll rate induced motor case burnthrough experienced on some earlier Malemute flights. Two flight tests, including a single stage Malemute II and a Terrier-Malemute II, were made by Sandia to qualify this modification. The three NASA operational flights that are the subject of this paper were made using the modified Malemute II motors.

Numerical analysis of a hypersonic turbulent and laminar flow using a commercial CFD solver

OpenAIRE

Pajčin Miroslav P.; Simonović Aleksandar M.; Ivanov Toni D.; Komarov Dragan M.; Stupar Slobodan N.

2017-01-01

Computational fluid dynamics computations for two hypersonic flow cases using the commercial ANSYS FLUENT 16.2 CFD software were done. In this paper, an internal and external hypersonic flow cases were considered and analysis of the hypersonic flow using different turbulence viscosity models available in ANSYS FLUENT 16.2 as well as the laminar viscosity model were done. The obtained results were after compared and commented upon. [Project of the Serbian Ministry of Education, Science and Tec...

Hypersonic Vehicle Propulsion System Simplified Model Development

Science.gov (United States)

Stueber, Thomas J.; Raitano, Paul; Le, Dzu K.; Ouzts, Peter

2007-01-01

This document addresses the modeling task plan for the hypersonic GN&C GRC team members. The overall propulsion system modeling task plan is a multi-step process and the task plan identified in this document addresses the first steps (short term modeling goals). The procedures and tools produced from this effort will be useful for creating simplified dynamic models applicable to a hypersonic vehicle propulsion system. The document continues with the GRC short term modeling goal. Next, a general description of the desired simplified model is presented along with simulations that are available to varying degrees. The simulations may be available in electronic form (FORTRAN, CFD, MatLab,...) or in paper form in published documents. Finally, roadmaps outlining possible avenues towards realizing simplified model are presented.

Hypersonic Air Flow with Finite Rate Chemistry

National Research Council Canada - National Science Library

Boyd, Ian

1997-01-01

... describe the effects of non-equilibrium flow chemistry, shock interaction, and turbulent mixing and combustion on the performance of vehicles and air breathing engines designed to fly in the hypersonic flow...

Project Minerva: A low cost manned Mars mission based on indigenous propellant production

Science.gov (United States)

Beder, David; Bryan, Richard; Bui, Tuyen; Caviezel, Kelly; Cinnamon, Mark; Daggert, Todd; Folkers, Mike; Fornia, Mark; Hanks, Natasha; Hamilton, Steve

1992-01-01

Project Minerva is a low-cost manned Mars mission designed to deliver a crew of four to the Martian surface using only two sets of two launches from the Kennedy Space Center. Key concepts which make this mission realizable are the use of near-term technologies and in-situ propellant production, following the scenario originally proposed by R. Zubrin. The first set of launches delivers two unmanned payloads into low Earth orbit (LEO): the first payload consists of an Earth Return Vehicle (ERV), a propellant production plant, and a set of robotic vehicles; the second payload consists of the trans-Mars injection (TMI) upper stage. In LEO, the two payloads are docked and the configuration is injected into a Mars transfer orbit. The landing on Mars is performed with the aid of multiple aerobraking maneuvers. On the Martian surface, the propellant production plant uses a Sabatier/electrolysis type process to combine nine tons of hydrogen with carbon dioxide from the Martian atmosphere to produce over a hundred tons of liquid oxygen and liquid methane, which are later used as the propellants for the rover expeditions and the manned return journey of the ERV. The systems necessary for the flights to and from Mars, as well as those needed for the stay on Mars, are discussed. These systems include the transfer vehicle design, life support, guidance and communications, rovers and telepresence, power generation, and propellant manufacturing. Also included are the orbital mechanics, the scientific goals, and the estimated mission costs.

Propellers in Saturn's rings

Science.gov (United States)

Sremcevic, M.; Stewart, G. R.; Albers, N.; Esposito, L. W.

2013-12-01

Theoretical studies and simulations have demonstrated the effects caused by objects embedded in planetary rings. Even if the objects are too small to be directly observed, each creates a much larger gravitational imprint on the surrounding ring material. These strongly depend on the mass of the object and range from "S" like propeller-shaped structures for about 100m-sized icy bodies to the opening of circumferential gaps as in the case of the embedded moons Pan and Daphnis and their corresponding Encke and Keeler Gaps. Since the beginning of the Cassini mission many of these smaller objects (~data from Cassini Ultraviolet Imaging Spectrograph (UVIS) and Imaging Science Subsystem (ISS) experiments. We show evidence that B ring seems to harbor two distinct populations of propellers: "big" propellers covering tens of degrees in azimuth situated in the densest part of B ring, and "small" propellers in less dense inner B ring that are similar in size and shape to known A ring propellers. The population of "big" propellers is exemplified with a single object which is observed for 5 years of Cassini data. The object is seen as a very elongated bright stripe (40 degrees wide) in unlit Cassini images, and dark stripe in lit geometries. In total we report observing the feature in images at 18 different epochs between 2005 and 2010. In UVIS occultations we observe this feature as an optical depth depletion in 14 out of 93 occultation cuts at corrotating longitudes compatible with imaging data. Combining the available Cassini data we infer that the object is a partial gap located at r=112,921km embedded in the high optical depth region of the B ring. The gap moves at Kepler speed appropriate for its radial location. Radial offsets of the gap locations in UVIS occultations are consistent with an asymmetric propeller shape. The asymmetry of the observed shape is most likely a consequence of the strong surface mass density gradient, as the feature is located at an edge between

Effect of Propeller Angle Relative to Flow on Aerodynamic Characteristics

Science.gov (United States)

Schueller, Joseph; Hubner, Paul

2017-11-01

As the interest in small unmanned air systems (UASs) for delivery and surveillance grows, new hybrid designs are being studied to take advantage of both quadcopters and fixed-wing aircraft. The tiltrotor design is able to combine the vertical take-off, hover, and landing of a multi-rotor copter with the efficiency of forward flight of a conventional airplane. However, literature documenting aerodynamic performance of the rotor as it rotates between the forward-flight and hover positions, especially in this low Reynolds number range, is limited. This data is critical for validating computational models and developing safe transition corridors. The objective of this research was to design, build and test a rotor thrust stand capable of rotating between the forward-flight and hover configurations suitable for senior design studies at low Reynolds number research. The poster covers the design of the rotating mechanism, the range and resolution of the load cell, and the thrust, torque and efficiency results for a conventional UAS motor and propeller for various advance ratios and thrust-line orientations. NSF Grant: EEC 1659710.

NATO Advanced Study Institute on Molecular Physics and Hypersonic Flows

CERN Document Server

1996-01-01

Molecular Physics and Hypersonic Flows bridges the gap between the fluid dynamics and molecular physics communities, emphasizing the role played by elementary processes in hypersonic flows. In particular, the work is primarily dedicated to filling the gap between microscopic and macroscopic treatments of the source terms to be inserted in the fluid dynamics codes. The first part of the book describes the molecular dynamics of elementary processes both in the gas phase and in the interaction with surfaces by using quantum mechanical and phenomenological approaches. A second group of contributions describes thermodynamics and transport properties of air components, with special attention to the transport of internal energy. A series of papers is devoted to the experimental and theoretical study of the flow of partially ionized gases. Subsequent contributions treat modern computational techniques for 3-D hypersonic flow. Non-equilibrium vibrational kinetics are then described, together with the coupling of vibra...

An implicit multigrid algorithm for computing hypersonic, chemically reacting viscous flows

International Nuclear Information System (INIS)

Edwards, J.R.

1996-01-01

An implicit algorithm for computing viscous flows in chemical nonequilibrium is presented. Emphasis is placed on the numerical efficiency of the time integration scheme, both in terms of periteration workload and overall convergence rate. In this context, several techniques are introduced, including a stable, O(m 2 ) approximate factorization of the chemical source Jacobian and implementations of V-cycle and filtered multigrid acceleration methods. A five species-seventeen reaction air model is used to calculate hypersonic viscous flow over a cylinder at conditions corresponding to flight at 5 km/s, 60 km altitude and at 11.36 km/s, 76.42 km altitude. Inviscid calculations using an eleven-species reaction mechanism including ionization are presented for a case involving 11.37 km/s flow at an altitude of 84.6 km. Comparisons among various options for the implicit treatment of the chemical source terms and among different multilevel approaches for convergence acceleration are presented for all simulations

Quadcopter thrust optimization with ducted-propeller

Directory of Open Access Journals (Sweden)

Kuantama Endrowednes

2017-01-01

Full Text Available In relation to quadcopter body frame model, propeller can be categorized into propeller with ducted and without ducted. This study present differences between those two using CFD (Computational Fluid Dynamics method. Both categories utilize two blade-propeller with diameter of 406 (mm. Propeller rotation generates acceleration per time unit on the volume of air. Based on the behavior of generated air velocity, ducted propeller can be modeled into three versions. The generated thrust and performance on each model were calculated to determine the best model. The use of ducted propeller increases the total weight of quadcopter and also total thrust. The influence of this modeling were analyzed in detail with variation of angular velocity propeller from 1000 (rpm to 9000 (rpm. Besides the distance between propeller tip and ducted barrier, the size of ducted is also an important part in thrust optimization and total weight minimization of quadcopter.

On air-chemistry reduction for hypersonic external flow applications

International Nuclear Information System (INIS)

Ibrahim, Ashraf; Suman, Sawan; Girimaji, Sharath S.

2015-01-01

Highlights: • The existence of the slow manifold for the air-mixture system is shown. • The QSSA estimate of the slow manifold is fairly accurate. • For mid-temperature range the reduction mechanisms could be useful. - Abstract: In external hypersonic flows, viscous and compressibility effects generate very high temperatures leading to significant chemical reactions among air constituents. Therefore, hypersonic flow computations require coupled calculations of flow and chemistry. Accurate and efficient computations of air-chemistry kinetics are of much importance for many practical applications but calculations accounting for detailed chemical kinetics can be prohibitively expensive. In this paper, we investigate the possibility of applying chemical kinetics reduction schemes for hypersonic air-chemistry. We consider two chemical kinetics sets appropriate for three different temperature ranges: 2500 K to 4500 K; 4500 K to 9000 K; and above 9000 K. By demonstrating the existence of the so-called the slow manifold in each of the chemistry sets, we show that judicious chemical kinetics reduction leading to significant computational savings is possible without much loss in accuracy

COMPLETE SEPARATION OF THE VERTICAL AND HORIZONTAL INDEPENDENT COMPONENTS OF THE FLIGHT IN POLICOPTER UAV NAU PKF "AURORA" AND MATHEMATICAL MODEL OF THIS FLIGHT

Directory of Open Access Journals (Sweden)

Volodymyr Kharchenko

2017-07-01

Full Text Available Purpose: This article presents a mathematical model and the experimental results of automatic flights of the policopter UAV NAU PKF "Aurora" of oktacopter scheme with additional elektroimpeler engines of horizontal thrust. Methods: UAV NAU PKF "Aurora" is developed for experimental flights in manual, semi-automatic and unmanned modes. The uniqueness and scientific novelty of data of flight testes is in a complete separation and isolation of vertical and horizontal components of the flight, which enables a fundamentally new way of moving of vehicle in the aerial space. This approach gives a ability to obtain all advantages and to eliminate disadvantages of helicopter and airplane in fundamentally new aircraft by structure and by function – namely in the policopter flyer with additional independent engines of the lateral thrust. Results: Obtained a new experimental data that allowed to better understand the nature of the physical forces, providing the flight of the policopter. Discussion: Revised a physical basis of the airscrew (propeller, namely on the example of flight of the policopter proved that most of the thrust of the propeller provided by the mechanical impulse (kinetic energy Ек=mv2/2 by the impulse, that a airscrew receives at his collisions with air molecules,but not by the gradient of air pressure below and above the airscrew. Is put forward a hypothesis of gravitational nature of the flight and introduced the notion of "functional antigravity", that a force completely identical in function and opposite on the direction of the force of gravity (gravity force. Deduced a mathematical formula of "functionally antigravitational" transport, namely:G·M·m/R2 = mI·v2/2 – for the flights of the aircraft with a mass m over universal astronomical body with a mass M, and m·g = mI·v2/2 – for the flights of the aircraft with mass m over a planet Earth.

Numerical analysis of a hypersonic turbulent and laminar flow using a commercial CFD solver

Directory of Open Access Journals (Sweden)

Pajčin Miroslav P.

2017-01-01

Full Text Available Computational fluid dynamics computations for two hypersonic flow cases using the commercial ANSYS FLUENT 16.2 CFD software were done. In this paper, an internal and external hypersonic flow cases were considered and analysis of the hypersonic flow using different turbulence viscosity models available in ANSYS FLUENT 16.2 as well as the laminar viscosity model were done. The obtained results were after compared and commented upon. [Project of the Serbian Ministry of Education, Science and Technological Development, Grant no. 35035

A Numerical Study of 2-D Surface Roughness Effects on the Growth of Wave Modes in Hypersonic Boundary Layers

Science.gov (United States)

Fong, Kahei Danny

stability equation (PSE) are performed to ensure the fidelity of the data and to study the relevant flow physics. All results unanimously confirm the conclusion that the relative location of the synchronization point with respect to the roughness element determines the roughness effect on the second mode. Namely, a roughness placed upstream of the synchronization point amplifies the unstable waves while placing a roughness downstream of the synchronization point damps the second-mode waves. The parametric study also shows that a tall roughness element within the local boundary-layer thickness results in a stronger damping effect, while the effect of the roughness width is relatively insignificant compared with the other roughness parameters. On the other hand, the fact that both LST and PSE successfully predict the damping effect only by analyzing the meanflow suggests the mechanism of the damping is by the meanflow alteration due to the existence of roughness elements, rather than new mode generation. In addition to studying the unstable waves, the drag force and heating with and without roughness have been investigated by comparing the numerical simulation data with experimental correlations. It is shown that the increase in drag force generated by the Mach wave around a roughness element in a hypersonic boundary layer is insignificant compared to the reduction of drag force by suppressing turbulent flow. The study also shows that, for a cold wall flow which is the case for practical flight applications, the Stanton number decreases as roughness elements smooth out the temperature gradient in the wall-normal direction. Based on the knowledge of roughness elements damping the second mode gained from the current study, a novel passive transition control method using judiciously placed roughness elements has been developed, and patented, during the course of this research. The main idea of the control method is that, with a given geometry and flow condition, it is possible

Flow visualization of a low density hypersonic flow field

International Nuclear Information System (INIS)

Masson, B.S.; Jumper, E.J.; Walters, E.; Segalman, T.Y.; Founds, N.D.

1989-01-01

Characteristics of laser induced iodine fluorescence (LIIF) in low density hypersonic flows are being investigated for use as a diagnostic technique. At low pressures, doppler broadening dominates the iodine absorption profile producing a fluorescence signal that is primarily temperature and velocity dependent. From this dependency, a low pressure flow field has the potential to be mapped for its velocity and temperature fields. The theory for relating iodine emission to the velocity and temperature fields of a hypersonic flow is discussed in this paper. Experimental observations are made of a fluorescencing free expansion and qualitatively related to the theory. 7 refs

An investigation of the effects of the propeller slipstream of a laminar wing boundary layer

Science.gov (United States)

Howard, R. M.; Miley, S. J.; Holmes, B. J.

1985-01-01

A research program is in progress to study the effects of the propeller slipstream on natural laminar flow. Flight and wind tunnel measurements of the wing boundary layer have been made using hot-film velocity sensor probes. The results show the boundary layer, at any given point, to alternate between laminar and turbulent states. This cyclic behavior is due to periodic external flow turbulence originating from the viscous wake of the propeller blades. Analytic studies show the cyclic laminar/turbulent boundary layer to result in a significantly lower wing section drag than a fully turbulent boundary layer. The application of natural laminar flow design philosophy yields drag reduction benefits in the slipstream affected regions of the airframe, as well as the unaffected regions.

Hypersonic - Model Analysis as a Service

DEFF Research Database (Denmark)

Acretoaie, Vlad; Störrle, Harald

2014-01-01

Hypersonic is a Cloud-based tool that proposes a new approach to the deployment of model analysis facilities. It is implemented as a RESTful Web service API o_ering analysis features such as model clone detection. This approach allows the migration of resource intensive analysis algorithms from...

Regarding the perturbed operating process of DB propellant rocket motor at extreme initial grain temperatures

Directory of Open Access Journals (Sweden)

Ioan ION

2012-03-01

Full Text Available Despite many decades of study, the combustion instability of several DB propellants is still of particular concern, especially at extreme grain temperature conditions of rocket motor operating. The purpose of the first part of the paper is to give an overview of our main experimental results on combustion instabilities and pressure oscillations in DB propellant segmented grain rocket motors (SPRM-01, large L/D ratio, working at extreme initial grain temperatures. Thus, we recorded some particular pressure-time traces with significant perturbed pressure signal that was FFT analysed. An updated mathematical model incorporating transient frequency-dependent combustion response, in conjunction with pressure-dependent burning, is applied to investigate and predict the DB propellant combustion instability phenomenon. The susceptibility of the tested motor SPRM-01 with DB propellant to get a perturbed working and to go unstable with pressure was evidenced and this risk has to be evaluated. In the last part of our paper we evaluated the influence of recorded perturbed thrust on the rocket behaviour on the trajectory. The study revealed that at firing-table initial conditions, this kind of perturbed motor operating may not lead to an unstable rocket flight, but the ballistic parameters would be influenced in an unacceptable manner.

Status of Turbulence Modeling for Hypersonic Propulsion Flowpaths

Science.gov (United States)

Georgiadis, Nicholas J.; Yoder, Dennis A.; Vyas, Manan A.; Engblom, William A.

2012-01-01

This report provides an assessment of current turbulent flow calculation methods for hypersonic propulsion flowpaths, particularly the scramjet engine. Emphasis is placed on Reynolds-averaged Navier-Stokes (RANS) methods, but some discussion of newer meth- ods such as Large Eddy Simulation (LES) is also provided. The report is organized by considering technical issues throughout the scramjet-powered vehicle flowpath including laminar-to-turbulent boundary layer transition, shock wave / turbulent boundary layer interactions, scalar transport modeling (specifically the significance of turbulent Prandtl and Schmidt numbers) and compressible mixing. Unit problems are primarily used to conduct the assessment. In the combustor, results from calculations of a direct connect supersonic combustion experiment are also used to address the effects of turbulence model selection and in particular settings for the turbulent Prandtl and Schmidt numbers. It is concluded that RANS turbulence modeling shortfalls are still a major limitation to the accuracy of hypersonic propulsion simulations, whether considering individual components or an overall system. Newer methods such as LES-based techniques may be promising, but are not yet at a maturity to be used routinely by the hypersonic propulsion community. The need for fundamental experiments to provide data for turbulence model development and validation is discussed.

Perseus B Taxi Tests in Preparation for a New Series of Flight Tests

Science.gov (United States)

1998-01-01

The Perseus B remotely piloted aircraft taxis on the runway at Edwards Air Force Base, California, before a series of development flights at NASA's Dryden flight Research Center. The Perseus B is the latest of three versions of the Perseus design developed by Aurora Flight Sciences under NASA's Environmental Research Aircraft and Sensor Technology (ERAST) program. Perseus B is a remotely piloted aircraft developed as a design-performance testbed under NASA's Environmental Research Aircraft and Sensor Technology (ERAST) project. Perseus is one of several flight vehicles involved in the ERAST project. A piston engine, propeller-powered aircraft, Perseus was designed and built by Aurora Flight Sciences Corporation, Manassas, Virginia. The objectives of Perseus B's ERAST flight tests have been to reach and maintain horizontal flight above altitudes of 60,000 feet and demonstrate the capability to fly missions lasting from 8 to 24 hours, depending on payload and altitude requirements. The Perseus B aircraft established an unofficial altitude record for a single-engine, propeller-driven, remotely piloted aircraft on June 27, 1998. It reached an altitude of 60,280 feet. In 1999, several modifications were made to the Perseus aircraft including engine, avionics, and flight-control-system improvements. These improvements were evaluated in a series of operational readiness and test missions at the Dryden Flight Research Center, Edwards, California. Perseus is a high-wing monoplane with a conventional tail design. Its narrow, straight, high-aspect-ratio wing is mounted atop the fuselage. The aircraft is pusher-designed with the propeller mounted in the rear. This design allows for interchangeable scientific-instrument payloads to be placed in the forward fuselage. The design also allows for unobstructed airflow to the sensors and other devices mounted in the payload compartment. The Perseus B that underwent test and development in 1999 was the third generation of the Perseus

N-S/DSMC hybrid simulation of hypersonic flow over blunt body including wakes

Science.gov (United States)

Li, Zhonghua; Li, Zhihui; Li, Haiyan; Yang, Yanguang; Jiang, Xinyu

2014-12-01

A hybrid N-S/DSMC method is presented and applied to solve the three-dimensional hypersonic transitional flows by employing the MPC (modular Particle-Continuum) technique based on the N-S and the DSMC method. A sub-relax technique is adopted to deal with information transfer between the N-S and the DSMC. The hypersonic flows over a 70-deg spherically blunted cone under different Kn numbers are simulated using the CFD, DSMC and hybrid N-S/DSMC method. The present computations are found in good agreement with DSMC and experimental results. The present method provides an efficient way to predict the hypersonic aerodynamics in near-continuum transitional flow regime.

Assessment of CFD capability for prediction of hypersonic shock interactions

Science.gov (United States)

Knight, Doyle; Longo, José; Drikakis, Dimitris; Gaitonde, Datta; Lani, Andrea; Nompelis, Ioannis; Reimann, Bodo; Walpot, Louis

2012-01-01

The aerothermodynamic loadings associated with shock wave boundary layer interactions (shock interactions) must be carefully considered in the design of hypersonic air vehicles. The capability of Computational Fluid Dynamics (CFD) software to accurately predict hypersonic shock wave laminar boundary layer interactions is examined. A series of independent computations performed by researchers in the US and Europe are presented for two generic configurations (double cone and cylinder) and compared with experimental data. The results illustrate the current capabilities and limitations of modern CFD methods for these flows.

Experimental Study of Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Aeroshell with Axisymmetric Surface Deflection Patterns

Science.gov (United States)

Hollis, Brian R.; Hollingsworth, Kevin E.

2017-01-01

A wind tunnel test program was conducted to obtain aeroheating environment data on Hypersonic Inflatable Aerodynamic Decelerator aeroshells with flexible thermal protection systems. Data were obtained on a set of rigid wind tunnel models with surface deflection patterns of various heights that simulated a range of potential in-flight aeroshell deformations. Wind tunnel testing was conducted at Mach 6 at unit Reynolds numbers from 2.1 × 10(exp 6)/ft to 8.3 × 10(exp 6)/ft and angles of attack from 0 deg to 18 deg. Boundary-layer transition onset and global surface heating distribution measurements were performed using phosphor thermography and flow field images were obtained through schlieren photography. Surface deflections were found to both promote early transition of the boundary layer and to augment heating levels for both laminar and turbulent flows. A complimentary computational flow field study was also performed to provide heating predictions for comparison with the measurements as well as boundary layer flow field properties for use in correlating the data. Correlations of the wind tunnel data were developed to predict deflection effects on boundary layer transition and surface heating and were applied to both the wind tunnel test conditions and to the trajectory of NASA's successful IRVE-3 flight test. In general, the correlations produced at least qualitative agreement with the wind tunnel data, although the heating levels were underpredicted for some of the larger surface deflections. For the flight conditions, the correlations suggested that peak heating levels on the leeward side conical flank of the IRVE-3 vehicle may have exceeded those at nose for times late in the trajectory after the peak heating time point. However, the flight estimates were based on a conservative assumption of surface deflection magnitude (i.e., larger) than likely was produced in flight.

Detached Eddy Simulations of Hypersonic Transition

Science.gov (United States)

Yoon, S.; Barnhardt, M.; Candler, G.

2010-01-01

This slide presentation reviews the use of Detached Eddy Simulation (DES) of hypersonic transistion. The objective of the study was to investigate the feasibility of using CFD in general, DES in particular, for prediction of roughness-induced boundary layer transition to turbulence and the resulting increase in heat transfer.

High burn rate solid composite propellants

Science.gov (United States)

Manship, Timothy D.

High burn rate propellants help maintain high levels of thrust without requiring complex, high surface area grain geometries. Utilizing high burn rate propellants allows for simplified grain geometries that not only make production of the grains easier, but the simplified grains tend to have better mechanical strength, which is important in missiles undergoing high-g accelerations. Additionally, high burn rate propellants allow for a higher volumetric loading which reduces the overall missile's size and weight. The purpose of this study is to present methods of achieving a high burn rate propellant and to develop a composite propellant formulation that burns at 1.5 inches per second at 1000 psia. In this study, several means of achieving a high burn rate propellant were presented. In addition, several candidate approaches were evaluated using the Kepner-Tregoe method with hydroxyl terminated polybutadiene (HTPB)-based propellants using burn rate modifiers and dicyclopentadiene (DCPD)-based propellants being selected for further evaluation. Propellants with varying levels of nano-aluminum, nano-iron oxide, FeBTA, and overall solids loading were produced using the HTPB binder and evaluated in order to determine the effect the various ingredients have on the burn rate and to find a formulation that provides the burn rate desired. Experiments were conducted to compare the burn rates of propellants using the binders HTPB and DCPD. The DCPD formulation matched that of the baseline HTPB mix. Finally, GAP-plasticized DCPD gumstock dogbones were attempted to be made for mechanical evaluation. Results from the study show that nano-additives have a substantial effect on propellant burn rate with nano-iron oxide having the largest influence. Of the formulations tested, the highest burn rate was a 84% solids loading mix using nano-aluminum nano-iron oxide, and ammonium perchlorate in a 3:1(20 micron: 200 micron) ratio which achieved a burn rate of 1.2 inches per second at 1000

The engineering of a nuclear thermal landing and ascent vehicle utilizing indigenous Martian propellant

International Nuclear Information System (INIS)

Zubrin, R.M.

1991-01-01

The following paper reports on a design study of a novel space transportation concept known as a ''NIMF'' (Nuclear rocket using Indigenous Martian Fuel.) The NIMF is a ballistic vehicle which obtains its propellant out of the Martian air by compression and liquefaction of atmospheric CO 2 . This propellant is subsequently used to generate rocket thrust at a specific impulse of 264 s by being heated to high temperature (2800 K) gas in the NIMFs' nuclear thermal rocket engines. The vehicle is designed to provide surface to orbit and surface to surface transportation, as well as housing, for a crew of three astronauts. It is capable of refueling itself for a flight to its maximum orbit in less than 50 days. The ballistic NIMF has a mass of 44.7 tonnes and, with the assumed 2800 K propellant temperature, is capable of attaining highly energetic (250 km by 34000 km elliptical) orbits. This allows it to rendezvous with interplanetary transfer vehicles which are only very loosely bound into orbit around Mars. If a propellant temperature of 2000 K is assumed, then low Mars orbit can be attained; while if 3100 K is assumed, then the ballistic NIMF is capable of injecting itself onto a minimum energy transfer orbit to Earth in a direct ascent from the Martian surface

Shape-Memory-Alloy Actuator For Flight Controls

Science.gov (United States)

Barret, Chris

1995-01-01

Report proposes use of shape-memory-alloy actuators, instead of hydraulic actuators, for aerodynamic flight-control surfaces. Actuator made of shape-memory alloy converts thermal energy into mechanical work by changing shape as it makes transitions between martensitic and austenitic crystalline phase states of alloy. Because both hot exhaust gases and cryogenic propellant liquids available aboard launch rockets, shape-memory-alloy actuators exceptionally suited for use aboard such rockets.

Computation of hypersonic flows with finite rate condensation and evaporation of water

Science.gov (United States)

Perrell, Eric R.; Candler, Graham V.; Erickson, Wayne D.; Wieting, Alan R.

1993-01-01

A computer program for modelling 2D hypersonic flows of gases containing water vapor and liquid water droplets is presented. The effects of interphase mass, momentum and energy transfer are studied. Computations are compared with existing quasi-1D calculations on the nozzle of the NASA Langley Eight Foot High Temperature Tunnel, a hypersonic wind tunnel driven by combustion of natural gas in oxygen enriched air.

Ceramic Matrix Composite (CMC) Thermal Protection Systems (TPS) and Hot Structures for Hypersonic Vehicles

Science.gov (United States)

Glass, David E.

2008-01-01

Thermal protection systems (TPS) and hot structures are required for a range of hypersonic vehicles ranging from ballistic reentry to hypersonic cruise vehicles, both within Earth's atmosphere and non-Earth atmospheres. The focus of this paper is on air breathing hypersonic vehicles in the Earth's atmosphere. This includes single-stage to orbit (SSTO), two-stage to orbit (TSTO) accelerators, access to space vehicles, and hypersonic cruise vehicles. This paper will start out with a brief discussion of aerodynamic heating and thermal management techniques to address the high heating, followed by an overview of TPS for rocket-launched and air-breathing vehicles. The argument is presented that as we move from rocket-based vehicles to air-breathing vehicles, we need to move away from the insulated airplane approach used on the Space Shuttle Orbiter to a wide range of TPS and hot structure approaches. The primary portion of the paper will discuss issues and design options for CMC TPS and hot structure components, including leading edges, acreage TPS, and control surfaces. The current state-of-the-art will be briefly discussed for some of the components. The two primary technical challenges impacting the use of CMC TPS and hot structures for hypersonic vehicles are environmental durability and fabrication, and will be discussed briefly.

Preliminary Structural Sensitivity Study of Hypersonic Inflatable Aerodynamic Decelerator Using Probabilistic Methods

Science.gov (United States)

Lyle, Karen H.

2014-01-01

Acceptance of new spacecraft structural architectures and concepts requires validated design methods to minimize the expense involved with technology validation via flighttesting. This paper explores the implementation of probabilistic methods in the sensitivity analysis of the structural response of a Hypersonic Inflatable Aerodynamic Decelerator (HIAD). HIAD architectures are attractive for spacecraft deceleration because they are lightweight, store compactly, and utilize the atmosphere to decelerate a spacecraft during re-entry. However, designers are hesitant to include these inflatable approaches for large payloads or spacecraft because of the lack of flight validation. In the example presented here, the structural parameters of an existing HIAD model have been varied to illustrate the design approach utilizing uncertainty-based methods. Surrogate models have been used to reduce computational expense several orders of magnitude. The suitability of the design is based on assessing variation in the resulting cone angle. The acceptable cone angle variation would rely on the aerodynamic requirements.

Portable Fluorescence Imaging System for Hypersonic Flow Facilities

Science.gov (United States)

Wilkes, J. A.; Alderfer, D. W.; Jones, S. B.; Danehy, P. M.

2003-01-01

A portable fluorescence imaging system has been developed for use in NASA Langley s hypersonic wind tunnels. The system has been applied to a small-scale free jet flow. Two-dimensional images were taken of the flow out of a nozzle into a low-pressure test section using the portable planar laser-induced fluorescence system. Images were taken from the center of the jet at various test section pressures, showing the formation of a barrel shock at low pressures, transitioning to a turbulent jet at high pressures. A spanwise scan through the jet at constant pressure reveals the three-dimensional structure of the flow. Future capabilities of the system for making measurements in large-scale hypersonic wind tunnel facilities are discussed.

Hypersonic force measurements using internal balance based on optical micromachined Fabry-Perot interferometry

Science.gov (United States)

Qiu, Huacheng; Min, Fu; Zhong, Shaolong; Song, Xin; Yang, Yanguang

2018-03-01

Force measurements using wind tunnel balance are necessary for determining a variety of aerodynamic performance parameters, while the harsh environment in hypersonic flows requires that the measurement instrument should be reliable and robust, in against strong electromagnetic interference, high vacuum, or metal (oxide) dusts. In this paper, we demonstrated a three-component internal balance for hypersonic aerodynamic force measurements, using novel optical micromachined Fabry-Perot interferometric (FPI) strain gauges as sensing elements. The FPI gauges were fabricated using Micro-Opto-Electro-Mechanical Systems (MOEMS) surface and bulk fabrication techniques. High-reflectivity coatings are used to form a high-finesse Fabry-Perot cavity, which benefits a high resolution. Antireflective and passivation coatings are used to reduce unwanted interferences. The FPI strain gauge based balance has been calibrated and evaluated in a Mach 5 hypersonic flow. The results are compared with the traditional technique using the foil resistive strain gauge balance, indicating that the proposed balance based on the MOEMS FPI strain gauge is reliable and robust and is potentially suitable for the hypersonic wind tunnel harsh environment.

MPD model for radar echo signal of hypersonic targets

Directory of Open Access Journals (Sweden)

Xu Xuefei

2014-08-01

Full Text Available The stop-and-go (SAG model is typically used for echo signal received by the radar using linear frequency modulation pulse compression. In this study, the authors demonstrate that this model is not applicable to hypersonic targets. Instead of SAG model, they present a more realistic echo signal model (moving-in-pulse duration (MPD for hypersonic targets. Following that, they evaluate the performances of pulse compression under the SAG and MPD models by theoretical analysis and simulations. They found that the pulse compression gain has an increase of 3 dB by using the MPD model compared with the SAG model in typical cases.

Anisotropic hypersonic phonon propagation in films of aligned ellipsoids.

Science.gov (United States)

Beltramo, Peter J; Schneider, Dirk; Fytas, George; Furst, Eric M

2014-11-14

A material with anisotropic elastic mechanical properties and a direction-dependent hypersonic band gap is fabricated using ac electric field-directed convective self-assembly of colloidal ellipsoids. The frequency of the gap, which is detected in the direction perpendicular to particle alignment and entirely absent parallel to alignment, and the effective sound velocities can be tuned by the particle aspect ratio. We hypothesize that the band gap originates from the primary eigenmode peak, the m-splitted (s,1,2) mode, of the particle resonating with the effective medium. These results reveal the potential for powerful control of the hypersonic phononic band diagram by combining anisotropic particles and self-assembly.

Analysis of Hypersonic Vehicle Wakes

Science.gov (United States)

2015-09-17

Fraction of Cyanide throughout the Flowfield ................................... 131 Figure 122. Mass Fraction of Cyanide at the Nose...hypersonic flow is that as M increases the conservation equations cannot be linearized. The flow properties must be modeled in a complex fashion and can no...ablation present to react with as well. These products of ablation, along with the dissociation and ionization of the gas, gives rise to complex

Hypersonic Threats to the Homeland

Science.gov (United States)

2017-03-28

ADAM) system . This ground based system protects 7 soldiers against rocket threats and utilizes a 10 kW laser with an effective range out to...early warning systems for response to hypersonic threats . The integration of directed energy defensive systems with Space Based Infrared Sensors (SBIRS...and early warning radars already in operation will save costs. By capitalizing on Terminal High Altitude Area Defense (THAAD) system capabilities

Experimental analysis and simulation of the dynamic response of a propeller pitch change actuator

OpenAIRE

LECLERCQ, Maxime; MALBURET, François; VERON, Philippe

2012-01-01

This paper focuses specifically on the control of the propeller pitch change mechanisms and their associated dynamics. The subject of this article is restricted to the mechanisms using a hydraulic single acting actuator. They function asymmetrically and are subject to important varying external loads under the full flight envelope. This phenomenon has an impact on their dynamic response.The question of the dynamics of these systems is rarely dealt with because, usually for aircraft applicatio...

Experimental analysis and simulation of the dynamix response of a propeller pitch change actuator

OpenAIRE

LECLERCQ, Maxime; MALBURET, François; VERON, Philippe

2012-01-01

This paper focuses specifically on the control of the propeller pitch change mechanisms and their associated dynamics. The subject of this article is restricted to the mechanisms using a hydraulic single acting actuator. They function asymmetrically and are subject to important varying external loads under the full flight envelope. This phenomenon has an impact on their dynamic response. The question of the dynamics of these systems is rarely dealt with because, usually for aircraft applic...

On two special values of temperature factor in hypersonic flow stagnation point

Science.gov (United States)

Bilchenko, G. G.; Bilchenko, N. G.

2018-03-01

The hypersonic aircraft permeable cylindrical and spherical surfaces laminar boundary layer heat and mass transfer control mathematical model properties are investigated. The nonlinear algebraic equations systems are obtained for two special values of temperature factor in the hypersonic flow stagnation point. The mappings bijectivity between heat and mass transfer local parameters and controls is established. The computation experiments results are presented: the domains of allowed values “heat-friction” are obtained.

Efficient multigrid computation of steady hypersonic flows

NARCIS (Netherlands)

Koren, B.; Hemker, P.W.; Murthy, T.K.S.

1991-01-01

In steady hypersonic flow computations, Newton iteration as a local relaxation procedure and nonlinear multigrid iteration as an acceleration procedure may both easily fail. In the present chapter, same remedies are presented for overcoming these problems. The equations considered are the steady,

Dynamics Evolution Investigation of Mack Mode Instability in a Hypersonic Boundary Layer by Bicoherence Spectrum Analysis

Science.gov (United States)

Han, Jian; Jiang, Nan

2012-07-01

The instability of a hypersonic boundary layer on a cone is investigated by bicoherence spectrum analysis. The experiment is conducted at Mach number 6 in a hypersonic wind tunnel. The time series signals of instantaneous fluctuating surface-thermal-flux are measured by Pt-thin-film thermocouple temperature sensors mounted at 28 stations on the cone surface along streamwise direction to investigate the development of the unstable disturbances. The bicoherence spectrum analysis based on wavelet transform is employed to investigate the nonlinear interactions of the instability of Mack modes in hypersonic laminar boundary layer transition. The results show that wavelet bicoherence is a powerful tool in studying the unstable mode nonlinear interaction of hypersonic laminar-turbulent transition. The first mode instability gives rise to frequency shifts to higher unstable modes at the early stage of hypersonic laminar-turbulent transition. The modulations subsequently lead to the second mode instability occurrence. The second mode instability governs the last stage of instability and final breakdown to turbulence with multi-scale disturbances growth.

Dynamics Evolution Investigation of Mack Mode Instability in a Hypersonic Boundary Layer by Bicoherence Spectrum Analysis

International Nuclear Information System (INIS)

Han Jian; Jiang Nan

2012-01-01

The instability of a hypersonic boundary layer on a cone is investigated by bicoherence spectrum analysis. The experiment is conducted at Mach number 6 in a hypersonic wind tunnel. The time series signals of instantaneous fluctuating surface-thermal-flux are measured by Pt-thin-film thermocouple temperature sensors mounted at 28 stations on the cone surface along streamwise direction to investigate the development of the unstable disturbances. The bicoherence spectrum analysis based on wavelet transform is employed to investigate the nonlinear interactions of the instability of Mack modes in hypersonic laminar boundary layer transition. The results show that wavelet bicoherence is a powerful tool in studying the unstable mode nonlinear interaction of hypersonic laminar-turbulent transition. The first mode instability gives rise to frequency shifts to higher unstable modes at the early stage of hypersonic laminar-turbulent transition. The modulations subsequently lead to the second mode instability occurrence. The second mode instability governs the last stage of instability and final breakdown to turbulence with multi-scale disturbances growth. (fundamental areas of phenomenology(including applications))

Development of HAN-based Liquid Propellant Thruster

Science.gov (United States)

Hisatsune, K.; Izumi, J.; Tsutaya, H.; Furukawa, K.

2004-10-01

Many of propellants that are applied to the conventional spacecraft propulsion system are toxic propellants. Because of its toxicity, considering the environmental pollution or safety on handling, it will be necessary to apply the "green" propellant to the spacecraft propulsion system. The purpose of this study is to apply HAN based liquid propellant (LP1846) to mono propellant thruster. Compared to the hydrazine that is used in conventional mono propellant thruster, HAN based propellant is not only lower toxic but also can obtain higher specific impulse. Moreover, HAN based propellant can be decomposed by the catalyst. It means there are the possibility of applying to the mono propellant thruster that can leads to the high reliability of the propulsion system.[1],[2] However, there are two technical subjects, to apply HAN based propellant to the mono propellant thruster. One is the high combustion temperature. The catalyst will be damaged under high temperature condition. The other is the low catalytic activity. It is the serious problem on application of HAN based propellant to the mono propellant thruster that is used for attitude control of spacecraft. To improve the catalytic activity of HAN based propellant, it is necessary to screen the best catalyst for HAN based propellant. The adsorption analysis is conducted by Monte Carlo Simulation to screen the catalyst metal for HAN and TEAN. The result of analysis shows the Iridium is the best catalyst metal for HAN and TEAN. Iridium is the catalyst metal that is used at conventional mono propellant thruster catalyst Shell405. Then, to confirm the result of analysis, the reaction test about catalyst is conducted. The result of this test is the same as the result of adsorption analysis. That means the adsorption analysis is effective in screening the catalyst metal. At the evaluating test, the various types of carrier of catalyst are also compared to Shell 405 to improve catalytic activity. The test result shows the

14 CFR 35.23 - Propeller control system.

Science.gov (United States)

2010-01-01

... propeller effect under the intended operating conditions. (4) The failure or corruption of data or signals... corruption of airplane-supplied data does not result in hazardous propeller effects. (e) The propeller... effect. (2) Failures or malfunctions directly affecting the propeller control system in a typical...

Wavelet Cross-Spectrum Analysis of Multi-Scale Disturbance Instability and Transition on Sharp Cone Hypersonic Boundary Layer

International Nuclear Information System (INIS)

Jian, Han; Nan, Jiang

2008-01-01

Experimental measurement of hypersonic boundary layer stability and transition on a sharp cone with a half angle of 5° is carried out at free-coming stream Mach number 6 in a hypersonic wind tunnel. Mean and fluctuation surface-thermal-flux characteristics of the hypersonic boundary layer flow are measured by Pt-thin-film thermocouple temperature sensors installed at 28 stations on the cone surface along longitudinal direction. At hypersonic speeds, the dominant flow instabilities demonstrate that the growth rate of the second mode tends to exceed that of the low-frequency mode. Wavelet-based cross-spectrum technique is introduced to obtain the multi-scale cross-spectral characteristics of the fluctuating signals in the frequency range of the second mode. Nonlinear interactions both of the second mode disturbance and the first mode disturbance are demonstrated to be dominant instabilities in the initial stage of laminar-turbulence transition for hypersonic shear flow. (fundamental areas of phenomenology (including applications))

Poor flight performance in deep-diving cormorants.

Science.gov (United States)

Watanabe, Yuuki Y; Takahashi, Akinori; Sato, Katsufumi; Viviant, Morgane; Bost, Charles-André

2011-02-01

Aerial flight and breath-hold diving present conflicting morphological and physiological demands, and hence diving seabirds capable of flight are expected to face evolutionary trade-offs regarding locomotory performances. We tested whether Kerguelen shags Phalacrocorax verrucosus, which are remarkable divers, have poor flight capability using newly developed tags that recorded their flight air speed (the first direct measurement for wild birds) with propeller sensors, flight duration, GPS position and depth during foraging trips. Flight air speed (mean 12.7 m s(-1)) was close to the speed that minimizes power requirement, rather than energy expenditure per distance, when existing aerodynamic models were applied. Flights were short (mean 92 s), with a mean summed duration of only 24 min day(-1). Shags sometimes stayed at the sea surface without diving between flights, even on the way back to the colony, and surface durations increased with the preceding flight durations; these observations suggest that shags rested after flights. Our results indicate that their flight performance is physiologically limited, presumably compromised by their great diving capability (max. depth 94 m, duration 306 s) through their morphological adaptations for diving, including large body mass (enabling a large oxygen store), small flight muscles (to allow for large leg muscles for underwater propulsion) and short wings (to decrease air volume in the feathers and hence buoyancy). The compromise between flight and diving, as well as the local bathymetry, shape the three-dimensional foraging range (<26 km horizontally, <94 m vertically) in this bottom-feeding cormorant.

A computational fluid dynamics simulation of the hypersonic flight of the Pegasus(TM) vehicle using an artificial viscosity model and a nonlinear filtering method. M.S. Thesis

Science.gov (United States)

Mendoza, John Cadiz

1995-01-01

The computational fluid dynamics code, PARC3D, is tested to see if its use of non-physical artificial dissipation affects the accuracy of its results. This is accomplished by simulating a shock-laminar boundary layer interaction and several hypersonic flight conditions of the Pegasus(TM) launch vehicle using full artificial dissipation, low artificial dissipation, and the Engquist filter. Before the filter is applied to the PARC3D code, it is validated in one-dimensional and two-dimensional form in a MacCormack scheme against the Riemann and convergent duct problem. For this explicit scheme, the filter shows great improvements in accuracy and computational time as opposed to the nonfiltered solutions. However, for the implicit PARC3D code it is found that the best estimate of the Pegasus experimental heat fluxes and surface pressures is the simulation utilizing low artificial dissipation and no filter. The filter does improve accuracy over the artificially dissipative case but at a computational expense greater than that achieved by the low artificial dissipation case which has no computational time penalty and shows better results. For the shock-boundary layer simulation, the filter does well in terms of accuracy for a strong impingement shock but not as well for weaker shock strengths. Furthermore, for the latter problem the filter reduces the required computational time to convergence by 18.7 percent.

Aeroelasticity, Aerothermoelasticity and Aeroelastic Scaling of Hypersonic Vehicles

National Research Council Canada - National Science Library

Freidmann, Peretz P; Powell, Kenneth G

2004-01-01

...) the behavior of a complete generic hypersonic vehicle. For problems (a) the unsteady airloads were computed using third order piston theory, as well a CFD based Euler and Navier-Stokes loads. For case (b...

Second-mode control in hypersonic boundary layers over assigned complex wall impedance

Science.gov (United States)

Sousa, Victor; Patel, Danish; Chapelier, Jean-Baptiste; Scalo, Carlo

2017-11-01

The durability and aerodynamic performance of hypersonic vehicles greatly relies on the ability to delay transition to turbulence. Passive aerodynamic flow control devices such as porous acoustic absorbers are a very attractive means to damp ultrasonic second-mode waves, which govern transition in hypersonic boundary layers under idealized flow conditions (smooth walls, slender geometries, small angles of attack). The talk will discuss numerical simulations modeling such absorbers via the time-domain impedance boundary condition (TD-IBC) approach by Scalo et al. in a hypersonic boundary layer flow over a 7-degree wedge at freestream Mach numbers M∞ = 7.3 and Reynolds numbers Rem = 1.46 .106 . A three-parameter impedance model tuned to the second-mode waves is tested first with varying resistance, R, and damping ratio, ζ, revealing complete mode attenuation for R workers at DLR-Göttingen.

Robust adaptive multivariable higher-order sliding mode flight control for air-breathing hypersonic vehicle with actuator failures

Directory of Open Access Journals (Sweden)

Peng Li

2016-10-01

Full Text Available This article proposes an adaptive multivariable higher-order sliding mode control for the longitudinal model of an air-breathing vehicle under system uncertainties and actuator failures. Firstly, a fast finite-time control law is designed for a chain of integrators. Secondly, based on the input/output feedback linearization technique, the system uncertainty and external disturbances are modeled as additive certainty and the actuator failures are modeled as multiplicative uncertainty. By using the proposed fast finite-time control law, a robust multivariable higher-order sliding mode control is designed for the air-breathing hypersonic vehicle with actuator failures. Finally, adaptive laws are proposed for the adaptation of the parameters in the robust multivariable higher-order sliding mode control. Thus, the bounds of the uncertainties are not needed in the control system design. Simulation results show the effectiveness of the proposed robust adaptive multivariable higher-order sliding mode control.

Progress with multigrid schemes for hypersonic flow problems

International Nuclear Information System (INIS)

Radespiel, R.; Swanson, R.C.

1995-01-01

Several multigrid schemes are considered for the numerical computation of viscous hypersonic flows. For each scheme, the basic solution algorithm employs upwind spatial discretization with explicit multistage time stepping. Two-level versions of the various multigrid algorithms are applied to the two-dimensional advection equation, and Fourier analysis is used to determine their damping properties. The capabilities of the multigrid methods are assessed by solving three different hypersonic flow problems. Some new multigrid schemes based on semicoarsening strategies are shown to be quite effective in relieving the stiffness caused by the high-aspect-ratio cells required to resolve high Reynolds number flows. These schemes exhibit good convergence rates for Reynolds numbers up to 200 X 10 6 and Mach numbers up to 25. 32 refs., 31 figs., 1 tab

Hypersonic nozzle/afterbody CFD code validation. I - Experimental measurements

Science.gov (United States)

Spaid, Frank W.; Keener, Earl R.

1993-01-01

This study was conducted to obtain a detailed experimental description of the flow field created by the interaction of a single-expansion-ramp-nozzle flow with a hypersonic external stream. Data were obtained from a generic nozzle/afterbody model in the 3.5-Foot Hypersonic Wind Tunnel of the NASA Ames Research Center in a cooperative experimental program involving Ames and the McDonnell Douglas Research Laboratories. This paper presents experimental results consisting primarily of surveys obtained with a five-hole total-pressure/flow-direction probe and a total-temperature probe. These surveys were obtained in the flow field created by the interaction between the underexpanded jet plume and the external flow.

Hydrodynamics of Ship Propellers

DEFF Research Database (Denmark)

Breslin, John P.; Andersen, Poul

This book deals with flows over propellers operating behind ships, and the hydrodynamic forces and moments which the propeller generates on the shaft and on the ship hull.The first part of the text is devoted to fundamentals of the flow about hydrofoil sections (with and without cavitation...... of an intermittently cavitating propeller in a wake and the pressures and forces it exerts on the shaft and on the ship hull is examined. A final chapter discusses the optimization of efficiency of compound propulsors. The authors have taken care to clearly describe physical concepts and mathematical steps. Appendices...

Modeling of Supersonic Combustion Systems for Sustained Hypersonic Flight

Directory of Open Access Journals (Sweden)

Stephen M. Neill

2017-11-01

Full Text Available Through Computational Fluid Dynamics and validation, an optimal scramjet combustor has been designed based on twin-strut Hydrogen injection to sustain flight at a desired speed of Mach 8. An investigation undertaken into the efficacy of supersonic combustion through various means of injection saw promising results for Hydrogen-based systems, whereby strut-style injectors were selected over transverse injectors based on their pressure recovery performance and combustive efficiency. The final configuration of twin-strut injectors provided robust combustion and a stable region of net thrust (1873 kN in the nozzle. Using fixed combustor inlet parameters and injection equivalence ratio, the finalized injection method advanced to the early stages of two-dimensional (2-D and three-dimensional (3-D scramjet engine integration. The overall investigation provided a feasible supersonic combustion system, such that Mach 8 sustained cruise could be achieved by the aircraft concept in a computational design domain.

Solid propellant impact tests

International Nuclear Information System (INIS)

Snow, E.C.

1976-03-01

Future space missions, as in the past, call for the continued use of radioisotopes as heat sources for thermoelectric power generators. In an effort to minimize the risk of radioactive contamination of the environment, a complete safety analysis of each such system is necessary. As a part of these analyses, the effects on such a system of a solid propellant fire environment resulting from a catastrophic launch pad abort must be considered. Several impact tests were conducted in which either a simulant MHW-FSA or a steel ball was dropped on the cold, unignited or the hot, burning surface of a block of UTP-3001 solid propellant. The rebound velocities were measured for both surface conditions of the propellant. The resulting coefficient of restitution, determined as the ratio of the components of the impact and rebound velocities perpendicular to the impact surface of the propellant, were not very dependent on whether the surface was cold or hot at the time of impact

Propeller and inflow vortex interaction : vortex response and impact on the propeller performance

NARCIS (Netherlands)

Yang, Y.; Zhou, T; Sciacchitano, A.; Veldhuis, L.L.M.; Eitelberg, G.

2016-01-01

The aerodynamic operating conditions of a propeller can include complex situations where vorticity from sources upstream can enter the propeller plane. In general, when the vorticity enters in a concentrated form of a vortex, the interaction between the vortex and blade is referred to as

New Propellants and Cryofuels

Science.gov (United States)

Palasezski, Bryan; Sullivan, Neil S.; Hamida, Jaha; Kokshenev, V.

2006-01-01

The proposed research will investigate the stability and cryogenic properties of solid propellants that are critical to NASA s goal of realizing practical propellant designs for future spacecraft. We will determine the stability and thermal properties of a solid hydrogen-liquid helium stabilizer in a laboratory environment in order to design a practical propellant. In particular, we will explore methods of embedding atomic species and metallic nano-particulates in hydrogen matrices suspended in liquid helium. We will also measure the characteristic lifetimes and diffusion of atomic species in these candidate cryofuels. The most promising large-scale advance in rocket propulsion is the use of atomic propellants; most notably atomic hydrogen stabilized in cryogenic environments, and metallized-gelled liquid hydrogen (MGH) or densified gelled hydrogen (DGH). The new propellants offer very significant improvements over classic liquid oxygen/hydrogen fuels because of two factors: (1) the high energy-release, and (ii) the density increase per unit energy release. These two changes can lead to significant reduced mission costs and increased payload to orbit weight ratios. An achievable 5 to 10 percent improvement in specific impulse for the atomic propellants or MGH fuels can result in a doubling or tripling of system payloads. The high-energy atomic propellants must be stored in a stabilizing medium such as solid hydrogen to inhibit or delay their recombination into molecules. The goal of the proposed research is to determine the stability and thermal properties of the solid hydrogen-liquid helium stabilizer. Magnetic resonance techniques will be used to measure the thermal lifetimes and the diffusive motions of atomic species stored in solid hydrogen grains. The properties of metallic nano-particulates embedded in hydrogen matrices will also be studied and analyzed. Dynamic polarization techniques will be developed to enhance signal/noise ratios in order to be able to

A reduced order aerothermodynamic modeling framework for hypersonic vehicles based on surrogate and POD

Directory of Open Access Journals (Sweden)

Chen Xin

2015-10-01

Full Text Available Aerothermoelasticity is one of the key technologies for hypersonic vehicles. Accurate and efficient computation of the aerothermodynamics is one of the primary challenges for hypersonic aerothermoelastic analysis. Aimed at solving the shortcomings of engineering calculation, computation fluid dynamics (CFD and experimental investigation, a reduced order modeling (ROM framework for aerothermodynamics based on CFD predictions using an enhanced algorithm of fast maximin Latin hypercube design is developed. Both proper orthogonal decomposition (POD and surrogate are considered and compared to construct ROMs. Two surrogate approaches named Kriging and optimized radial basis function (ORBF are utilized to construct ROMs. Furthermore, an enhanced algorithm of fast maximin Latin hypercube design is proposed, which proves to be helpful to improve the precisions of ROMs. Test results for the three-dimensional aerothermodynamic over a hypersonic surface indicate that: the ROMs precision based on Kriging is better than that by ORBF, ROMs based on Kriging are marginally more accurate than ROMs based on POD-Kriging. In a word, the ROM framework for hypersonic aerothermodynamics has good precision and efficiency.

Advanced Control System Design for Hypersonic Vehicles, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Guidance and control system design for hypersonic vehicles is more challenging than their subsonic and supersonic counterparts. Some of these challenges are (i)...

Parametric Analysis of a Hypersonic Inlet using Computational Fluid Dynamics

Science.gov (United States)

Oliden, Daniel

For CFD validation, hypersonic flow fields are simulated and compared with experimental data specifically designed to recreate conditions found by hypersonic vehicles. Simulated flow fields on a cone-ogive with flare at Mach 7.2 are compared with experimental data from NASA Ames Research Center 3.5" hypersonic wind tunnel. A parametric study of turbulence models is presented and concludes that the k-kl-omega transition and SST transition turbulence model have the best correlation. Downstream of the flare's shockwave, good correlation is found for all boundary layer profiles, with some slight discrepancies of the static temperature near the surface. Simulated flow fields on a blunt cone with flare above Mach 10 are compared with experimental data from CUBRC LENS hypervelocity shock tunnel. Lack of vibrational non-equilibrium calculations causes discrepancies in heat flux near the leading edge. Temperature profiles, where non-equilibrium effects are dominant, are compared with the dissociation of molecules to show the effects of dissociation on static temperature. Following the validation studies is a parametric analysis of a hypersonic inlet from Mach 6 to 20. Compressor performance is investigated for numerous cowl leading edge locations up to speeds of Mach 10. The variable cowl study showed positive trends in compressor performance parameters for a range of Mach numbers that arise from maximizing the intake of compressed flow. An interesting phenomenon due to the change in shock wave formation for different Mach numbers developed inside the cowl that had a negative influence on the total pressure recovery. Investigation of the hypersonic inlet at different altitudes is performed to study the effects of Reynolds number, and consequently, turbulent viscous effects on compressor performance. Turbulent boundary layer separation was noted as the cause for a change in compressor performance parameters due to a change in Reynolds number. This effect would not be

Evaluation of CFD Turbulent Heating Prediction Techniques and Comparison With Hypersonic Experimental Data

Science.gov (United States)

Dilley, Arthur D.; McClinton, Charles R. (Technical Monitor)

2001-01-01

Results from a study to assess the accuracy of turbulent heating and skin friction prediction techniques for hypersonic applications are presented. The study uses the original and a modified Baldwin-Lomax turbulence model with a space marching code. Grid converged turbulent predictions using the wall damping formulation (original model) and local damping formulation (modified model) are compared with experimental data for several flat plates. The wall damping and local damping results are similar for hot wall conditions, but differ significantly for cold walls, i.e., T(sub w) / T(sub t) hypersonic vehicles. Based on the results of this study, it is recommended that the local damping formulation be used with the Baldwin-Lomax and Cebeci-Smith turbulence models in design and analysis of Hyper-X and future hypersonic vehicles.

NASA Marshall Space Flight Center Controls Systems Design and Analysis Branch

Science.gov (United States)

Gilligan, Eric

2014-01-01

Marshall Space Flight Center maintains a critical national capability in the analysis of launch vehicle flight dynamics and flight certification of GN&C algorithms. MSFC analysts are domain experts in the areas of flexible-body dynamics and control-structure interaction, thrust vector control, sloshing propellant dynamics, and advanced statistical methods. Marshall's modeling and simulation expertise has supported manned spaceflight for over 50 years. Marshall's unparalleled capability in launch vehicle guidance, navigation, and control technology stems from its rich heritage in developing, integrating, and testing launch vehicle GN&C systems dating to the early Mercury-Redstone and Saturn vehicles. The Marshall team is continuously developing novel methods for design, including advanced techniques for large-scale optimization and analysis.

Systems Challenges for Hypersonic Vehicles

Science.gov (United States)

Hunt, James L.; Laruelle, Gerard; Wagner, Alain

1997-01-01

This paper examines the system challenges posed by fully reusable hypersonic cruise airplanes and access to space vehicles. Hydrocarbon and hydrogen fueled airplanes are considered with cruise speeds of Mach 5 and 10, respectively. The access to space matrix is examined. Airbreathing and rocket powered, single- and two-stage vehicles are considered. Reference vehicle architectures are presented. Major systems/subsystems challenges are described. Advanced, enhancing systems concepts as well as common system technologies are discussed.

The engineering of a nuclear thermal landing and ascent vehicle utilizing indigenous Martian propellant

Science.gov (United States)

Zubrin, Robert M.

1990-01-01

A design study of a novel space transportation concept called NIMF (Nuclear rocket using Indigenous Martian Fuel) is reported. In this concept, Martian CO2 gas, which constitutes 95 percent of the atmosphere, is liquified by simple compression to about 100 psi and remains stable without refrigeration. When heated and exhausted out of a rocket nozzle, a specific impulse of about 264 s can be achieved, sufficient for flights from the surface to highly energetic orbits or from one point on the surface to any other point. The propellant acquisition system can travel with the vehicle, allowing it to refuel itself each time it lands. The concept offers unequalled potential to achieve planetwide mobility, allowing complete global access for the exploration of Mars. By eliminating the necessity of transporting ascent propellant to Mars, the NIMF can also significantly reduce the initial mass in LEO and of a manned Mars mission.

Flexible Thermal Protection System Development for Hypersonic Inflatable Aerodynamic Decelerators

Science.gov (United States)

DelCorso, Joseph A.; Bruce, Walter E., III; Hughes, Stephen J.; Dec, John A.; Rezin, Marc D.; Meador, Mary Ann B.; Guo, Haiquan; Fletcher, Douglas G.; Calomino, Anthony M.; Cheatwood, McNeil

2012-01-01

The Hypersonic Inflatable Aerodynamic Decelerators (HIAD) project has invested in development of multiple thermal protection system (TPS) candidates to be used in inflatable, high downmass, technology flight projects. Flexible TPS is one element of the HIAD project which is tasked with the research and development of the technology ranging from direct ground tests, modelling and simulation, characterization of TPS systems, manufacturing and handling, and standards and policy definition. The intent of flexible TPS is to enable large deployable aeroshell technologies, which increase the drag performance while significantly reducing the ballistic coefficient of high-mass entry vehicles. A HIAD requires a flexible TPS capable of surviving aerothermal loads, and durable enough to survive the rigors of construction, handling, high density packing, long duration exposure to extrinsic, in-situ environments, and deployment. This paper provides a comprehensive overview of key work being performed within the Flexible TPS element of the HIAD project. Included in this paper is an overview of, and results from, each Flexible TPS research and development activity, which includes ground testing, physics-based thermal modelling, age testing, margins policy, catalysis, materials characterization, and recent developments with new TPS materials.

Dynamic Testing of the NASA Hypersonic Project Combined Cycle Engine Testbed for Mode Transition Experiments

Science.gov (United States)

2011-01-01

NASA is interested in developing technology that leads to more routine, safe, and affordable access to space. Access to space using airbreathing propulsion systems has potential to meet these objectives based on Airbreathing Access to Space (AAS) system studies. To this end, the NASA Fundamental Aeronautics Program (FAP) Hypersonic Project is conducting fundamental research on a Turbine Based Combined Cycle (TBCC) propulsion system. The TBCC being studied considers a dual flow-path inlet system. One flow-path includes variable geometry to regulate airflow to a turbine engine cycle. The turbine cycle provides propulsion from take-off to supersonic flight. The second flow-path supports a dual-mode scramjet (DMSJ) cycle which would be initiated at supersonic speed to further accelerate the vehicle to hypersonic speed. For a TBCC propulsion system to accelerate a vehicle from supersonic to hypersonic speed, a critical enabling technology is the ability to safely and effectively transition from the turbine to the DMSJ-referred to as mode transition. To experimentally test methods of mode transition, a Combined Cycle Engine (CCE) Large-scale Inlet testbed was designed with two flow paths-a low speed flow-path sized for a turbine cycle and a high speed flow-path designed for a DMSJ. This testbed system is identified as the CCE Large-Scale Inlet for Mode Transition studies (CCE-LIMX). The test plan for the CCE-LIMX in the NASA Glenn Research Center (GRC) 10- by 10-ft Supersonic Wind Tunnel (10x10 SWT) is segmented into multiple phases. The first phase is a matrix of inlet characterization (IC) tests to evaluate the inlet performance and establish the mode transition schedule. The second phase is a matrix of dynamic system identification (SysID) experiments designed to support closed-loop control development at mode transition schedule operating points for the CCE-LIMX. The third phase includes a direct demonstration of controlled mode transition using a closed loop control

Shock-tunnel combustor testing for hypersonic vehicles

Science.gov (United States)

Loomis, Mark P.

1994-01-01

Proposed configurations for the next generation of transatmospheric vehicles will rely on air breathing propulsion systems during all or part of their mission. At flight Mach numbers greater than about 7 these engines will operate in the supersonic combustion ramjet mode (scramjet). Ground testing of these engine concepts above Mach 8 requires high pressure, high enthalpy facilities such as shock tunnels and expansion tubes. These impulse, or short duration facilities have test times on the order of a millisecond, requiring high speed instrumentation and data systems. One such facility ideally suited for scramjet testing is the NASA-Ames 16-Inch shock tunnel, which over the last two years has completed a series of tests for the NASP (National Aero-Space Plane) program at simulated flight Mach numbers ranging from 12-16. The focus of the experimental programs consisted of a series of classified tests involving a near-full scale hydrogen fueled scramjet combustor model in the semi-free jet method of engine testing whereby the compressed forebody flow ahead of the cowl inlet is reproduced (see appendix A). The AIMHYE-1 (Ames Integrated Modular Hypersonic Engine) test entry for the NASP program was completed in April 1993, while AIMHYE-2 was completed in May 1994. The test entries were regarded as successful, resulting in some of the first data of its kind on the performance of a near full scale scramjet engine at Mach 12-16. The data was distributed to NASP team members for use in design system verification and development. Due to the classified nature of the hardware and data, the data reports resulting from this work are classified and have been published as part of the NASP literature. However, an unclassified AIAA paper resulted from the work and has been included as appendix A. It contains an overview of the test program and a description of some of the important issues.

Molecular Diagnostics for the Study of Hypersonic Flows

Science.gov (United States)

2000-04-01

UNCLASSIFIED Defense Technical Information Center Compilation Part Notice ADPO10744 TITLE: Molecular Diagnostics for the Study of Hypersonic Flows...following component part numbers comprise the compilation report: ADP010736 thru ADPO10751 UNCLASSIFIED 5-1 Molecular Diagnostics for the Study of

Application of a Near Infrared Imaging System for Thermographic Imaging of the Space Shuttle during Hypersonic Re-Entry

Science.gov (United States)

Zalameda, Joseph N.; Tietjen, Alan B.; Horvath, Thomas J.; Tomek, Deborah M.; Gibson, David M.; Taylor, Jeff C.; Tack, Steve; Bush, Brett C.; Mercer, C. David; Shea, Edward J.

2010-01-01

High resolution calibrated near infrared (NIR) imagery was obtained of the Space Shuttle s reentry during STS-119, STS-125, and STS-128 missions. The infrared imagery was collected using a US Navy NP-3D Orion aircraft using a long-range infrared optical package referred to as Cast Glance. The slant ranges between the Space Shuttle and Cast Glance were approximately 26-41 nautical miles at point of closest approach. The Hypersonic Thermodynamic Infrared Measurements (HYTHIRM) project was a NASA Langley led endeavor sponsored by the NASA Engineering Safety Center, the Space Shuttle Program Office and the NASA Aeronautics Research Mission Directorate to demonstrate a quantitative thermal imaging capability. HYTHIRM required several mission tools to acquire the imagery. These tools include pre-mission acquisition simulations of the Shuttle trajectory in relationship to the Cast Glance aircraft flight path, radiance modeling to predict the infrared response of the Shuttle, and post mission analysis tools to process the infrared imagery to quantitative temperature maps. The spatially resolved global thermal measurements made during the Shuttle s hypersonic reentry provides valuable flight data for reducing the uncertainty associated with present day ground-to-flight extrapolation techniques and current state-of-the-art empirical boundary-layer transition or turbulent heating prediction methods. Laminar and turbulent flight data is considered critical for the development of turbulence models supporting NASA s next-generation spacecraft. This paper will provide the motivation and details behind the use of an upgraded NIR imaging system used onboard a Navy Cast Glance aircraft and describe the characterizations and procedures performed to obtain quantitative temperature maps. A brief description and assessment will be provided of the previously used analog NIR camera along with image examples from Shuttle missions STS-121, STS-115, and solar tower test. These thermal

Cavitation simulation on marine propellers

DEFF Research Database (Denmark)

Shin, Keun Woo

Cavitation on marine propellers causes thrust breakdown, noise, vibration and erosion. The increasing demand for high-efficiency propellers makes it difficult to avoid the occurrence of cavitation. Currently, practical analysis of propeller cavitation depends on cavitation tunnel test, empirical...... criteria and inviscid flow method, but a series of model test is costly and the other two methods have low accuracy. Nowadays, computational fluid dynamics by using a viscous flow solver is common for practical industrial applications in many disciplines. Cavitation models in viscous flow solvers have been...... hydrofoils and conventional/highly-skewed propellers are performed with one of three cavitation models proven in 2D analysis. 3D cases also show accuracy and robustness of numerical method in simulating steady and unsteady sheet cavitation on complicated geometries. Hydrodynamic characteristics of cavitation...

On Challenges for Hypersonic Turbulent Simulations

International Nuclear Information System (INIS)

Yee, H.C.; Sjogreen, B.

2009-01-01

This short note discusses some of the challenges for design of suitable spatial numerical schemes for hypersonic turbulent flows, including combustion, and thermal and chemical nonequilibrium flows. Often, hypersonic turbulent flows in re-entry space vehicles and space physics involve mixed steady strong shocks and turbulence with unsteady shocklets. Material mixing in combustion poses additional computational challenges. Proper control of numerical dissipation in numerical methods beyond the standard shock-capturing dissipation at discontinuities is an essential element for accurate and stable simulations of the subject physics. On one hand, the physics of strong steady shocks and unsteady turbulence/shocklet interactions under the nonequilibrium environment is not well understood. On the other hand, standard and newly developed high order accurate (fourth-order or higher) schemes were developed for homogeneous hyperbolic conservation laws and mixed hyperbolic and parabolic partial differential equations (PDEs) (without source terms). The majority of finite rate chemistry and thermal nonequilibrium simulations employ methods for homogeneous time-dependent PDEs with a pointwise evaluation of the source terms. The pointwise evaluation of the source term might not be the best choice for stability, accuracy and minimization of spurious numerics for the overall scheme

Theseus Landing Following Maiden Flight

Science.gov (United States)

1996-01-01

The Theseus prototype research aircraft shows off its high aspect-ratio wing as it comes in for a landing on Rogers Dry Lake after its first test flight from NASA's Dryden Flight Research Center, Edwards, California, on May 24, 1996. The Theseus aircraft, built and operated by Aurora Flight Sciences Corporation, Manassas, Virginia, was a unique aircraft flown at NASA's Dryden Flight Research Center, Edwards, California, under a cooperative agreement between NASA and Aurora. Dryden hosted the Theseus program, providing hangar space and range safety for flight testing. Aurora Flight Sciences was responsible for the actual flight testing, vehicle flight safety, and operation of the aircraft. The Theseus remotely piloted aircraft flew its maiden flight on May 24, 1996, at Dryden. During its sixth flight on November 12, 1996, Theseus experienced an in-flight structural failure that resulted in the loss of the aircraft. As of the beginning of the year 2000, Aurora had not rebuilt the aircraft. Theseus was built for NASA under an innovative, $4.9 million fixed-price contract by Aurora Flight Sciences Corporation and its partners, West Virginia University, Morgantown, West Virginia, and Fairmont State College, Fairmont, West Virginia. The twin-engine, unpiloted vehicle had a 140-foot wingspan, and was constructed largely of composite materials. Powered by two 80-horsepower, turbocharged piston engines that drove twin 9-foot-diameter propellers, Theseus was designed to fly autonomously at high altitudes, with takeoff and landing under the active control of a ground-based pilot in a ground control station 'cockpit.' With the potential ability to carry 700 pounds of science instruments to altitudes above 60,000 feet for durations of greater than 24 hours, Theseus was intended to support research in areas such as stratospheric ozone depletion and the atmospheric effects of future high-speed civil transport aircraft engines. Instruments carried aboard Theseus also would be able

Performance evaluation and design of flight vehicle control systems

CERN Document Server

Falangas, Eric T

2015-01-01

This book will help students, control engineers and flight dynamics analysts to model and conduct sophisticated and systemic analyses of early flight vehicle designs controlled with multiple types of effectors and to design and evaluate new vehicle concepts in terms of satisfying mission and performance goals. Performance Evaluation and Design of Flight Vehicle Control Systems begins by creating a dynamic model of a generic flight vehicle that includes a range of elements from airplanes and launch vehicles to re-entry vehicles and spacecraft. The models may include dynamic effects dealing with structural flexibility, as well as dynamic coupling between structures and actuators, propellant sloshing, and aeroelasticity, and they are typically used for control analysis and design. The book shows how to efficiently combine different types of effectors together, such as aero-surfaces, TVC, throttling engines and RCS, to operate as a system by developing a mixing logic atrix. Methods of trimming a vehicle controll...

Thermal Vacuum Test Correlation of a Zero Propellant Load Case Thermal Capacitance Propellant Gauging Analytical Model

Science.gov (United States)

Mckim, Stephen A.

2016-01-01

This thesis describes the development and correlation of a thermal model that forms the foundation of a thermal capacitance spacecraft propellant load estimator. Specific details of creating the thermal model for the diaphragm propellant tank used on NASA's Magnetospheric Multiscale spacecraft using ANSYS and the correlation process implemented are presented. The thermal model was correlated to within plus or minus 3 degrees Celsius of the thermal vacuum test data, and was determined sufficient to make future propellant predictions on MMS. The model was also found to be relatively sensitive to uncertainties in applied heat flux and mass knowledge of the tank. More work is needed to improve temperature predictions in the upper hemisphere of the propellant tank where predictions were found to be 2 to 2.5 C lower than the test data. A road map for applying the model to predict propellant loads on the actual MMS spacecraft toward its end of life in 2017-2018 is also presented.

Drag and Torque on Locked Screw Propeller

Directory of Open Access Journals (Sweden)

Tomasz Tabaczek

2014-09-01

Full Text Available Few data on drag and torque on locked propeller towed in water are available in literature. Those data refer to propellers of specific geometry (number of blades, blade area, pitch and skew of blades. The estimation of drag and torque of an arbitrary propeller considered in analysis of ship resistance or propulsion is laborious. The authors collected and reviewed test data available in the literature. Based on collected data there were developed the empirical formulae for estimation of hydrodynamic drag and torque acting on locked screw propeller. Supplementary CFD computations were carried out in order to prove the applicability of the formulae to modern moderately skewed screw propellers.

Noise from Two-Blade Propellers

Science.gov (United States)

Stowell, E Z; Deming, A F

1936-01-01

The two-blade propeller, one of the most powerful sources of sound known, has been studied with the view of obtaining fundamental information concerning the noise emission. In order to eliminate engine noise, the propeller was mounted on an electric motor. A microphone was used to pick up the sound whose characteristics were studied electrically. The distribution of noise throughout the frequency range, as well as the spatial distribution about the propeller, was studied. The results are given in the form of polar diagrams. An appendix of common acoustical terms is included.

A reduced order aerothermodynamic modeling framework for hypersonic vehicles based on surrogate and POD

OpenAIRE

Chen Xin; Liu Li; Long Teng; Yue Zhenjiang

2015-01-01

Aerothermoelasticity is one of the key technologies for hypersonic vehicles. Accurate and efficient computation of the aerothermodynamics is one of the primary challenges for hypersonic aerothermoelastic analysis. Aimed at solving the shortcomings of engineering calculation, computation fluid dynamics (CFD) and experimental investigation, a reduced order modeling (ROM) framework for aerothermodynamics based on CFD predictions using an enhanced algorithm of fast maximin Latin hypercube design ...

Assessment of predictive capabilities for aerodynamic heating in hypersonic flow

Science.gov (United States)

Knight, Doyle; Chazot, Olivier; Austin, Joanna; Badr, Mohammad Ali; Candler, Graham; Celik, Bayram; Rosa, Donato de; Donelli, Raffaele; Komives, Jeffrey; Lani, Andrea; Levin, Deborah; Nompelis, Ioannis; Panesi, Marco; Pezzella, Giuseppe; Reimann, Bodo; Tumuklu, Ozgur; Yuceil, Kemal

2017-04-01

The capability for CFD prediction of hypersonic shock wave laminar boundary layer interaction was assessed for a double wedge model at Mach 7.1 in air and nitrogen at 2.1 MJ/kg and 8 MJ/kg. Simulations were performed by seven research organizations encompassing both Navier-Stokes and Direct Simulation Monte Carlo (DSMC) methods as part of the NATO STO AVT Task Group 205 activity. Comparison of the CFD simulations with experimental heat transfer and schlieren visualization suggest the need for accurate modeling of the tunnel startup process in short-duration hypersonic test facilities, and the importance of fully 3-D simulations of nominally 2-D (i.e., non-axisymmmetric) experimental geometries.

Development of Fast Deterministic Physically Accurate Solvers for Kinetic Collision Integral for Applications of Near Space Flight and Control Devices

Science.gov (United States)

2015-08-31

flow non-equilibrium is important, such as hypersonic flight at moderate to high altitudes and rocket propulsion with plume - atmosphere interaction...d2=1.8990E-4 kg/m 3 , v2 =281.098, in the case of the Mach 1.55 wave, and d1=1.0676E-4 kg/m 3 , v1=1225.851 m/s, T1=300 K and d2=3.35386E-4 kg/m 3 , v2

Propeller Flaps: A Literature Review.

Science.gov (United States)

Sisti, Andrea; D'Aniello, Carlo; Fortezza, Leonardo; Tassinari, Juri; Cuomo, Roberto; Grimaldi, Luca; Nisi, Giuseppe

2016-01-01

Since their introduction in 1991, propeller flaps are increasingly used as a surgical approach to loss of substance. The aim of this study was to evaluate the indications and to verify the outcomes and the complication rates using this reconstructing technique through a literature review. A search on PubMed was performed using "propeller flap", "fasciocutaneous flap", "local flap" or "pedicled flap" as key words. We selected clinical studies using propeller flaps as a reconstructing technique. We found 119 studies from 1991 to 2015. Overall, 1,315 propeller flaps were reported in 1,242 patients. Most frequent indications included loss of substance following tumor excision, repair of trauma-induced injuries, burn scar contractures, pressure sores and chronic infections. Complications were observed in 281/1242 patients (22.6%) occurring more frequently in the lower limbs (31.8%). Partial flap necrosis and venous congestion were the most frequent complications. The complications' rate was significantly higher in infants (70 years old) but there was not a significant difference between the sexes. Trend of complication rate has not improved during the last years. Propeller flaps showed a great success rate with low morbidity, quick recovery, good aesthetic outcomes and reduced cost. The quality and volume of the transferred soft tissue, the scar orientation and the possibility of direct donor site closure should be considered in order to avoid complications. Indications for propeller flaps are small- or medium-sized defects located in a well-vascularized area with healthy surrounding tissues. Copyright © 2016 International Institute of Anticancer Research (Dr. John G. Delinassios), All rights reserved.

Thermal Design of Vapor Cooling of Flight Vehicle Structures Using LH2 Boil-Off

Science.gov (United States)

Wang, Xiao-Yen; Zoeckler, Joseph

2015-01-01

Using hydrogen boil-off vapor to cool the structure of a flight vehicle cryogenic upper stage can reduce heat loads to the stage and increase the usable propellant in the stage or extend the life of the stage. The hydrogen vapor can be used to absorb incoming heat as it increases in temperature before being vented overboard. In theory, the amount of heat leaking into the hydrogen tank from the structure will be reduced if the structure is cooled using the propellant boil-off vapor. However, the amount of boil-off vapor available to be used for cooling and the reduction in heat leak to the propellant tank are dependent to each other. The amount of heat leak reduction to the LH2 tank also depends on the total heat load on the stage and the vapor cooling configurations.

Disposal of Liquid Propellants

Science.gov (United States)

1990-03-13

propellant includes an oxi- dizer (hydroxylammoniuin nitrate), a fuel (triethanolammonium nitrate), and water . In an- ticipation of widespread (both...are also included. 20. DISTRIBUTION/ AVAILABILIT ’." OF ABMTRACT 21 ABSTRACT SECURITY CLASSIF.CATICIN IUNCLASSIFIEDIUNLIMITED 0 SAME AS RPT. 0 OTIC...trieth- anolammoiur nitrate), anG water . In anticipation of widespread (both conti- nental U.S. and abroac) use of the propellant, USATHAMA began a

Armstrong Flight Research Center Research Technology and Engineering 2017

Science.gov (United States)

Voracek, David F. (Editor)

2018-01-01

I am delighted to present this report of accomplishments at NASA's Armstrong Flight Research Center. Our dedicated innovators possess a wealth of performance, safety, and technical capabilities spanning a wide variety of research areas involving aircraft, electronic sensors, instrumentation, environmental and earth science, celestial observations, and much more. They not only perform tasks necessary to safely and successfully accomplish Armstrong's flight research and test missions but also support NASA missions across the entire Agency. Armstrong's project teams have successfully accomplished many of the nation's most complex flight research projects by crafting creative solutions that advance emerging technologies from concept development and experimental formulation to final testing. We are developing and refining technologies for ultra-efficient aircraft, electric propulsion vehicles, a low boom flight demonstrator, air launch systems, and experimental x-planes, to name a few. Additionally, with our unique location and airborne research laboratories, we are testing and validating new research concepts. Summaries of each project highlighting key results and benefits of the effort are provided in the following pages. Technology areas for the projects include electric propulsion, vehicle efficiency, supersonics, space and hypersonics, autonomous systems, flight and ground experimental test technologies, and much more. Additional technical information is available in the appendix, as well as contact information for the Principal Investigator of each project. I am proud of the work we do here at Armstrong and am pleased to share these details with you. We welcome opportunities for partnership and collaboration, so please contact us to learn more about these cutting-edge innovations and how they might align with your needs.

A theoretical and experimental investigation of propeller performance methodologies

Science.gov (United States)

Korkan, K. D.; Gregorek, G. M.; Mikkelson, D. C.

1980-01-01

This paper briefly covers aspects related to propeller performance by means of a review of propeller methodologies; presentation of wind tunnel propeller performance data taken in the NASA Lewis Research Center 10 x 10 wind tunnel; discussion of the predominent limitations of existing propeller performance methodologies; and a brief review of airfoil developments appropriate for propeller applications.

Numerical simulation for the influence of laser-induced plasmas addition on air mass capture of hypersonic inlet

Science.gov (United States)

Zhao, Wei; Dou, Zhiguo; Li, Qian

2012-03-01

The theory of laser-induced plasmas addition to hypersonic airflow off a vehicle to increase air mass capture and improve the performance of hypersonic inlets at Mach numbers below the design value is explored. For hypersonic vehicles, when flying at mach numbers lower than the design one, we can increase the mass capture ratio of inlet through laser-induced plasmas injection to the hypersonic flow upstream of cowl lip to form a virtual cowl. Based on the theory, the model of interaction between laser-induced plasmas and hypersonic flow was established. The influence on the effect of increasing mass capture ratio was studied at different positions of laser-induced plasmas region for the external compression hypersonic inlet at Mach 5 while the design value is 6, the power of plasmas was in the range of 1-8mJ. The main results are as follows: 1. the best location of the plasma addition region is near the intersection of the nose shock of the vehicle with the continuation of the cowl line, and slightly below that line. In that case, the shock generated by the heating is close to the shock that is a reflection of the vehicle nose shock off the imaginary solid surface-extension of the cowl. 2. Plasma addition does increase mass capture, and the effect becomes stronger as more energy is added, the peak value appeared when the power of plasma was about 4mJ, when the plasma energy continues to get stronger, the mass capture will decline slowly.

A Combined CFD/Characteristic Method for Prediction and Design of Hypersonic Inlet with Nose Bluntness

Science.gov (United States)

Gao, Wenzhi; Li, Zhufei; Yang, Jiming

Leading edge bluntness is widely used in hypersonic inlet design for thermal protection[1]. Detailed research of leading edge bluntness on hypersonic inlet has been concentrated on shock shape correlation[2], boundary layer flow[3], inlet performance[4], etc. It is well known that blunted noses cause detached bow shocks which generate subsonic regions around the noses and entropy layers in the flowfield.

Theseus on Take-off for First Flight

Science.gov (United States)

1996-01-01

The Theseus prototype research aircraft takes off for its first test flight from NASA's Dryden Flight Research Center, Edwards, California, on May 24, 1996. The Theseus aircraft, built and operated by Aurora Flight Sciences Corporation, Manassas, Virginia, was a unique aircraft flown at NASA's Dryden Flight Research Center, Edwards, California, under a cooperative agreement between NASA and Aurora. Dryden hosted the Theseus program, providing hangar space and range safety for flight testing. Aurora Flight Sciences was responsible for the actual flight testing, vehicle flight safety, and operation of the aircraft. The Theseus remotely piloted aircraft flew its maiden flight on May 24, 1996, at Dryden. During its sixth flight on November 12, 1996, Theseus experienced an in-flight structural failure that resulted in the loss of the aircraft. As of the beginning of the year 2000, Aurora had not rebuilt the aircraft. Theseus was built for NASA under an innovative, $4.9 million fixed-price contract by Aurora Flight Sciences Corporation and its partners, West Virginia University, Morgantown, West Virginia, and Fairmont State College, Fairmont, West Virginia. The twin-engine, unpiloted vehicle had a 140-foot wingspan, and was constructed largely of composite materials. Powered by two 80-horsepower, turbocharged piston engines that drove twin 9-foot-diameter propellers, Theseus was designed to fly autonomously at high altitudes, with takeoff and landing under the active control of a ground-based pilot in a ground control station 'cockpit.' With the potential ability to carry 700 pounds of science instruments to altitudes above 60,000 feet for durations of greater than 24 hours, Theseus was intended to support research in areas such as stratospheric ozone depletion and the atmospheric effects of future high-speed civil transport aircraft engines. Instruments carried aboard Theseus also would be able to validate satellite-based global environmental change measurements. Dryden

Cavitation noise studies on marine propellers

Science.gov (United States)

Sharma, S. D.; Mani, K.; Arakeri, V. H.

1990-04-01

Experimental observations are described of cavitation inception and noise from five model propellers, three basic and two modified, tested in the open jet section of the Indian Institute of Science high-speed water tunnel facility. Extensive experiments on the three basic propellers of different design, which included visualization of cavitation and measurements of noise, showed that the dominant type of cavitation was in the form of tip vortex cavitation, accompanied by leading edge suction side sheet cavitation in its close vicinity, and the resultant noise depended on parameters such as the advance coefficient, the cavitation number, and the propeller geometry. Of these, advance coefficient was found to have the maximum influence not only on cavitation noise but also on the inception of cavitation. Noise levels and frequencies of spectra obtained from all the three basic propellers at conditions near inception and different advance coefficient values, when plotted in the normalized form as suggested by Blake, resulted in a universal spectrum which would be useful for predicting cavitation noise at prototype scales when a limited extent of cavitation is expected in the same form as observed on the present models. In an attempt to delay the onset of tip vortex cavitation, the blades of two of the three basic propellers were modified by drilling small holes in the tip and leading edge areas. Studies on the modified propellers showed that the effectiveness of the blade modification was apparently stronger at low advance coefficient values and depended on the blade sectional profile. Measurements of cavitation noise indicated that the modification also improved the acoustic performance of the propellers as it resulted in a complete attenuation of the low-frequency spectral peaks, which were prominent with the basic propellers. In addition to the above studies, which were conducted under uniform flow conditions, one of the basic propellers was tested in the simulated

Characteristic Model-Based Robust Model Predictive Control for Hypersonic Vehicles with Constraints

Directory of Open Access Journals (Sweden)

Jun Zhang

2017-06-01

Full Text Available Designing robust control for hypersonic vehicles in reentry is difficult, due to the features of the vehicles including strong coupling, non-linearity, and multiple constraints. This paper proposed a characteristic model-based robust model predictive control (MPC for hypersonic vehicles with reentry constraints. First, the hypersonic vehicle is modeled by a characteristic model composed of a linear time-varying system and a lumped disturbance. Then, the identification data are regenerated by the accumulative sum idea in the gray theory, which weakens effects of the random noises and strengthens regularity of the identification data. Based on the regenerated data, the time-varying parameters and the disturbance are online estimated according to the gray identification. At last, the mixed H2/H∞ robust predictive control law is proposed based on linear matrix inequalities (LMIs and receding horizon optimization techniques. Using active tackling system constraints of MPC, the input and state constraints are satisfied in the closed-loop control system. The validity of the proposed control is verified theoretically according to Lyapunov theory and illustrated by simulation results.

Downstream Effects on Orbiter Leeside Flow Separation for Hypersonic Flows

Science.gov (United States)

Buck, Gregory M.; Pulsonetti, Maria V.; Weilmuenster, K. James

2005-01-01

Discrepancies between experiment and computation for shuttle leeside flow separation, which came to light in the Columbia accident investigation, are resolved. Tests were run in the Langley Research Center 20-Inch Hypersonic CF4 Tunnel with a baseline orbiter model and two extended trailing edge models. The extended trailing edges altered the wing leeside separation lines, moving the lines toward the fuselage, proving that wing trailing edge modeling does affect the orbiter leeside flow. Computations were then made with a wake grid. These calculations more closely matched baseline experiments. Thus, the present findings demonstrate that it is imperative to include the wake flow domain in CFD calculations in order to accurately predict leeside flow separation for hypersonic vehicles at high angles of attack.

Examination of uniform momentum zones in hypersonic turbulent boundary layers

Science.gov (United States)

Williams, Owen; Helm, Clara; Martin, Pino

2017-11-01

The presence of uniform momentum zones (UMZs) separated by regions of high shear is now well-established in incompressible flows, with the mean number of such zones increasing in a log-linear fashion with Reynolds number. While known to be present in supersonic and hypersonic boundary layers, the properties of these UMZs and the appropriate Reynolds number for comparison with incompressible results have not previously been investigated. A large, previously published DNS database of hypersonic boundary layers is used in this investigation, with Mach numbers up to 12 and wall temperatures from cold to adiabatic, resulting in a wide range of outer layer Reynolds numbers. UMZs are examined using a range of parameters in both conventional inner and semi-local scalings, and Reynolds number trends examined.

Resonant influence of a longitudinal hypersonic field on the radiation from channeled electrons

International Nuclear Information System (INIS)

Grigoryan, L.Sh.; Mkrtchyan, A.R.; Mkrtchyan, A.H.; Khachatryan, H.F.; Prade, H.; Wagner, W.; Piestrup, M.A.

2001-01-01

The wave function of a planar/axially channeled electron with energy 10 MeV≤E<<1 GeV under the influence of a longitudinal hypersonic wave excited in a single crystal is calculated. Conditions for the resonant influence of the hypersonic wave on the quantum state of the channeled electron are deduced. Expressions for the wave function that are applicable in the case of resonance are obtained. Angular and spectral distributions of the radiation intensity from the planar/axially channeled electron are also calculated. The possibility of significant amplification of channeling radiation by a hypersonic wave is substantiated. It is found that the hypersound can excite inverse radiative transitions through which the transversal energy of the channeled electron is increased. These transitions have a resonant nature and can lead to a considerable intensification of the electron channeling radiation. In the case of axial channeling, the resonance radiation is sustained also by direct radiative transitions of the electron

Tests on thirteen navy type model propellers

Science.gov (United States)

Durand, W F

1927-01-01

The tests on these model propellers were undertaken for the purpose of determining the performance coefficients and characteristics for certain selected series of propellers of form and type as commonly used in recent navy designs. The first series includes seven propellers of pitch ratio varying by 0.10 to 1.10, the area, form of blade, thickness, etc., representing an arbitrary standard propeller which had shown good results. The second series covers changes in thickness of blade section, other things equal, and the third series, changes in blade area, other things equal. These models are all of 36-inch diameter. Propellers A to G form the series on pitch ratio, C, N. I. J the series on thickness of section, and K, M, C, L the series on area. (author)

Theseus First Flight - May 24, 1996

Science.gov (United States)

1996-01-01

The Theseus prototype research aircraft shows off its high aspect-ratio wing as it lifts off from Rogers Dry Lake during its first test flight from NASA's Dryden Flight Research Center, Edwards, California, on May 24, 1996. The Theseus aircraft, built and operated by Aurora Flight Sciences Corporation, Manassas, Virginia, was a unique aircraft flown at NASA's Dryden Flight Research Center, Edwards, California, under a cooperative agreement between NASA and Aurora. Dryden hosted the Theseus program, providing hangar space and range safety for flight testing. Aurora Flight Sciences was responsible for the actual flight testing, vehicle flight safety, and operation of the aircraft. The Theseus remotely piloted aircraft flew its maiden flight on May 24, 1996, at Dryden. During its sixth flight on November 12, 1996, Theseus experienced an in-flight structural failure that resulted in the loss of the aircraft. As of the beginning of the year 2000, Aurora had not rebuilt the aircraft. Theseus was built for NASA under an innovative, $4.9 million fixed-price contract by Aurora Flight Sciences Corporation and its partners, West Virginia University, Morgantown, West Virginia, and Fairmont State College, Fairmont, West Virginia. The twin-engine, unpiloted vehicle had a 140-foot wingspan, and was constructed largely of composite materials. Powered by two 80-horsepower, turbocharged piston engines that drove twin 9-foot-diameter propellers, Theseus was designed to fly autonomously at high altitudes, with takeoff and landing under the active control of a ground-based pilot in a ground control station 'cockpit.' With the potential ability to carry 700 pounds of science instruments to altitudes above 60,000 feet for durations of greater than 24 hours, Theseus was intended to support research in areas such as stratospheric ozone depletion and the atmospheric effects of future high-speed civil transport aircraft engines. Instruments carried aboard Theseus also would be able to

Hypersonic modulation of light in three-dimensional photonic and phononic band-gap materials.

Science.gov (United States)

Akimov, A V; Tanaka, Y; Pevtsov, A B; Kaplan, S F; Golubev, V G; Tamura, S; Yakovlev, D R; Bayer, M

2008-07-18

The elastic coupling between the a-SiO2 spheres composing opal films brings forth three-dimensional periodic structures which besides a photonic stop band are predicted to also exhibit complete phononic band gaps. The influence of elastic crystal vibrations on the photonic band structure has been studied by injection of coherent hypersonic wave packets generated in a metal transducer by subpicosecond laser pulses. These studies show that light with energies close to the photonic band gap can be efficiently modulated by hypersonic waves.

The use of computer-generated color graphic images for transient thermal analysis. [for hypersonic aircraft

Science.gov (United States)

Edwards, C. L. W.; Meissner, F. T.; Hall, J. B.

1979-01-01

Color computer graphics techniques were investigated as a means of rapidly scanning and interpreting large sets of transient heating data. The data presented were generated to support the conceptual design of a heat-sink thermal protection system (TPS) for a hypersonic research airplane. Color-coded vector and raster displays of the numerical geometry used in the heating calculations were employed to analyze skin thicknesses and surface temperatures of the heat-sink TPS under a variety of trajectory flight profiles. Both vector and raster displays proved to be effective means for rapidly identifying heat-sink mass concentrations, regions of high heating, and potentially adverse thermal gradients. The color-coded (raster) surface displays are a very efficient means for displaying surface-temperature and heating histories, and thereby the more stringent design requirements can quickly be identified. The related hardware and software developments required to implement both the vector and the raster displays for this application are also discussed.

Large Propellant Tank Cryo-Cooler (LPTC)

Data.gov (United States)

National Aeronautics and Space Administration — In rocket test and launch facilities, cryogenic propellants stored in tanks boils off due to heat leakage, with the following impacts:Ø   Waste, propellants boil off...

Seawater Immersion of GEM II Propellant

National Research Council Canada - National Science Library

Merrill, Calude

1999-01-01

... (% AP lost/week aged in seawater) and intercepts that depend on sample size. Friction and impact data on dried aged propellant samples showed no increased burning hazard compared with propellant not exposed to water...

Effect of surface roughness on the heating rates of large-angled hypersonic blunt cones

Science.gov (United States)

Irimpan, Kiran Joy; Menezes, Viren

2018-03-01

Surface-roughness caused by the residue of an ablative Thermal Protection System (TPS) can alter the turbulence level and surface heating rates on a hypersonic re-entry capsule. Large-scale surface-roughness that could represent an ablated TPS, was introduced over the forebody of a 120° apex angle blunt cone, in order to test for its influence on surface heating rates in a hypersonic freestream of Mach 8.8. The surface heat transfer rates measured on smooth and roughened models under the same freestream conditions were compared. The hypersonic flow-fields of the smooth and rough-surfaced models were visualized to analyse the flow physics. Qualitative numerical simulations and pressure measurements were carried out to have an insight into the high-speed flow physics. Experimental observations under moderate Reynolds numbers indicated a delayed transition and an overall reduction of 17-46% in surface heating rates on the roughened model.

Radiation from channeled positrons in a hypersonic wave field

International Nuclear Information System (INIS)

Mkrtchyan, A.R.; Gasparyan, R.A.; Gabrielyan, R.G.

1987-01-01

The radiation emitted by channeled positrons in a longitudinal or transverse standing hypersonic wave field is considered. In the case of plane channeling the spectral distribution of the radiation intensity is shown to be of a resonance nature depending on the hypersound frequency

Marshall Space Flight Center Faculty Fellowship Program

Science.gov (United States)

Six, N. F.; Karr, G.

2017-01-01

The research projects conducted by the 2016 Faculty Fellows at NASA Marshall Space Flight Center included propulsion studies on propellant issues, and materials investigations involving plasma effects and friction stir welding. Spacecraft Systems research was conducted on wireless systems and 3D printing of avionics. Vehicle Systems studies were performed on controllers and spacecraft instruments. The Science and Technology group investigated additive construction applied to Mars and Lunar regolith, medical uses of 3D printing, and unique instrumentation, while the Test Laboratory measured pressure vessel leakage and crack growth rates.

78 FR 4038 - Critical Parts for Airplane Propellers

Science.gov (United States)

2013-01-18

... manufacturers are not required to provide information concerning propeller critical part design, manufacture, or... engineering, manufacturing, and service management processes should provide clear information for propeller... manufacture critical parts for airplane propellers update their manuals to record engineering, manufacture...

Ionic liquid propellants: future fuels for space propulsion.

Science.gov (United States)

Zhang, Qinghua; Shreeve, Jean'ne M

2013-11-11

Use of green propellants is a trend for future space propulsion. Hypergolic ionic liquid propellants, which are environmentally-benign while exhibiting energetic performances comparable to hydrazine, have shown great potential to meet the requirements of developing nontoxic high-performance propellant formulations for space propulsion applications. This Concept article presents a review of recent advances in the field of ionic liquid propellants. Copyright © 2013 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim.

Engineering method for aero-propulsive characteristics at hypersonic Mach numbers

Science.gov (United States)

Goradia, Suresh; Torres, Abel O.; Stack, Sharon H.; Everhart, Joel L.

1991-01-01

An engineering method has been developed for the rapid analysis of external aerodynamics and propulsive performance characteristics of airbreathing vehicles at hypersonic Mach numbers. This method, based on the theory of characteristics, has been developed to analyze fuselage-wing body combinations and body flaps with blunt or sharp leading/trailing edges. Arbitrary ratio of specific heat for the flowing medium can be specified in the program. Furthermore, the capability exists in the code to compute the inviscid inlet mass capture and momentum flux. The method is under development for computations of pressure distribution, and flow characteristics in the inlet, along with the effect of viscosity. Correlative studies have been performed for representative hypersonic configurations using the current method. The results of these correlations for various aerodynamics parameters are encouraging.

Analysis of Windward Side Hypersonic Boundary Layer Transition on Blunted Cones at Angle of Attack

Science.gov (United States)

2017-01-09

correlated with PSE/LST N-Factors. 15. SUBJECT TERMS boundary layer transition, hypersonic, ground test 16. SECURITY CLASSIFICATION OF: 17. LIMITATION ...Maccoll) solution e condition at boundary layer edge w condition at wall, viscous ∞ condition in freestream Conventions LST Linear Stability Theory PSE...STATES AIR FORCE AFRL-RQ-WP-TP-2017-0169 ANALYSIS OF WINDWARD SIDE HYPERSONIC BOUNDARY LAYER TRANSITION ON BLUNTED CONES AT ANGLE OF ATTACK Roger

Flight Tests of a 0.13-Scale Model of the Convair XFY-1 Vertically Rising Airplane in a Setup Simulating that Proposed for Captive-Flight Tests in a Hangar, TED No. NACA DE 368

Science.gov (United States)

Lovell, Powell M., Jr.

1953-01-01

An experimental investigation has been conducted to determine the dynamic stability and control characteristics of a 0.13-scale free-flight model of the Convair XFY-1 airplane in test setups representing the setup proposed for use in the first flight tests of the full-scale airplane in the Moffett Field airship hangar. The investigation was conducted in two parts: first, tests with the model flying freely in an enclosure simulating the hangar, and second, tests with the model partially restrained by an overhead line attached to the propeller spinner and ground lines attached to the wing and tail tips. The results of the tests indicated that the airplane can be flown without difficulty in the Moffett Field airship hangar if it does not approach too close to the hangar walls. If it does approach too close to the walls, the recirculation of the propeller slipstream might cause sudden trim changes which would make smooth flight difficult for the pilot to accomplish. It appeared that the tethering system proposed by Convair could provide generally satisfactory restraint of large-amplitude motions caused by control failure or pilot error without interfering with normal flying or causing any serious instability or violent jerking motions as the tethering lines restrained the model.

A Collaborative Analysis Tool for Integrating Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

Science.gov (United States)

Stanley, Thomas Troy; Alexander, Reginald

1999-01-01

Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. The deficiencies in the scramjet powered concept led to a revival of interest in Rocket-Based Combined-Cycle (RBCC) propulsion systems. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. At this point the transitions to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scram4jet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance.

An engineering code to analyze hypersonic thermal management systems

Science.gov (United States)

Vangriethuysen, Valerie J.; Wallace, Clark E.

1993-01-01

Thermal loads on current and future aircraft are increasing and as a result are stressing the energy collection, control, and dissipation capabilities of current thermal management systems and technology. The thermal loads for hypersonic vehicles will be no exception. In fact, with their projected high heat loads and fluxes, hypersonic vehicles are a prime example of systems that will require thermal management systems (TMS) that have been optimized and integrated with the entire vehicle to the maximum extent possible during the initial design stages. This will not only be to meet operational requirements, but also to fulfill weight and performance constraints in order for the vehicle to takeoff and complete its mission successfully. To meet this challenge, the TMS can no longer be two or more entirely independent systems, nor can thermal management be an after thought in the design process, the typical pervasive approach in the past. Instead, a TMS that was integrated throughout the entire vehicle and subsequently optimized will be required. To accomplish this, a method that iteratively optimizes the TMS throughout the vehicle will not only be highly desirable, but advantageous in order to reduce the manhours normally required to conduct the necessary tradeoff studies and comparisons. A thermal management engineering computer code that is under development and being managed at Wright Laboratory, Wright-Patterson AFB, is discussed. The primary goal of the code is to aid in the development of a hypersonic vehicle TMS that has been optimized and integrated on a total vehicle basis.

Supersonic Combustion of Hydrogen Jets System in Hypersonic Stream

International Nuclear Information System (INIS)

Zhapbasbaev, U.K.; Makashev, E.P.

2003-01-01

The data of calculated theoretical investigations of diffusive combustion of plane supersonic hydrogen jets in hypersonic stream received with Navier-Stokes parabola equations closed by one-para metrical (k-l) model of turbulence and multiply staged mechanism of hydrogen oxidation are given. Combustion mechanisms depending on the operating parameters are discussing. The influences of air stream composition and ways off fuel feed to the length of ignition delay and level quantity of hydrogen bum-out have been defined. The calculated theoretical results of investigations permit to make the next conclusions: 1. The diffusive combustion of the system of plane supersonic hydrogen jets in hypersonic flow happens in the cellular structures with alternation zones of intensive running of chemical reactions with their inhibition zones. 2. Gas dynamic and heat Mach waves cause a large - scale viscous formation intensifying mixing of fuel with oxidizer. 3. The system ignition of plane supersonic hydrogen jets in hypersonic airy co-flow happens with the formation of normal flame front of hydrogen airy mixture with transition to the diffusive combustion. 4. The presence of active particles in the flow composition initiates the ignition of hydrogen - airy mixture, provides the intensive running of chemical reactions and shortens the length of ignition delay. 5. The supersonic combustion of hydrogel-airy mixture is characterized by two zones: the intensive chemical reactions with an active energy heat release is occurring in the first zone and in the second - a slow hydrogen combustion limited by the mixing of fuel with oxidizer. (author)

Improved Hypersonic Inlet Performance Using Validated Strut Compression Designs

Science.gov (United States)

Bulman, M. J.; Stout, P. W.; Fernandez, R.

1997-01-01

Aerojet is currently executing two Strutjet propulsion contracts: one a Rocket Based Combined Cycle (RBCC) engine for a NASA-Marshall Space Flight Center (MSFC) Advanced Reusable Transportation Technology (ARTT) program, the second a Dual Mode Ram/Scramjet engine for a USAF Wright Laboratories Storable Fuel Scramjet Flow Path Concepts program. The engines employed in both programs operate at supersonic and low hypersonic speeds and use inlets employing forebody external and sidewall compression. Aerojet has developed and validated a successful design methodology applicable to these inlet types. Design features include an integrated vehicle forebody, external side compression struts, strut sidewall and throat bleed, a throat shock trap, and variable geometry internal contraction. Computation Fluid Dynamic (CFD) predictions and test data show these inlets allow substantially increased flow turning angles over other designs. These increased flow turning angles allow shorter and lighter engines than current designs, which in turn enables higher performing vehicles with broad operating characteristics. This paper describes the designs of two different inlets evaluated by the NASA-MSFC and USAF programs, discusses the results of wind tunnel tests performed by NASA-Lewis Research Center, and provides correlations of test data with CFD predictions. Parameters of interest include low Mach number starting capability, start sensitivity as a function of back pressure at various contraction ratios, flow turning angles, strut and throat bleed effects, and pressure recovery at various Mach numbers.

Multiscale Computational Analysis of Nitrogen and Oxygen Gas-Phase Thermochemistry in Hypersonic Flows

Science.gov (United States)

Bender, Jason D.

Understanding hypersonic aerodynamics is important for the design of next-generation aerospace vehicles for space exploration, national security, and other applications. Ground-level experimental studies of hypersonic flows are difficult and expensive; thus, computational science plays a crucial role in this field. Computational fluid dynamics (CFD) simulations of extremely high-speed flows require models of chemical and thermal nonequilibrium processes, such as dissociation of diatomic molecules and vibrational energy relaxation. Current models are outdated and inadequate for advanced applications. We describe a multiscale computational study of gas-phase thermochemical processes in hypersonic flows, starting at the atomic scale and building systematically up to the continuum scale. The project was part of a larger effort centered on collaborations between aerospace scientists and computational chemists. We discuss the construction of potential energy surfaces for the N4, N2O2, and O4 systems, focusing especially on the multi-dimensional fitting problem. A new local fitting method named L-IMLS-G2 is presented and compared with a global fitting method. Then, we describe the theory of the quasiclassical trajectory (QCT) approach for modeling molecular collisions. We explain how we implemented the approach in a new parallel code for high-performance computing platforms. Results from billions of QCT simulations of high-energy N2 + N2, N2 + N, and N2 + O2 collisions are reported and analyzed. Reaction rate constants are calculated and sets of reactive trajectories are characterized at both thermal equilibrium and nonequilibrium conditions. The data shed light on fundamental mechanisms of dissociation and exchange reactions -- and their coupling to internal energy transfer processes -- in thermal environments typical of hypersonic flows. We discuss how the outcomes of this investigation and other related studies lay a rigorous foundation for new macroscopic models for

Ares I Flight Control System Design

Science.gov (United States)

Jang, Jiann-Woei; Alaniz, Abran; Hall, Robert; Bedrossian, Nazareth; Hall, Charles; Ryan, Stephen; Jackson, Mark

2010-01-01

The Ares I launch vehicle represents a challenging flex-body structural environment for flight control system design. This paper presents a design methodology for employing numerical optimization to develop the Ares I flight control system. The design objectives include attitude tracking accuracy and robust stability with respect to rigid body dynamics, propellant slosh, and flex. Under the assumption that the Ares I time-varying dynamics and control system can be frozen over a short period of time, the flight controllers are designed to stabilize all selected frozen-time launch control systems in the presence of parametric uncertainty. Flex filters in the flight control system are designed to minimize the flex components in the error signals before they are sent to the attitude controller. To ensure adequate response to guidance command, step response specifications are introduced as constraints in the optimization problem. Imposing these constraints minimizes performance degradation caused by the addition of the flex filters. The first stage bending filter design achieves stability by adding lag to the first structural frequency to phase stabilize the first flex mode while gain stabilizing the higher modes. The upper stage bending filter design gain stabilizes all the flex bending modes. The flight control system designs provided here have been demonstrated to provide stable first and second stage control systems in both Draper Ares Stability Analysis Tool (ASAT) and the MSFC 6DOF nonlinear time domain simulation.

Stability of hypersonic boundary-layer flows with chemistry

Science.gov (United States)

Reed, Helen L.; Stuckert, Gregory K.; Haynes, Timothy S.

1993-01-01

The effects of nonequilibrium chemistry and three dimensionality on the stability characteristics of hypersonic flows are discussed. In two-dimensional (2-D) and axisymmetric flows, the inclusion of chemistry causes a shift of the second mode of Mack to lower frequencies. This is found to be due to the increase in size of the region of relative supersonic flow because of the lower speeds of sound in the relatively cooler boundary layers. Although this shift in frequency is present in both the equilibrium and nonequilibrium air results, the equilibrium approximation predicts modes which are not observed in the nonequilibrium calculations (for the flight conditions considered). These modes are superpositions of incoming and outgoing unstable disturbances which travel supersonically relative to the boundary-layer edge velocity. Such solutions are possible because of the finite shock stand-off distance. Their corresponding wall-normal profiles exhibit an oscillatory behavior in the inviscid region between the boundary-layer edge and the bow shock. For the examination of three-dimensional (3-D) effects, a rotating cone is used as a model of a swept wing. An increase of stagnation temperature is found to be only slightly stabilizing. The correlation of transition location (N = 9) with parameters describing the crossflow profile is discussed. Transition location does not correlate with the traditional crossflow Reynolds number. A new parameter that appears to correlate for boundary-layer flow was found. A verification with experiments on a yawed cone is provided.

Advances in LO2 Propellant Conditioning

Science.gov (United States)

Mehta, Gopal; Orth, Michael; Stone, William; Perry, Gretchen; Holt, Kimberly; Suter, John

1994-01-01

This paper describes the cryogenic testing and analysis that has recently been completed as part of a multi-year effort to develop a new, more robust and operable LO2 propellant conditioning system. Phase 1 of the program consisted of feasibility demonstrations ot four novel propellant conditioning concepts. A no-bleed, passive propellant conditioning option was shown for the first time to successfully provide desired propellant inlet conditions. The benefits of passive conditioning are reduced operations costs, decreased hardware costs, enhanced operability and increased reliability on future expendable launch vehicles In Phase 2 of the test program, effects of major design parameters were studied and design correlation for future vehicle design were developed. Simultaneously, analytical models were developed and validated. Over 100 tests were conducted with a full-scale feedline using LN2 as the test fluid. A circulation pump provided a range of pressure and flow conditions. The test results showed that the passive propellant conditioning system is insensitive to variations in many of the parameters. The test program provides the validation necessary to incorporate the passive conditioning system into the baseline of future vehicles. Modeling of these systems using computational fluid dynamics seems highly promising.

Cryogenic Propellant Storage and Transfer (CPST) Technology Maturation: Establishing a Foundation for a Technology Demonstration Mission (TDM)

Science.gov (United States)

Doherty, Michael P.; Meyer, Michael L.; Motil, Susan M.; Ginty, Carol A.

2014-01-01

As part of U.S. National Space Policy, NASA is seeking an innovative path for human space exploration, which strengthens the capability to extend human and robotic presence throughout the solar system. NASA is laying the groundwork to enable humans to safely reach multiple potential destinations, including asteroids, Lagrange points, the Moon and Mars. In support of this, NASA is embarking on the Technology Demonstration Mission Cryogenic Propellant Storage and Transfer (TDM CPST) Project to test and validate key cryogenic capabilities and technologies required for future exploration elements, opening up the architecture for large cryogenic propulsion stages (CPS) and propellant depots. The TDM CPST project will provide an on-orbit demonstration of the capability to store, transfer, and measure cryogenic propellants for a duration which is relevant to enable long term human space exploration missions beyond low Earth orbit (LEO). Recognizing that key cryogenic fluid management technologies anticipated for on-orbit (flight) demonstration needed to be matured to a readiness level appropriate for infusion into the design of the flight demonstration, the NASA Headquarters Space Technology Mission Directorate authorized funding for a one-year (FY12) ground based technology maturation program. The strategy, proposed by the CPST Project Manager, focused on maturation through modeling, studies, and ground tests of the storage and fluid transfer Cryogenic Fluid Management (CFM) technology sub-elements and components that were not already at a Technology Readiness Level (TRL) of 5. A technology maturation plan (TMP) was subsequently approved which described: the CFM technologies selected for maturation, the ground testing approach to be used, quantified success criteria of the technologies, hardware and data deliverables, and a deliverable to provide an assessment of the technology readiness after completion of the test, study or modeling activity. This paper will present

Radiation shielding estimates for manned Mars space flight

International Nuclear Information System (INIS)

Dudkin, V.E.; Kovalev, E.E.; Kolomensky, A.V.; Sakovich, V.A.; Semenov, V.F.; Demin, V.P.; Benton, E.V.

1992-01-01

In the analysis of the required radiation shielding for spacecraft during a Mars flight, the specific effects of solar activity (SA) on the intensity of galactic and solar cosmic rays were taken into consideration. Three spaceflight periods were considered: (1) maximum SA; (2) minimum SA; and (3) intermediate SA, when intensities of both galactic and solar cosmic rays are moderately high. Scenarios of spaceflights utilizing liquid-propellant rocket engines, low-and intermediate-thrust nuclear electrojet engines, and nuclear rocket engines, all of which have been designed in the Soviet Union, are reviewed. Calculations were performed on the basis of a set of standards for radiation protection approved by the U.S.S.R. State Committee for Standards. It was found that the lowest estimated mass of a Mars spacecraft, including the radiation shielding mass, obtained using a combination of a liquid propellant engine with low and intermediate thrust nuclear electrojet engines, would be 500-550 metric tons. (author)

Design and simulation on the morphing composite propeller (Conference Presentation)

Science.gov (United States)

Chen, Fanlong; Li, Qinyu; Liu, Liwu; Lan, Xin; Liu, Yanju; Leng, Jinsong

2017-04-01

As one of the most crucial part of the unmanned underwater vehicle (UUV), the composite propeller plays an important role on the UUV's performance. As the composite propeller behaves excellent properties in hydroelastic facet and acoustic suppression, it attracts increasing attentions all over the globe. This paper goes a step further based on this idea, and comes up with a novel concept of "morphing composite propeller" (MCP) to improve the performance of the conventional composite propeller (CCP) to anticipate the improved propeller can perform better to propel the UUV. Based on the new concept, a novel MCP is designed. Each blade of the propeller is assembled with an active rotatable flap (ARF) to change the blade's local camber with flap rotation. Then the transmission mechanism (TM) has been designed and housed in the propeller blade to push the ARF. With the ARF rotating, the UUV can be propelled by different thrusts under certain rotation velocities of the propeller. Based on the design, the Fluent is exploited to analyze the fluid dynamics around the propeller. Finally, based on the design and hydrodynamic analysis, the structural response for the novel morphing composite propeller is calculated. The propeller blade is simplified and layered with composite materials. And the structure response of an MCP is obtained with various rotation angle under the hydrodynamic pressure. This simulation can instruct the design and fabrication techniques of the MCP.

Solution-Space Screening of a Hypersonic Endurance Demonstrator

Science.gov (United States)

Chudoba, Bernd; Coleman, Gary; Oza, Amit; Gonzalez, Lex; Czysz, Paul

2012-01-01

This report documents a parametric sizing study performed to develop a program strategy for research and development and procurement of a feasible next-generation hypersonic air-breathing endurance demonstrator. Overall project focus has been on complementing technical and managerial decision-making during the earliest conceptual design phase towards minimization of operational, technical, and managerial risks.

Wind Tunnel Measurements of Shuttle Orbiter Global Heating with Comparisons to Flight

Science.gov (United States)

Berry, Scott A.; Merski, N. Ronald; Blanchard, Robert C.

2002-01-01

An aerothermodynamic database of global heating images was acquired of the Shuttle Orbiter in the NASA Langley Research Center 20-Inch Mach 6 Air Tunnel. These results were obtained for comparison to the global infrared images of the Orbiter in flight from the infrared sensing aeroheating flight experiment (ISAFE). The most recent ISAFE results from STS-103, consisted of port side images, at hypersonic conditions, of the surface features that result from the strake vortex scrubbing along the side of the vehicle. The wind tunnel results were obtained with the phosphor thermography system, which also provides global information and thus is ideally suited for comparison to the global flight results. The aerothermodynamic database includes both windward and port side heating images of the Orbiter for a range of angles of attack (20 to 40 deg), freestream unit Reynolds number (1 x 10(exp 6))/ft to 8 x 10(exp 6)/ft, body flap deflections (0, 5, and 10 deg), speed brake deflections (0 and 45 deg), as well as with boundary layer trips for forced transition to turbulence heating results. Sample global wind tunnel heat transfer images were extrapolated to flight conditions for comparison to Orbiter flight data. A windward laminar case for an angle of attack of 40 deg was extrapolated to Mach 11.6 flight conditions for comparison to STS-2 flight thermocouple results. A portside wind tunnel image for an angle of attack of 25 deg was extrapolated for Mach 5 flight conditions for comparison to STS-103 global surface temperatures. The comparisons showed excellent qualitative agreement, however the extrapolated wind tunnel results over-predicted the flight surface temperatures on the order of 5% on the windward surface and slightly higher on the portside.

Wind-Tunnel Balance Characterization for Hypersonic Research Applications

Science.gov (United States)

Lynn, Keith C.; Commo, Sean A.; Parker, Peter A.

2012-01-01

Wind-tunnel research was recently conducted at the NASA Langley Research Center s 31-Inch Mach 10 Hypersonic Facility in support of the Mars Science Laboratory s aerodynamic program. Researchers were interested in understanding the interaction between the freestream flow and the reaction control system onboard the entry vehicle. A five-component balance, designed for hypersonic testing with pressurized flow-through capability, was used. In addition to the aerodynamic forces, the balance was exposed to both thermal gradients and varying internal cavity pressures. Historically, the effect of these environmental conditions on the response of the balance have not been fully characterized due to the limitations in the calibration facilities. Through statistical design of experiments, thermal and pressure effects were strategically and efficiently integrated into the calibration of the balance. As a result of this new approach, researchers were able to use the balance continuously throughout the wide range of temperatures and pressures and obtain real-time results. Although this work focused on a specific application, the methodology shown can be applied more generally to any force measurement system calibration.

Numerical analysis of hypersonic turbulent film cooling flows

Science.gov (United States)

Chen, Y. S.; Chen, C. P.; Wei, H.

1992-01-01

As a building block, numerical capabilities for predicting heat flux and turbulent flowfields of hypersonic vehicles require extensive model validations. Computational procedures for calculating turbulent flows and heat fluxes for supersonic film cooling with parallel slot injections are described in this study. Two injectant mass flow rates with matched and unmatched pressure conditions using the database of Holden et al. (1990) are considered. To avoid uncertainties associated with the boundary conditions in testing turbulence models, detailed three-dimensional flowfields of the injection nozzle were calculated. Two computational fluid dynamics codes, GASP and FDNS, with the algebraic Baldwin-Lomax and k-epsilon models with compressibility corrections were used. It was found that the B-L model which resolves near-wall viscous sublayer is very sensitive to the inlet boundary conditions at the nozzle exit face. The k-epsilon models with improved wall functions are less sensitive to the inlet boundary conditions. The testings show that compressibility corrections are necessary for the k-epsilon model to realistically predict the heat fluxes of the hypersonic film cooling problems.

Robust stabilization control based on guardian maps theory for a longitudinal model of hypersonic vehicle.

Science.gov (United States)

Liu, Yanbin; Liu, Mengying; Sun, Peihua

2014-01-01

A typical model of hypersonic vehicle has the complicated dynamics such as the unstable states, the nonminimum phases, and the strong coupling input-output relations. As a result, designing a robust stabilization controller is essential to implement the anticipated tasks. This paper presents a robust stabilization controller based on the guardian maps theory for hypersonic vehicle. First, the guardian maps theories are provided to explain the constraint relations between the open subsets of complex plane and the eigenvalues of the state matrix of closed-loop control system. Then, a general control structure in relation to the guardian maps theories is proposed to achieve the respected design demands. Furthermore, the robust stabilization control law depending on the given general control structure is designed for the longitudinal model of hypersonic vehicle. Finally, a simulation example is provided to verify the effectiveness of the proposed methods.

Thermal Vacuum Test Correlation of A Zero Propellant Load Case Thermal Capacitance Propellant Gauging Analytics Model

Science.gov (United States)

McKim, Stephen A.

2016-01-01

This thesis describes the development and test data validation of the thermal model that is the foundation of a thermal capacitance spacecraft propellant load estimator. Specific details of creating the thermal model for the diaphragm propellant tank used on NASA's Magnetospheric Multiscale spacecraft using ANSYS and the correlation process implemented to validate the model are presented. The thermal model was correlated to within plus or minus 3 degrees Centigrade of the thermal vacuum test data, and was found to be relatively insensitive to uncertainties in applied heat flux and mass knowledge of the tank. More work is needed, however, to refine the thermal model to further improve temperature predictions in the upper hemisphere of the propellant tank. Temperatures predictions in this portion were found to be 2-2.5 degrees Centigrade lower than the test data. A road map to apply the model to predict propellant loads on the actual MMS spacecraft toward its end of life in 2017-2018 is also presented.

Preliminary Studies on Aerodynamic Control with Direct Current Discharge at Hypersonic Speed

Science.gov (United States)

Watanabe, Yasumasa; Takama, Yoshiki; Imamura, Osamu; Watanuki, Tadaharu; Suzuki, Kojiro

A new idea of an aerodynamic control device for hypersonic vehicles using plasma discharges is presented. The effect of DC plasma discharge on a hypersonic flow is examined with both experiments and CFD analyses. It is revealed that the surface pressure upstream of plasma area significantly increases, which would be preferable in realizing a new aerodynamic control devices. Such pressure rise is also observed in the result of analyses of the Navier-Stokes equations with energy addition that simulates the Joule heating of a plasma discharge. It is revealed that the pressure rise due to the existence of the plasma discharge can be qualitatively explained as an effect of Joule heating.

Combustion characteristics of SMX and SMX based propellants

Science.gov (United States)

Reese, David A.

This work investigates the combustion of the new solid nitrate ester 2,3-hydroxymethyl-2,3-dinitro-1,4-butanediol tetranitrate (SMX, C6H 8N6O16). SMX was synthesized for the first time in 2008. It has a melting point of 85 °C and oxygen balance of 0% to CO 2, allowing it to be used as an energetic additive or oxidizer in solid propellants. In addition to its neat combustion characteristics, this work also explores the use of SMX as a potential replacement for nitroglycerin (NG) in double base gun propellants and as a replacement for ammonium perchlorate in composite rocket propellants. The physical properties, sensitivity characteristics, and combustion behaviors of neat SMX were investigated. Its combustion is stable at pressures of up to at least 27.5 MPa (n = 0.81). The observed flame structure is nearly identical to that of other double base propellant ingredients, with a primary flame attached at the surface, a thick isothermal dark zone, and a luminous secondary flame wherein final recombination reactions occur. As a result, the burning rate and primary flame structure can be modeled using existing one-dimensional steady state techniques. A zero gas-phase activation energy approximation results in a good fit between modeled and observed behavior. Additionally, SMX was considered as a replacement for nitroglycerin in a double base propellant. Thermochemical calculations indicate improved performance when compared with the common double base propellant JA2 at SMX loadings above 40 wt-%. Also, since SMX is a room temperature solid, migration may be avoided. Like other nitrate esters, SMX is susceptible to decomposition over long-term storage due to the presence of excess acid in the crystals; the addition of stabilizers (e.g., derivatives of urea) during synthesis should be sufficient to prevent this. the addition of Both unplasticized and plasticized propellants were formulated. Thermal analysis of unplasticized propellant showed a distinct melt

Study of the coupling between real gas effects and rarefied effects on hypersonic aerodynamics

Science.gov (United States)

Chen, Song; Hu, Yuan; Sun, Quanhua

2012-11-01

Hypersonic vehicles travel across the atmosphere at very high speed, and the surrounding gas experiences complicated physical and chemical processes. These processes produce real gas effects at high temperature and rarefied gas effects at high altitude where the two effects are coupled through molecular collisions. In this study, we aim to identify the individual real gas and rarefied gas effects by simulating hypersonic flow over a 2D cylinder, a sphere and a blunted cone using a continuum-based CFD approach and the direct simulation Monte Carlo method. It is found that physical processes such as vibrational excitation and chemical reaction will reduce significantly the shock stand-off distance and flow temperature for flows having small Knudsen number. The calculated skin friction and surface heat flux will decrease when the real gas effects are considered in simulations. The trend, however, gets weakened as the Knudsen number increases. It is concluded that the rarefied gas effects weaken the real gas effects on hypersonic flows.

Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Technology Development Overview

Science.gov (United States)

Hughes, Stephen J.; Cheatwood, F. McNeil; Calomino, Anthony M.; Wright, Henry S.

2013-01-01

The successful flight of the Inflatable Reentry Vehicle Experiment (IRVE)-3 has further demonstrated the potential value of Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology. This technology development effort is funded by NASA's Space Technology Mission Directorate (STMD) Game Changing Development Program (GCDP). This paper provides an overview of a multi-year HIAD technology development effort, detailing the projects completed to date and the additional testing planned for the future. The effort was divided into three areas: Flexible Systems Development (FSD), Mission Advanced Entry Concepts (AEC), and Flight Validation. FSD consists of a Flexible Thermal Protection Systems (FTPS) element, which is investigating high temperature materials, coatings, and additives for use in the bladder, insulator, and heat shield layers; and an Inflatable Structures (IS) element which includes manufacture and testing (laboratory and wind tunnel) of inflatable structures and their associated structural elements. AEC consists of the Mission Applications element developing concepts (including payload interfaces) for missions at multiple destinations for the purpose of demonstrating the benefits and need for the HIAD technology as well as the Next Generation Subsystems element. Ground test development has been pursued in parallel with the Flight Validation IRVE-3 flight test. A larger scale (6m diameter) HIAD inflatable structure was constructed and aerodynamically tested in the National Full-scale Aerodynamics Complex (NFAC) 40ft by 80ft test section along with a duplicate of the IRVE-3 3m article. Both the 6m and 3m articles were tested with instrumented aerodynamic covers which incorporated an array of pressure taps to capture surface pressure distribution to validate Computational Fluid Dynamics (CFD) model predictions of surface pressure distribution. The 3m article also had a duplicate IRVE-3 Thermal Protection System (TPS) to test in addition to testing with the

All-theoretical prediction of cabin noise due to impingement of propeller vortices on a wing structure

Science.gov (United States)

Martinez, R.; Cole, J. E., III; Martini, K.; Westagard, A.

1987-01-01

Reported calculations of structure-borne cabin noise for a small twin engine aircraft powered by tractor propellers rely on the following three-stage methodological breakup of the problem: (1) the unsteady-aerodynamic prediction of wing lift harmonics caused by the whipping action of the vortex system trailed from each propeller; (2) the associated wing/fuselage structural response; (3) the cabin noise field for the computed wall vibration. The first part--the estimate of airloads--skirts a full-fledged aeroelastic situation by assuming the wing to be fixed in space while cancelling the downwash field of the cutting vortices. The model is based on an approximate high-frequency lifting-surface theory justified by the blade rate and flight Mach number of application. Its results drive a finite-element representation of the wing accounting for upper and lower skin surfaces, spars, ribs, and the presence of fuel. The fuselage, modeled as a frame-stiffened cylindrical shell, is bolted to the wing.

Quantifying Non-Equilibrium in Hypersonic Flows Using Entropy Generation

Science.gov (United States)

2007-03-01

do this, two experimental cases performed at the Calspan- University of Buffalo Research Center ( CUBRC ) were modeled using Navier-Stokes based CFD...data provided by the CUBRC hypersonic wind tunnel facility (Holden and Wadhams, 2004). The wall data in Figure 9 and Figure 10 reveals some difference

Advanced Metal Rubber Sensors for Hypersonic Decelerator Entry Systems, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — NanoSonic proposes to design and develop light-weight, low-modulus, and durable Metal Rubber™ sensors for aeroelastic analysis of Hypersonic Decelerator Entry...

Development of a Marine Propeller With Nonplanar Lifting Surfaces

DEFF Research Database (Denmark)

Andersen, Poul; Friesch, Jürgen; Kappel, Jens J.

2005-01-01

The principle of non-planar lifting surfaces is applied to the design of modern aircraft wings to obtain better lift to drag ratios. Whereas a pronounced fin or winglet at the wingtip has been developed for aircraft, the application of the nonplanar principle to marine propellers, dealt...... with in this paper, has led to the KAPPEL propeller with blades curved toward the suction side integrating the fin or winglet into the propeller blade. The combined theoretical, experimental, and practical approach to develop and design marine propellers with nonplanar lifting surfaces has resulted in propellers...

Galerkin CFD solvers for use in a multi-disciplinary suite for modeling advanced flight vehicles

Science.gov (United States)

Moffitt, Nicholas J.

This work extends existing Galerkin CFD solvers for use in a multi-disciplinary suite. The suite is proposed as a means of modeling advanced flight vehicles, which exhibit strong coupling between aerodynamics, structural dynamics, controls, rigid body motion, propulsion, and heat transfer. Such applications include aeroelastics, aeroacoustics, stability and control, and other highly coupled applications. The suite uses NASA STARS for modeling structural dynamics and heat transfer. Aerodynamics, propulsion, and rigid body dynamics are modeled in one of the five CFD solvers below. Euler2D and Euler3D are Galerkin CFD solvers created at OSU by Cowan (2003). These solvers are capable of modeling compressible inviscid aerodynamics with modal elastics and rigid body motion. This work reorganized these solvers to improve efficiency during editing and at run time. Simple and efficient propulsion models were added, including rocket, turbojet, and scramjet engines. Viscous terms were added to the previous solvers to create NS2D and NS3D. The viscous contributions were demonstrated in the inertial and non-inertial frames. Variable viscosity (Sutherland's equation) and heat transfer boundary conditions were added to both solvers but not verified in this work. Two turbulence models were implemented in NS2D and NS3D: Spalart-Allmarus (SA) model of Deck, et al. (2002) and Menter's SST model (1994). A rotation correction term (Shur, et al., 2000) was added to the production of turbulence. Local time stepping and artificial dissipation were adapted to each model. CFDsol is a Taylor-Galerkin solver with an SA turbulence model. This work improved the time accuracy, far field stability, viscous terms, Sutherland?s equation, and SA model with NS3D as a guideline and added the propulsion models from Euler3D to CFDsol. Simple geometries were demonstrated to utilize current meshing and processing capabilities. Air-breathing hypersonic flight vehicles (AHFVs) represent the ultimate

A thin-shock-layer solution for nonequilibrium, inviscid hypersonic flows in earth, Martian, and Venusian atmospheres

Science.gov (United States)

Grose, W. L.

1971-01-01

An approximate inverse solution is presented for the nonequilibrium flow in the inviscid shock layer about a vehicle in hypersonic flight. The method is based upon a thin-shock-layer approximation and has the advantage of being applicable to both subsonic and supersonic regions of the shock layer. The relative simplicity of the method makes it ideally suited for programming on a digital computer with a significant reduction in storage capacity and computing time required by other more exact methods. Comparison of nonequilibrium solutions for an air mixture obtained by the present method is made with solutions obtained by two other methods. Additional cases are presented for entry of spherical nose cones into representative Venusian and Martian atmospheres. A digital computer program written in FORTRAN language is presented that permits an arbitrary gas mixture to be employed in the solution. The effects of vibration, dissociation, recombination, electronic excitation, and ionization are included in the program.

Changing of ballistic parameters from aged gun propellants

NARCIS (Netherlands)

Klerk, W.P.C. de; Driel, C.A. van

2003-01-01

The various properties of an SB and a DB gun propellant were investigated before and after artificial ageing. It was found that the decrease of nitrocellulose (NC) molecular weight, due to ageing of gun propellants, leads to a decrease of the mechanical integrity of the propellant grains. The effect

Drag Reduction by Laser-Plasma Energy Addition in Hypersonic Flow

International Nuclear Information System (INIS)

Oliveira, A. C.; Minucci, M. A. S.; Toro, P. G. P.; Chanes, J. B. Jr; Myrabo, L. N.

2008-01-01

An experimental study was conducted to investigate the drag reduction by laser-plasma energy addition in a low density Mach 7 hypersonic flow. The experiments were conducted in a shock tunnel and the optical beam of a high power pulsed CO 2 TEA laser operating with 7 J of energy and 30 MW peak power was focused to generate the plasma upstream of a hemispherical model installed in the tunnel test section. The non-intrusive schlieren optical technique was used to visualize the effects of the energy addition to hypersonic flow, from the plasma generation until the mitigation of the shock wave profile over the model surface. Aside the optical technique, a piezoelectric pressure transducer was used to measure the impact pressure at stagnation point of the hemispherical model and the pressure reduction could be observed

Laser-propelled ram accelerator

Energy Technology Data Exchange (ETDEWEB)

Sasoh, A. [Tohoku Univ., Sendai (Japan). Inst. of Fluid Science

2000-11-01

The concept of 'laser-propelled ram accelerator (L-RAMAC)' is proposed. Theoretically it is capable of achieving a higher launch speed than that by a chemical ram accelerator because a higher specific energy can be input to the propellant gas. The laser beam is supplied through the muzzle, focused as an annulus behind the base of the projectile. The performance of L-RAMAC is analized based on generalized Rankine-Hugoniot relations, suggesting that a superorbital muzzle speed is achievable out of this device. (orig.)

Propeller installation effects on turboprop aircraft acoustics

Science.gov (United States)

Chirico, Giulia; Barakos, George N.; Bown, Nicholas

2018-06-01

Propeller installation options for a twin-engined turboprop aircraft are evaluated at cruise conditions, aiming to identify the quieter configuration. Computational fluid dynamics is used to investigate the near-field acoustics and transfer functions are employed to estimate the interior cabin noise. Co-rotating and counter-rotating installation options are compared. The effect of propeller synchrophasing is also considered. The employed method captures the complexity of the acoustic field generated by the interactions of the propeller sound fields among each other and with the airframe, showing also the importance of simulating the whole problem to predict the actual noise on a flying aircraft. Marked differences among the various layouts are observed. The counter-rotating top-in option appears the best in terms of acoustics, the top-out propeller rotation leading to louder noise because of inflow conditions and the occurrence of constructive acoustic interferences. Synchrophasing is shown to be beneficial for co-rotating propellers, specially regarding the interior noise, because of favorable effects in the interaction between the propeller direct sound field and the noise due to the airframe. An angle closer to the maximum relative blade shift was found to be the best choice, yielding, however, higher sound levels than those provided by the counter-rotating top-in layout.

THE PROPELLER AND THE FROG

International Nuclear Information System (INIS)

Pan, Margaret; Chiang, Eugene

2010-01-01

'Propellers' in planetary rings are disturbances in ring material excited by moonlets that open only partial gaps. We describe a new type of co-orbital resonance that can explain the observed non-Keplerian motions of propellers. The resonance is between the moonlet underlying the propeller and co-orbiting ring particles downstream of the moonlet where the gap closes. The moonlet librates within the gap about an equilibrium point established by co-orbiting material and stabilized by the Coriolis force. In the limit of small libration amplitude, the libration period scales linearly with the gap azimuthal width and inversely as the square root of the co-orbital mass. The new resonance recalls but is distinct from conventional horseshoe and tadpole orbits; we call it the 'frog' resonance, after the relevant term in equine hoof anatomy. For a ring surface density and gap geometry appropriate for the propeller Bleriot in Saturn's A ring, our theory predicts a libration period of ∼4 years, similar to the ∼3.7 year period over which Bleriot's orbital longitude is observed to vary. These librations should be subtracted from the longitude data before any inferences about moonlet migration are made.

Energy coefficients for a propeller series

DEFF Research Database (Denmark)

Olsen, Anders Smærup

2004-01-01

The efficiency for a propeller is calculated by energy coefficients. These coefficients are related to four types of losses, i.e. the axial, the rotational, the frictional, and the finite blade number loss, and one gain, i.e. the axial gain. The energy coefficients are derived by use...... of the potential theory with the propeller modelled as an actuator disk. The efficiency based on the energy coefficients is calculated for a propeller series. The results show a good agreement between the efficiency based on the energy coefficients and the efficiency obtained by a vortex-lattice method....

Theseus Waits on Lakebed for First Flight

Science.gov (United States)

1996-01-01

The Theseus prototype remotely-piloted aircraft (RPA) waits on the lakebed before its first test flight from NASA's Dryden Flight Research Center, Edwards, California, on May 24, 1996. The Theseus aircraft, built and operated by Aurora Flight Sciences Corporation, Manassas, Virginia, was a unique aircraft flown at NASA's Dryden Flight Research Center, Edwards, California, under a cooperative agreement between NASA and Aurora. Dryden hosted the Theseus program, providing hangar space and range safety for flight testing. Aurora Flight Sciences was responsible for the actual flight testing, vehicle flight safety, and operation of the aircraft. The Theseus remotely piloted aircraft flew its maiden flight on May 24, 1996, at Dryden. During its sixth flight on November 12, 1996, Theseus experienced an in-flight structural failure that resulted in the loss of the aircraft. As of the beginning of the year 2000, Aurora had not rebuilt the aircraft. Theseus was built for NASA under an innovative, $4.9 million fixed-price contract by Aurora Flight Sciences Corporation and its partners, West Virginia University, Morgantown, West Virginia, and Fairmont State College, Fairmont, West Virginia. The twin-engine, unpiloted vehicle had a 140-foot wingspan, and was constructed largely of composite materials. Powered by two 80-horsepower, turbocharged piston engines that drove twin 9-foot-diameter propellers, Theseus was designed to fly autonomously at high altitudes, with takeoff and landing under the active control of a ground-based pilot in a ground control station 'cockpit.' With the potential ability to carry 700 pounds of science instruments to altitudes above 60,000 feet for durations of greater than 24 hours, Theseus was intended to support research in areas such as stratospheric ozone depletion and the atmospheric effects of future high-speed civil transport aircraft engines. Instruments carried aboard Theseus also would be able to validate satellite-based global environmental

Solid propellant processing factor in rocket motor design

Science.gov (United States)

1971-01-01

The ways are described by which propellant processing is affected by choices made in designing rocket engines. Tradeoff studies, design proof or scaleup studies, and special design features are presented that are required to obtain high product quality, and optimum processing costs. Processing is considered to include the operational steps involved with the lining and preparation of the motor case for the grain; the procurement of propellant raw materials; and propellant mixing, casting or extrusion, curing, machining, and finishing. The design criteria, recommended practices, and propellant formulations are included.

14 CFR 25.907 - Propeller vibration and fatigue.

Science.gov (United States)

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Propeller vibration and fatigue. 25.907... vibration and fatigue. This section does not apply to fixed-pitch wood propellers of conventional design. (a... propeller to show that failure due to fatigue will be avoided throughout the operational life of the...

14 CFR 23.907 - Propeller vibration and fatigue.

Science.gov (United States)

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Propeller vibration and fatigue. 23.907... General § 23.907 Propeller vibration and fatigue. This section does not apply to fixed-pitch wood... evaluation of the propeller to show that failure due to fatigue will be avoided throughout the operational...

CFD simulation on Kappel propeller with a hull wake field

DEFF Research Database (Denmark)

Shin, Keun Woo; Andersen, Poul; Møller Bering, Rasmus

2013-01-01

Marine propellers are designed not for the open-water operation, but for the operation behind a hull due to the inhomogeneous hull wake and thrust deduction. The adaptation for the hull wake is important for the propulsive efficiency and cavitation risk especially on single-screw ships. CFD...... simulations for a propeller with a hull model have showed acceptable agreement with a model test result in the thrust and torque (Larsson et al. 2010). In the current work, a measured hull wake is applied to the simulation instead of modelling a hull, because the hull geometry is mostly not available...... for propeller designers and the computational effort can be reduced by excluding the hull. The CFD simulation of a propeller flow with a hull wake is verified in order to use CFD as a propeller design tool. A Kappel propeller, which is an innovative tip-modified propeller, is handled. Kappel propellers...

Computation of hypersonic axisymmetric flows of equilibrium gas over blunt bodies

International Nuclear Information System (INIS)

Hejranfar, K.; Esfahanian, V.; Moghadam, R.K.

2005-01-01

An appropriate combination of the thin-layer Navier-Stokes (TLNS) and parabolized Navier-Stokes (PNS) solvers is used to accurately and efficiently compute hypersonic flowfields of equilibrium air around blunt-body configurations. The TLNS equations are solved in the nose region to provide the initial data plane needed for the solution of the PNS equations. Then the PNS equations are employed to efficiently compute the flowfield for the afterbody region by using a space marching procedure. Both the TLNS and the PNS equations are numerically solved by using the implicit non-iterative finite-difference algorithm of Beam and Warming. A shock fitting technique is used in both the TLNS and PNS codes to obtain accurate solution in the vicinity of the shock. To validate the results of the developed TLNS code, hypersonic laminar flow over a sphere at Mach number of 11.26 is computed. To demonstrate the accuracy and efficiency of using the present TLNS-PNS methodology, the computations are performed for hypersonic flow over 5 o long slender blunt cone at Mach number of 19.25. The results of these computations are found to be in good agreement with available numerical and experimental data. The effects of real gas on the flowfield characteristics are also studied in both the TLNS and PNS solutions. (author)

EFD and CFD Characterization of a CLT Propeller

Directory of Open Access Journals (Sweden)

Daniele Bertetta

2012-01-01

Full Text Available In the present paper an experimental and numerical analysis of an unconventional CLT propeller is carried out. Two different numerical approaches, a potential panel method and an RANSE solver, are employed. Cavitation tunnel experiments are carried out in order to measure, as usual, thrust, torque, and cavity extension for different propeller working points. Moreover, LDV measurements are performed to have a deep insight into the complex wake behind the propeller and to analyze the dynamics of generated tip vortexes. The numerical/experimental analysis and comparison of results highlight the peculiarities of this kind of propellers, the possibility to increase efficiency and reduce cavitation risk, in order to exploit the design approaches already well proven for conventional propellers also in the case of these unconventional geometries.

Experimental study of high-speed counter-rotation propeller on low speed wind range; Dojiku hantengata kosoku propeller no teisokuiki ni okeru fudo jikken

Energy Technology Data Exchange (ETDEWEB)

NONE

1997-03-01

Collaborative research was conducted by National Aerospace Laboratory and Japan Aircraft Development Company in the period of fiscal 1988-1992 into methods for testing aircraft with advanced propeller in low-speed wind tunnel. The propulsion efficiency of the currently available high-bypass turbofan engine is approximately 60% in the vicinity of Mach number 0.85. Propeller-driven aircraft, whose propulsion efficiency is as high as 80% in the low Mach number domain, are scarcely in practical use in the domain of Mach number 0.75 or higher. There are studies reported abroad as well as in Japan for the propeller-driven aircraft to enjoy higher propeller propulsion efficiency even in the vicinity of Mach number 0.8 by modifying the propeller diameter, number of blades, and blade sections, etc. This paper describes the experimental research into the high-speed counter-rotation propeller. A counter-rotation propeller 0.3m in diameter and provided with coaxially arranged 8times2 SR-2 blades is evaluated for pitch angles during the takeoff and landing modes, for thrust characteristics at the pitch angle for the cruising mode, and for propeller backwash and noise. 15 refs., 72 figs., 9 tabs.

Technology of foamed propellants

Energy Technology Data Exchange (ETDEWEB)

Boehnlein-Mauss, Jutta; Kroeber, Hartmut [Fraunhofer Institut fuer Chemische Technologie ICT, Pfinztal (Germany)

2009-06-15

Foamed propellants are based on crystalline explosives bonded in energetic reaction polymers. Due to their porous structures they are distinguished by high burning rates. Energy content and material characteristics can be varied by using different energetic fillers, energetic polymers and porous structures. Foamed charges can be produced easily by the reaction injection moulding process. For the manufacturing of foamed propellants a semi-continuous remote controlled production plant in pilot scale was set up and a modified reaction injection moulding process was applied. (Abstract Copyright [2009], Wiley Periodicals, Inc.)

Experimental Studies of Shock Interaction Phenomena Associated with Hypersonic Airbreathing Propulsion

National Research Council Canada - National Science Library

Holden, Michael

2001-01-01

... and double cone configurations in hypersonic flow. In the best Navier-Stokes solutions the structure and density of the flowfield was captured exactly over both the hollow cylinder/flare and double cone models...

A Laser-Based Diagnostic Suite for Hypersonic Test Facilities, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — In this SBIR effort, Los Gatos Research (LGR) proposes to develop a suite of laser-based diagnostics for the study of reactive and non-reactive hypersonic flows....

In-flight performance optimization for rotorcraft with redundant controls

Science.gov (United States)

Ozdemir, Gurbuz Taha

A conventional helicopter has limits on performance at high speeds because of the limitations of main rotor, such as compressibility issues on advancing side or stall issues on retreating side. Auxiliary lift and thrust components have been suggested to improve performance of the helicopter substantially by reducing the loading on the main rotor. Such a configuration is called the compound rotorcraft. Rotor speed can also be varied to improve helicopter performance. In addition to improved performance, compound rotorcraft and variable RPM can provide a much larger degree of control redundancy. This additional redundancy gives the opportunity to further enhance performance and handling qualities. A flight control system is designed to perform in-flight optimization of redundant control effectors on a compound rotorcraft in order to minimize power required and extend range. This "Fly to Optimal" (FTO) control law is tested in simulation using the GENHEL model. A model of the UH-60, a compound version of the UH-60A with lifting wing and vectored thrust ducted propeller (VTDP), and a generic compound version of the UH-60A with lifting wing and propeller were developed and tested in simulation. A model following dynamic inversion controller is implemented for inner loop control of roll, pitch, yaw, heave, and rotor RPM. An outer loop controller regulates airspeed and flight path during optimization. A Golden Section search method was used to find optimal rotor RPM on a conventional helicopter, where the single redundant control effector is rotor RPM. The FTO builds off of the Adaptive Performance Optimization (APO) method of Gilyard by performing low frequency sweeps on a redundant control for a fixed wing aircraft. A method based on the APO method was used to optimize trim on a compound rotorcraft with several redundant control effectors. The controller can be used to optimize rotor RPM and compound control effectors through flight test or simulations in order to

Interior near-field acoustical holography in flight.

Science.gov (United States)

Williams, E G; Houston, B H; Herdic, P C; Raveendra, S T; Gardner, B

2000-10-01

In this paper boundary element methods (BEM) are mated with near-field acoustical holography (NAH) in order to determine the normal velocity over a large area of a fuselage of a turboprop airplane from a measurement of the pressure (hologram) on a concentric surface in the interior of the aircraft. This work represents the first time NAH has been applied in situ, in-flight. The normal fuselage velocity was successfully reconstructed at the blade passage frequency (BPF) of the propeller and its first two harmonics. This reconstructed velocity reveals structure-borne and airborne sound-transmission paths from the engine to the interior space.

Real-Gas Correction Factors for Hypersonic Flow Parameters in Helium

Science.gov (United States)

Erickson, Wayne D.

1960-01-01

The real-gas hypersonic flow parameters for helium have been calculated for stagnation temperatures from 0 F to 600 F and stagnation pressures up to 6,000 pounds per square inch absolute. The results of these calculations are presented in the form of simple correction factors which must be applied to the tabulated ideal-gas parameters. It has been shown that the deviations from the ideal-gas law which exist at high pressures may cause a corresponding significant error in the hypersonic flow parameters when calculated as an ideal gas. For example the ratio of the free-stream static to stagnation pressure as calculated from the thermodynamic properties of helium for a stagnation temperature of 80 F and pressure of 4,000 pounds per square inch absolute was found to be approximately 13 percent greater than that determined from the ideal-gas tabulation with a specific heat ratio of 5/3.

Design and Analysis of Morpheus Lander Flight Control System

Science.gov (United States)

Jang, Jiann-Woei; Yang, Lee; Fritz, Mathew; Nguyen, Louis H.; Johnson, Wyatt R.; Hart, Jeremy J.

2014-01-01

The Morpheus Lander is a vertical takeoff and landing test bed vehicle developed to demonstrate the system performance of the Guidance, Navigation and Control (GN&C) system capability for the integrated autonomous landing and hazard avoidance system hardware and software. The Morpheus flight control system design must be robust to various mission profiles. This paper presents a design methodology for employing numerical optimization to develop the Morpheus flight control system. The design objectives include attitude tracking accuracy and robust stability with respect to rigid body dynamics and propellant slosh. Under the assumption that the Morpheus time-varying dynamics and control system can be frozen over a short period of time, the flight controllers are designed to stabilize all selected frozen-time control systems in the presence of parametric uncertainty. Both control gains in the inner attitude control loop and guidance gains in the outer position control loop are designed to maximize the vehicle performance while ensuring robustness. The flight control system designs provided herein have been demonstrated to provide stable control systems in both Draper Ares Stability Analysis Tool (ASAT) and the NASA/JSC Trick-based Morpheus time domain simulation.

Nondestructive examination of the Tropical Rainfall Measuring Mission (TRMM) reaction control subsystem (RCS) propellant tanks

Science.gov (United States)

Free, James M.

1993-01-01

This paper assesses the feasibility of using eddy current nondestructive examination to determine flaw sizes in completely assembled hydrazine propellant tanks. The study was performed by the NASA Goddard Space Flight Center for the Tropical Rainfall Measuring Mission (TRMM) project to help determine whether existing propellant tanks could meet the fracture analysis requirements of the current pressure vessel specification, MIL-STD-1522A and, therefore be used on the TRMM spacecraft. After evaluating several nondestructive test methods, eddy current testing was selected as the most promising method for determining flaw sizes on external and internal surfaces of completely assembled tanks. Tests were conducted to confirm the detection capability of the eddy current NDE, procedures were developed to inspect two candidate tanks, and the test support equipment was designed. The non-spherical tank eddy current NDE test program was terminated when the decision was made to procure new tanks for the TRMM propulsion subsystem. The information on the development phase of this test program is presented in this paper as a reference for future investigation on the subject.

High-Fidelity Kinetics and Radiation Transport for NLTE Hypersonic Flows, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — The modeling of NLTE hypersonic flows combines several disciplines: chemistry, kinetics, radiation transport, fluid mechanics, and surface science. No single code or...

Effective high-order solver with thermally perfect gas model for hypersonic heating prediction

International Nuclear Information System (INIS)

Jiang, Zhenhua; Yan, Chao; Yu, Jian; Qu, Feng; Ma, Libin

2016-01-01

Highlights: • Design proper numerical flux for thermally perfect gas. • Line-implicit LUSGS enhances efficiency without extra memory consumption. • Develop unified framework for both second-order MUSCL and fifth-order WENO. • The designed gas model can be applied to much wider temperature range. - Abstract: Effective high-order solver based on the model of thermally perfect gas has been developed for hypersonic heat transfer computation. The technique of polynomial curve fit coupling to thermodynamics equation is suggested to establish the current model and particular attention has been paid to the design of proper numerical flux for thermally perfect gas. We present procedures that unify five-order WENO (Weighted Essentially Non-Oscillatory) scheme in the existing second-order finite volume framework and a line-implicit method that improves the computational efficiency without increasing memory consumption. A variety of hypersonic viscous flows are performed to examine the capability of the resulted high order thermally perfect gas solver. Numerical results demonstrate its superior performance compared to low-order calorically perfect gas method and indicate its potential application to hypersonic heating predictions for real-life problem.

Investigation on utilization of liquid propellant in ballistic range experiments

Energy Technology Data Exchange (ETDEWEB)

Saso, Akihiro; Oba, Shinji; Takayama, Kazuyoshi [Tohoku University, Sendai (Japan)

1999-10-31

Experiments were conducted in a ballistic range using a HAN (hydroxylammonium nitrate)-based liquid monopropellant, LP1846. In a 25-mm-bore single-stage gun, using bulk-loaded propellant of 10 to 35 g, a muzzle speed up to 1.0 km/s was obtained. Time variations of propellant chamber pressures and in-tube projectile velocity profiles were measured. The liquid propellant combustion was initiated accompanying a delay time which was created due to the pyrolysis of the propellant. In order to obtain reliable ballistic range performance, the method of propellant loading was revealed to be critical. Since the burning rate of the liquid propellant is relatively low, the peak acceleration and the muzzle speed strongly depend on the rupture pressure of a diaphragm that was inserted between the launch tube and the propellant chamber. (author)

Burning properties and mechanical integrity of aged gun propellants

NARCIS (Netherlands)

Driel, C.A. van; Klerk, W.P.C. de

2017-01-01

Conventional gun propellants exhibit the phenomenon of nitrocellulose (NC) decomposition. Besides an effect on thermal stability of propellants, decomposition of NC has an effect on the mechanical integrity of the propellant grains. Enhanced grain fracture may lead to unacceptable changes of the

SiC Matrix Composites for High Temperature Hypersonic Vehicle Applications, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Durable high temperature materials are required for hypersonic engine and structural thermal protection systems. In particular, 2700ºF or greater capable structural...

Numerical Investigation of a Generic Scramjet Configuration

OpenAIRE

Karl, Sebastian

2011-01-01

A Supersonic Combustion Ramjet (scramjet) is, at least in theory, an efficient air-breathing propulsion system for sustained hypersonic flight at Mach numbers above approximately M=5. Important design issues for such hypersonic propulsion systems, are the lack of ground based facilities capable of testing a full-sized engine at cruise flight conditions and the absence of general scaling laws for the extrapolation of wind tunnel data to flight configurations. Therefore, there is a strong need ...

Hypersonic Technology Developments with EU Co-Funded Projects

Science.gov (United States)

2010-09-01

and Hypersonic Systems and Technologies Conference, AIAA-2006-8109, 06-09/11 2006, Canberra, Australia. [18] Karl S., Hannemann K., Steelant J. and...Canberra, Australia. [23] Haidn, O., Ciezki, H., Hannemann , K. and Karl., S., Selected Supersonic Combustion Activities at DLR within the European...LAPCAT Project, 2nd European Conference for Aerospace Sciences (EUCASS), July 2007, Brussels, Belgium. [24] Martinez-Schram J. , Karl S., Hannemann K

Molecular-Based Optical Diagnostics for Hypersonic Nonequilibrium Flows

Science.gov (United States)

Danehy, Paul; Bathel, Brett; Johansen, Craig; Winter, Michael; O'Byrne, Sean; Cutler, Andrew

2015-01-01

This presentation package consists of seven different talks rolled up into one. These talks are all invited orals presentations in a special session at the Aviation 2015 conference and represent contributions that were made to a recent AIAA book that will be published entitled 'Hypersonic Nonequilibrium Flows: Fundamentals and Recent Advances'. Slide 5 lists the individual presentations that will be given during the special session.

Experimental Research on Air Propellers III

Science.gov (United States)

Durand, W F; Lesley, E P

1920-01-01

Report presents the results of wind tunnel tests of propellers that examined the influence of the following characteristics: (1) nominal pitch ratio 1.3 combined with a certain number of the more common or standard forms and proportions; (2) driving face slightly rounded or convex; (3) change in the location of the maximum thickness ordinate of the blade section; (4) pushing forward the leading edge of the blade, thus giving a rounded convex surface on the leading side of the driving face. (5) a series of values for the constant "angle of attack" in forming propellers with radially increasing pitch. In accordance with these purposes tests were carried out on 28 propellers.

Shuttle APS propellant thermal conditioner study

Science.gov (United States)

Pearson, W. E.

1971-01-01

A study program was performed to allow selection of thermal conditioner assemblies for superheating O2 and H2 at supercritical pressures. The application was the auxiliary propulsion system (APS) for the space shuttle vehicle. The O2/H2 APS propellant feed system included propellant conditioners, of which the thermal conditioner assemblies were a part. Cryogens, pumped to pressures above critical, were directed to the thermal conditioner assembly included: (1) a gas generator assembly with ignition system and bipropellant valves, which burned superheated O2 and H2 at rich conditions; (2) a heat exchanger assembly for thermal conditioning of the cryogenic propellant; and (3) a dump nozzle for heat exchanger exhaust.

Heat transfer in heterogeneous propellant combustion systems

International Nuclear Information System (INIS)

Brewster, M.Q.

1992-01-01

This paper reports that heat transfer plays an important role in several critical areas of heterogeneous, solid-propellant combustion systems. These areas include heat feedback to the propellant surface, heat transfer between burning aluminum droplets and their surroundings, heat transfer to internal insulation systems, and heat transfer to aft-end equipment. Gas conduction dominates heat feedback to the propellant surface in conventional ammonium perchlorate (AP) composite propellants, although particle radiative feedback also plays a significant role in combustion of metalized propellants. Particle radiation plays a dominant role in heat transfer to internal insulation, compared with that of convection. However, conduction by impingement of burning aluminum particles, which has not been extensively studied, may also be significant. Radiative heat loss plays an important role in determining the burning rate of molten aluminum particles due to a highly luminous, oxide particle-laden, detached flame envelope. Radiation by aluminum oxide smoke particles also plays a dominant role in heat transfer from the exhaust plume to aft-end equipment. Uncertainties in aluminum oxide particle-size distribution and optical properties still make it difficult to predict radiative plume heat transfer accurately from first principles

A model for supersonic and hypersonic impactors for nanoparticles

International Nuclear Information System (INIS)

Abouali, Omid; Ahmadi, Goodarz

2005-01-01

In this study the performance of supersonic and hypersonic impactors for collection efficiency of nanoparticles (in the size range of 2-100 nm) under various operating conditions is analyzed. Axisymmetric forms of the compressible Navier-Stokes and energy equations are solved and the airflow and thermal condition in the impactor are evaluated. A Lagrangian particle trajectory analysis procedure is used and the deposition rates of different size particles under various operating conditions are studied. For dilute particle concentrations, the assumption of one-way interaction is used and the effect of particles on gas flow field is ignored. The importance of drag, lift and Brownian forces on particle motions in supersonic impactors is discussed. Sensitivity of the simulation results to the use of different assumptions for the Cunningham correction coefficient is studied. It is shown that accurate evaluation of the gas mean free path and the Cunningham correction factor is important for accurate simulation of nano-particle transport and deposition in supersonic/hypersonic impactors. The computer simulation results are compared favorably with the available experimental data

Powered Flight The Engineering of Aerospace Propulsion

CERN Document Server

Greatrix, David R

2012-01-01

Whilst most contemporary books in the aerospace propulsion field are dedicated primarily to gas turbine engines, there is often little or no coverage of other propulsion systems and devices such as propeller and helicopter rotors or detailed attention to rocket engines. By taking a wider viewpoint, Powered Flight - The Engineering of Aerospace Propulsion aims to provide a broader context, allowing observations and comparisons to be made across systems that are overlooked by focusing on a single aspect alone. The physics and history of aerospace propulsion are built on step-by-step, coupled with the development of an appreciation for the mathematics involved in the science and engineering of propulsion. Combining the author’s experience as a researcher, an industry professional and a lecturer in graduate and undergraduate aerospace engineering, Powered Flight - The Engineering of Aerospace Propulsion covers its subject matter both theoretically and with an awareness of the practicalities of the industry. To ...

Simulation and Optimization of Control of Selected Phases of Gyroplane Flight

Directory of Open Access Journals (Sweden)

Wienczyslaw Stalewski

2018-02-01

Full Text Available Optimization methods are increasingly used to solve problems in aeronautical engineering. Typically, optimization methods are utilized in the design of an aircraft airframe or its structure. The presented study is focused on improvement of aircraft flight control procedures through numerical optimization. The optimization problems concern selected phases of flight of a light gyroplane—a rotorcraft using an unpowered rotor in autorotation to develop lift and an engine-powered propeller to provide thrust. An original methodology of computational simulation of rotorcraft flight was developed and implemented. In this approach the aircraft motion equations are solved step-by-step, simultaneously with the solution of the Unsteady Reynolds-Averaged Navier–Stokes equations, which is conducted to assess aerodynamic forces acting on the aircraft. As a numerical optimization method, the BFGS (Broyden–Fletcher–Goldfarb–Shanno algorithm was adapted. The developed methodology was applied to optimize the flight control procedures in selected stages of gyroplane flight in direct proximity to the ground, where proper control of the aircraft is critical to ensure flight safety and performance. The results of conducted computational optimizations proved the qualitative correctness of the developed methodology. The research results can be helpful in the design of easy-to-control gyroplanes and also in the training of pilots for this type of rotorcraft.

Experimental results of a Mach 10 conical-flow derived waverider to 14-X hypersonic aerospace vehicle

Directory of Open Access Journals (Sweden)

Tiago Cavalcanti Rolim

2011-05-01

Full Text Available This paper presents a research in the development of the 14-X hypersonic airspace vehicle at Institute for Advanced Studies (IEAv from Department of Science and Aerospace Technology (DCTA of the Brazilian Air Force (FAB. The 14-X project objective is to develop a higher efficient satellite launch alternative, using a Supersonic Combustion Ramjet (SCRAMJET engine and waverider aerodynamics. For this development, the waverider technology is under investigation in Prof. Henry T. Nagamatsu Aerothermodynamics and Hypersonics Laboratory (LHTN, in IEAv/DCTA. The investigation has been conducted through ground test campaigns in Hypersonic Shock Tunnel T3. The 14-X Waverider Vehicle characteristic was verified in shock tunnel T3 where surface static pressures and pitot pressure for Mach number 10 were measured and, using Schlieren photographs Diagnostic Method, it was possible to identify a leading-edge attached shock wave in 14-X lower surface.

A study of upwind schemes on the laminar hypersonic heating predictions for the reusable space vehicle

Science.gov (United States)

Qu, Feng; Sun, Di; Zuo, Guang

2018-06-01

With the rapid development of the Computational Fluid Dynamics (CFD), Accurate computing hypersonic heating is in a high demand for the design of the new generation reusable space vehicle to conduct deep space exploration. In the past years, most researchers try to solve this problem by concentrating on the choice of the upwind schemes or the definition of the cell Reynolds number. However, the cell Reynolds number dependencies and limiter dependencies of the upwind schemes, which are of great importance to their performances in hypersonic heating computations, are concerned by few people. In this paper, we conduct a systematic study on these properties respectively. Results in our test cases show that SLAU (Simple Low-dissipation AUSM-family) is with a much higher level of accuracy and robustness in hypersonic heating predictions. Also, it performs much better in terms of the limiter dependency and the cell Reynolds number dependency.

Dissociation–recombination models in hypersonic boundary layer O2/O flows

International Nuclear Information System (INIS)

Armenise, I.; Esposito, F.

2012-01-01

Graphical abstract: In hypersonic boundary layers, in which the temperature strongly decreases from the edge to the body surface, the coupling of transport phenomena and chemical kinetics causes a strong vibrational non-equilibrium, as demonstrated by the vibrational distributions and the pseudo-first-order dissociation constants. In this work a pure O2/O mixture has been investigated to evaluate the role of new multiquanta atom-molecule collision rate coefficients, calculated by means of a quasiclassical trajectory (QCT) method. Highlights: ► We evaluate the vibrational non-equilibrium in oxygen hypersonic boundary layer flows. ► We adopt a state-to-state vibrational kinetics model. ► We use updated quasicassical trajectory atom–molecule collision rate coefficients. ► Multiquanta transitions and direct dissociation–recombination are important. ► We calculate the heat flux through the boundary layer. - Abstract: A recent complete set of oxygen atom–molecule collision rate coefficients, calculated by means of a quasiclassical trajectory (QCT) method, has been used to evaluate the vibrational non-equilibrium in hypersonic boundary layer flows. The importance of multiquanta transitions has been demonstrated. Moreover a new ‘direct dissociation–recombination’ (DDR) model has been adopted and the corresponding results differ from the ones obtained with the ladder-climbing (LC) model, characterized by the extrapolation of bound-to-bound transitions to the continuum. The heat flux through the boundary layer and at the surface has been calculated too.

Development of hydrazinium nitroformate based solid propellants

NARCIS (Netherlands)

Schöyer, H.F.R.; Schnorhk, A.J.; Korting, P.A.O.G.; Lit, P.J. van; Mul, J.M.; Gadiot, G.; Meulenbrugge, J.J.

1995-01-01

The development of new high-performance propellant combinations requires the establishment of safety and handling characteristics and thermodynamic decomposition and explosive properties. This paper addresses the early development phases of a new composite solid propellant based on HNF as oxidizer

In-Space Propellant Production Using Water

Science.gov (United States)

Notardonato, William; Johnson, Wesley; Swanger, Adam; McQuade, William

2012-01-01

A new era of space exploration is being planned. Manned exploration architectures under consideration require the long term storage of cryogenic propellants in space, and larger science mission directorate payloads can be delivered using cryogenic propulsion stages. Several architecture studies have shown that in-space cryogenic propulsion depots offer benefits including lower launch costs, smaller launch vehicles, and enhanced mission flexibility. NASA is currently planning a Cryogenic Propellant Storage and Transfer (CPST) technology demonstration mission that will use existing technology to demonstrate long duration storage, acquisition, mass gauging, and transfer of liquid hydrogen in low Earth orbit. This mission will demonstrate key technologies, but the CPST architecture is not designed for optimal mission operations for a true propellant depot. This paper will consider cryogenic propellant depots that are designed for operability. The operability principles considered are reusability, commonality, designing for the unique environment of space, and use of active control systems, both thermal and fluid. After considering these operability principles, a proposed depot architecture will be presented that uses water launch and on orbit electrolysis and liquefaction. This could serve as the first true space factory. Critical technologies needed for this depot architecture, including on orbit electrolysis, zero-g liquefaction and storage, rendezvous and docking, and propellant transfer, will be discussed and a developmental path forward will be presented. Finally, use of the depot to support the NASA Science Mission Directorate exploration goals will be presented.

Observer-based linear parameter varying H∞ tracking control for hypersonic vehicles

Directory of Open Access Journals (Sweden)

Yiqing Huang

2016-11-01

Full Text Available This article aims to develop observer-based linear parameter varying output feedback H∞ tracking controller for hypersonic vehicles. Due to the complexity of an original nonlinear model of the hypersonic vehicle dynamics, a slow–fast loop linear parameter varying polytopic model is introduced for system stability analysis and controller design. Then, a state observer is developed by linear parameter varying technique in order to estimate the unmeasured attitude angular for slow loop system. Also, based on the designed linear parameter varying state observer, a kind of attitude tracking controller is presented to reduce tracking errors for all bounded reference attitude angular inputs. The closed-loop linear parameter varying system is proved to be quadratically stable by Lypapunov function technique. Finally, simulation results show that the developed linear parameter varying H∞ controller has good tracking capability for reference commands.

Numerical Analysis of Ice Impacts on Azimuth Propeller

Science.gov (United States)

2013-09-01

uniformity as one solid object, though the built-on process is recognized as well. The number of blades for this propeller varies to control cavitation ...Figure 7). The CRP allows the hydrodynamic advantage of regaining lost rotational energy from the first propeller, which is the conventional...it speeds flow through and accelerates it past the propellers. Decelerating duct slows the flow and allows for greatly reduced cavitation and

High-Performance, Space-Storable, Bi-Propellant Program Status

Science.gov (United States)

Schneider, Steven J.

2002-01-01

Bipropellant propulsion systems currently represent the largest bus subsystem for many missions. These missions range from low Earth orbit satellite to geosynchronous communications and planetary exploration. The payoff of high performance bipropellant systems is illustrated by the fact that Aerojet Redmond has qualified a commercial NTO/MMH engine based on the high Isp technology recently delivered by this program. They are now qualifying a NTO/hydrazine version of this engine. The advanced rhenium thrust chambers recently provided by this program have raised the performance of earth storable propellants from 315 sec to 328 sec of specific impulse. The recently introduced rhenium technology is the first new technology introduced to satellite propulsion in 30 years. Typically, the lead time required to develop and qualify new chemical thruster technology is not compatible with program development schedules. These technology development programs must be supported by a long term, Base R&T Program, if the technology s to be matured. This technology program then addresses the need for high performance, storable, on-board chemical propulsion for planetary rendezvous and descent/ascent. The primary NASA customer for this technology is Space Science, which identifies this need for such programs as Mars Surface Return, Titan Explorer, Neptune Orbiter, and Europa Lander. High performance (390 sec) chemical propulsion is estimated to add 105% payload to the Mars Sample Return mission or alternatively reduce the launch mass by 33%. In many cases, the use of existing (flight heritage) propellant technology is accommodated by reducing mission objectives and/or increasing enroute travel times sacrificing the science value per unit cost of the program. Therefore, a high performance storable thruster utilizing fluorinated oxidizers with hydrazine is being developed.

Shock stand off Calculations for Hemisphere in Hypersonic Flows

International Nuclear Information System (INIS)

Hanif, M.; Ghaffar, A.; Bilal, S.; Zahir, S.; Khan, M.A.

2004-01-01

The shape and location of shock has been studied by solving the axi symmetric Navier Stokes Equations for a hemisphere in hypersonic flow. The effect of Mach number on shock stand-off distance has been investigated. It is found that the shock location varies with Mach number and the free stream conditions at a given nose radius. (author)

NASA Langley's AirSTAR Testbed: A Subscale Flight Test Capability for Flight Dynamics and Control System Experiments

Science.gov (United States)

Jordan, Thomas L.; Bailey, Roger M.

2008-01-01

As part of the Airborne Subscale Transport Aircraft Research (AirSTAR) project, NASA Langley Research Center (LaRC) has developed a subscaled flying testbed in order to conduct research experiments in support of the goals of NASA s Aviation Safety Program. This research capability consists of three distinct components. The first of these is the research aircraft, of which there are several in the AirSTAR stable. These aircraft range from a dynamically-scaled, twin turbine vehicle to a propeller driven, off-the-shelf airframe. Each of these airframes carves out its own niche in the research test program. All of the airplanes have sophisticated on-board data acquisition and actuation systems, recording, telemetering, processing, and/or receiving data from research control systems. The second piece of the testbed is the ground facilities, which encompass the hardware and software infrastructure necessary to provide comprehensive support services for conducting flight research using the subscale aircraft, including: subsystem development, integrated testing, remote piloting of the subscale aircraft, telemetry processing, experimental flight control law implementation and evaluation, flight simulation, data recording/archiving, and communications. The ground facilities are comprised of two major components: (1) The Base Research Station (BRS), a LaRC laboratory facility for system development, testing and data analysis, and (2) The Mobile Operations Station (MOS), a self-contained, motorized vehicle serving as a mobile research command/operations center, functionally equivalent to the BRS, capable of deployment to remote sites for supporting flight tests. The third piece of the testbed is the test facility itself. Research flights carried out by the AirSTAR team are conducted at NASA Wallops Flight Facility (WFF) on the Eastern Shore of Virginia. The UAV Island runway is a 50 x 1500 paved runway that lies within restricted airspace at Wallops Flight Facility. The

Miniaturized compact water-cooled pitot-pressure probe for flow-field surveys in hypersonic wind tunnels

Science.gov (United States)

Ashby, George C.

1988-01-01

An experimental investigation of the design of pitot probes for flowfield surveys in hypersonic wind tunnels is reported. The results show that a pitot-pressure probe can be miniaturized for minimum interference effects by locating the transducer in the probe support body and water-cooling it so that the pressure-settling time and transducer temperature are compatible with hypersonic tunnel operation and flow conditions. Flowfield surveys around a two-to-one elliptical cone model in a 20-inch Mach 6 wind tunnel using such a probe show that probe interference effects are essentially eliminated.

AN AIRPLANE WITH UNCONVENTIONALLY PLACED PROPELLER POWER UNIT

Directory of Open Access Journals (Sweden)

Jan Červinka

2017-02-01

Full Text Available The significance of the influence of operating propellers on the aircraft aerodynamic characteristics is well-known. Wind tunnel testing of an airplane model with operating propellers is a complex task regarding the required similarity of the full-scale and the model case. Matching sufficient similarity in axial and rotational velocities in the propeller slipstream is the primordial condition for the global aerodynamic similarity of the windtunnel testing. An example of the model power units with related devices is presented. Examples of the wind tunnel testing results illustrate the extent of the propeller influence on aerodynamic characteristics of an aircraft of unconventional configuration with power units positioned at the fuselage afterbody.

Numerical Simulation of Transitional, Hypersonic Flows using a Hybrid Particle-Continuum Method

Science.gov (United States)

Verhoff, Ashley Marie

Analysis of hypersonic flows requires consideration of multiscale phenomena due to the range of flight regimes encountered, from rarefied conditions in the upper atmosphere to fully continuum flow at low altitudes. At transitional Knudsen numbers there are likely to be localized regions of strong thermodynamic nonequilibrium effects that invalidate the continuum assumptions of the Navier-Stokes equations. Accurate simulation of these regions, which include shock waves, boundary and shear layers, and low-density wakes, requires a kinetic theory-based approach where no prior assumptions are made regarding the molecular distribution function. Because of the nature of these types of flows, there is much to be gained in terms of both numerical efficiency and physical accuracy by developing hybrid particle-continuum simulation approaches. The focus of the present research effort is the continued development of the Modular Particle-Continuum (MPC) method, where the Navier-Stokes equations are solved numerically using computational fluid dynamics (CFD) techniques in regions of the flow field where continuum assumptions are valid, and the direct simulation Monte Carlo (DSMC) method is used where strong thermodynamic nonequilibrium effects are present. Numerical solutions of transitional, hypersonic flows are thus obtained with increased physical accuracy relative to CFD alone, and improved numerical efficiency is achieved in comparison to DSMC alone because this more computationally expensive method is restricted to those regions of the flow field where it is necessary to maintain physical accuracy. In this dissertation, a comprehensive assessment of the physical accuracy of the MPC method is performed, leading to the implementation of a non-vacuum supersonic outflow boundary condition in particle domains, and more consistent initialization of DSMC simulator particles along hybrid interfaces. The relative errors between MPC and full DSMC results are greatly reduced as a

On nitrogen condensation in hypersonic nozzle flows: Numerical method and parametric study

KAUST Repository

Lin, Longyuan; Cheng, Wan; Luo, Xisheng; Qin, Fenghua

2013-01-01

A numerical method for calculating two-dimensional planar and axisymmetric hypersonic nozzle flows with nitrogen condensation is developed. The classical nucleation theory with an empirical correction function and the modified Gyarmathy model

Improved-Delayed-Detached-Eddy Simulation of cavity-induced transition in hypersonic boundary layer

International Nuclear Information System (INIS)

Xiao, Lianghua; Xiao, Zhixiang; Duan, Zhiwei; Fu, Song

2015-01-01

Highlights: • This work is about hypersonic cavity-induced transition with IDDES approach. • The length-to-width-to-depth ratio of the cavity is 19.9:3.57:1 at AoA −10° and −15°. • Flow remains laminar at −10°, transition occurs at −15° and cavity changed from open to close type. • Streamwise vortices, impingement shock, traveling shocks and exit shock are observed. • Breakdown of these vortices triggering rapid flow transition. - Abstract: Hypersonic flow transition from laminar to turbulent due to the surface irregularities, like local cavities, can greatly affect the surface heating and skin friction. In this work, the hypersonic flows over a three-dimensional rectangular cavity with length-to-width-to-depth ratio, L:W:D, of 19.9:3.57:1 at two angles of attack (AoA) were numerically studied with Improved-Delayed-Detached-Eddy Simulation (IDDES) method to highlight the mechanism of transition triggered by the cavity. The present approach was firstly applied to the transonic flow over M219 rectangular cavity. The results, including the fluctuating pressure and frequency, agreed with experiment well. In the hypersonic case at Mach number about 9.6 the cavity is seen as “open” at AoA of −10° but “closed” at AoA of −15° unconventional to the two-dimensional cavity case where the flow always exhibits closed cavity feature when the length-to-depth ratio L/D is larger than 14. For the open cavity flow, the shear layer is basically steady and the flow maintains laminar. For the closed cavity case, the external flow goes into the cavity and impinges on the bottom floor. High intensity streamwise vortices, impingement shock and exit shock are observed causing breakdown of these vortices triggering rapid flow transition

Green plasticizers for multibase gun propellants (Lecture)

NARCIS (Netherlands)

Schoolderman, C.; Driel, C.A. van; Zebregs, M.

2007-01-01

TNO Defence, Security and Safety has a long history of research on gun propellants. Areas investigated are formulating (new ingredients, optimization), manufacturing, charge design and lifetime assessment [1,2,3,4,5]. In conventional propellants inert plasticizers are used to alter performance,

Small-Scale Shock Testing of Propellants and Ingredients

National Research Council Canada - National Science Library

Dawley, S

2004-01-01

.... The use of small-scale gap testing to evaluate the shock sensitivity of individual propellant ingredients and propellant formulations is a valuable method for experimentally establishing shock...

Effects of superhydrophobic surface on the propeller wake

Science.gov (United States)

Choi, Hongseok; Lee, Jungjin; Park, Hyungmin

2017-11-01

This study investigates the change in propeller wake when the superhydrophobic surface is applied on the propeller blade. The propeller rotates in a quiescent water tank, facing its bottom, with a rotational Reynolds number of 96000. To measure the three-dimensional flow fields, we use stereo PIV and a water prism is installed at the camera-side tank wall. Two cameras are tilted 30 degrees from the normal axis of the tank wall, satisfying schiempflug condition. Superhydrophobic surface is made by coating hydrophobic nanoparticles on the propeller blade. Measurements are done on two vertical planes (at the center of propeller hub and the blade tip), and are ensemble averaged being classified by blade phase of 0 and 90 degrees. Velocity fluctuation, turbulent kinetic energy, and vorticity are evaluated. With superhydrophobic surface, it is found that the turbulence level is significantly (20 - 30 %) reduced with a small penalty (less than 5%) in the streamwise momentum (i.e., thrust) generation. This is because the cone shaped propeller wake gets narrower and organized vortex structures are broken with the superhydrophobic surfaces. More detailed flow analysis will be given. Supported by NRF (NRF-2016R1C1B2012775, NRF-2016M2B2A9A02945068) programs of Korea government.

On nitrogen condensation in hypersonic nozzle flows: Numerical method and parametric study

KAUST Repository

Lin, Longyuan

2013-12-17

A numerical method for calculating two-dimensional planar and axisymmetric hypersonic nozzle flows with nitrogen condensation is developed. The classical nucleation theory with an empirical correction function and the modified Gyarmathy model are used to describe the nucleation rate and the droplet growth, respectively. The conservation of the liquid phase is described by a finite number of moments of the size distribution function. The moment equations are then combined with the Euler equations and are solved by the finite-volume method. The numerical method is first validated by comparing its prediction with experimental results from the literature. The effects of nitrogen condensation on hypersonic nozzle flows are then numerically examined. The parameters at the nozzle exit under the conditions of condensation and no-condensation are evaluated. For the condensation case, the static pressure, the static temperature, and the amount of condensed fluid at the nozzle exit decrease with the increase of the total temperature. Compared with the no-condensation case, both the static pressure and temperature at the nozzle exit increase, and the Mach number decreases due to the nitrogen condensation. It is also indicated that preheating the nitrogen gas is necessary to avoid the nitrogen condensation even for a hypersonic nozzle with a Mach number of 5 operating at room temperatures. © 2013 Springer-Verlag Berlin Heidelberg.

Supercritical Fluid Processing of Propellant Polymers

Science.gov (United States)

1991-01-01

of insoluble material present in the polymeric neopentyl glycol azelate (NPGA) binder. Laub (2) summarizes (from Refs 3 & 4) the various polymeric...binders used in composite propellant formulations for DOD missile systems. e.g., these include: polyneopentyl glycol azelate (NPGA) in HAWK. hydroxy...systems. Composite smokeless propellants containing polyethylene glycol (PEG), polyethylene glycol adepate (PEGA) and polycaprolactone are currently under

HMX based enhanced energy LOVA gun propellant

Energy Technology Data Exchange (ETDEWEB)

Sanghavi, R.R. [High Energy Materials Research Laboratory, Pune 411021 (India)]. E-mail: sanghavirr@yahoo.co.uk; Kamale, P.J. [High Energy Materials Research Laboratory, Pune 411021 (India); Shaikh, M.A.R. [High Energy Materials Research Laboratory, Pune 411021 (India); Shelar, S.D. [High Energy Materials Research Laboratory, Pune 411021 (India); Kumar, K. Sunil [High Energy Materials Research Laboratory, Pune 411021 (India); Singh, Amarjit [High Energy Materials Research Laboratory, Pune 411021 (India)

2007-05-08

Efforts to develop gun propellants with low vulnerability have recently been focused on enhancing the energy with a further improvement in its sensitivity characteristics. These propellants not only prevent catastrophic disasters due to unplanned initiation of currently used gun propellants (based on nitrate esters) but also realize enhanced energy levels to increase the muzzle velocity of the projectiles. Now, in order to replace nitroglycerine, which is highly sensitive to friction and impact, nitramines meet the requirements as they offer superior energy due to positive heat of formation, typical stoichiometry with higher decomposition temperatures and also owing to negative oxygen balance are less sensitive than stoichiometrically balanced NG. RDX has been widely reported for use in LOVA propellant. In this paper we have made an effort to present the work on scantily reported nitramine HMX based LOVA gun propellant while incorporating energetic plasticizer glycidyl azide polymer to enhance the energy level. HMX is known to be thermally stable at higher temperature than RDX and also proved to be less vulnerable to small scale shaped charge jet attack as its decomposition temperature is 270 deg. C. HMX also offers improved impulse due to its superior heat of formation (+17 kcal/mol) as compared to RDX (+14 kcal/mol). It has also been reported that a break point will not appear until 35,000 psi for propellant comprising of 5 {mu}m HMX. Since no work has been reported in open literature regarding replacement of RDX by HMX, the present studies were carried out.

HyPlane for Space Tourism and Business Transportation

Science.gov (United States)

Savino, R.

In the present work a preliminary study on a small hypersonic airplane for a long duration space tourism mission is presented. It is also consistent with a point-to-point medium range (5000-6000 km) hypersonic trip, in the frame of the "urgent business travel" market segment. The main ideas is to transfer technological solutions developed for aeronautical and space atmospheric re-entry systems to the design of such a hypersonic airplane. A winged vehicle characterized by high aerodynamic efficiency and able to manoeuvre along the flight path, in all aerodynamic regimes encountered, is taken into consideration. Rocket-Based Combined Cycle and Turbine-Based Combined Cycle engines are investigated to ensure higher performances in terms of flight duration and range. Different flight-paths are also considered, including sub-orbital parabolic trajectories and steady state hypersonic cruise. The former, in particular, takes advantage of the high aerodynamic efficiency during the unpowered phase, in combination with a periodic engine actuation, to guarantee a long duration oscillating flight path. These trajectories offer Space tourists the opportunity of extended missions, characterized by repeated periods of low-gravity at altitudes high enough to ensure a wide view of the Earth from Space.

Hypersonic Transition and Turbulence with Non-Equilibrium Thermochemistry

Science.gov (United States)

2009-08-31

from the literamre. In summary, this AFOSR MURI project has resulted in the production of new knowledge that should significantly improve the accuracy...behavior. The accumulated knowledge and understanding are expected to help development of better dissipation models for compressible flow fields. 2.23.2...8ffipüC<Pressurt Modieung suggestions from physics study <T acautttc Hypersonic Mach numbers Supersonic Mach numbers * skier * *a Subsonic

Prediction of propeller-induced hull-pressure fluctuations

NARCIS (Netherlands)

Van Wijngaarden, H.C.J.

2011-01-01

The cavitating propeller often forms the primary source of noise and vibration on board ships. The propeller induces hydroacoustic pressure fluctuations due to the passing blades and, more importantly, the dynamic activity of cavities in the propeller’s immediate vicinity. The accurate prediction of

ASTP (SA-210) Launch vehicle operational flight trajectory. Part 3: Final documentation

Science.gov (United States)

Carter, A. B.; Klug, G. W.; Williams, N. W.

1975-01-01

Trajectory data are presented for a nominal and two launch window trajectory simulations. These trajectories are designed to insert a manned Apollo spacecraft into a 150/167 km. (81/90 n. mi.) earth orbit inclined at 51.78 degrees for rendezvous with a Soyuz spacecraft, which will be orbiting at approximately 225 km. (121.5 n. mi.). The launch window allocation defined for this launch is 500 pounds of S-IVB stage propellant. The launch window opening trajectory simulation depicts the earliest launch time deviation from a planar flight launch which conforms to this constraint. The launch window closing trajectory simulation was developed for the more stringent Air Force Eastern Test Range (AFETR) flight azimuth restriction of 37.4 degrees east-of-north. These trajectories enclose a 12.09 minute launch window, pertinent features of which are provided in a tabulation. Planar flight data are included for mid-window reference.

Adaptive fuzzy tracking control for a constrained flexible air-breathing hypersonic vehicle based on actuator compensation

Directory of Open Access Journals (Sweden)

Peng Fei Wang

2016-10-01

Full Text Available The design of an adaptive fuzzy tracking control for a flexible air-breathing hypersonic vehicle with actuator constraints is discussed. Based on functional decomposition methodology, velocity and altitude controllers are designed. Fuzzy logic systems are applied to approximate the lumped uncertainty of each subsystem of air-breathing hypersonic vehicle model. Every controllers contain only one adaptive parameter that needs to be updated online with a minimal-learning-parameter scheme. The back-stepping design is not demanded by converting the altitude subsystem into the normal output-feedback formulation, which predigests the design of a controller. The special contribution is that novel auxiliary systems are developed to compensate both the tracking errors and desired control laws, based on which the explored controller can still provide effective tracking of velocity and altitude commands when the inputs are saturated. Finally, reference trajectory tracking simulation shows the effectiveness of the proposed method in its application to air-breathing hypersonic vehicle control.

75 FR 12148 - Airworthiness Directives; Ontic Engineering and Manufacturing, Inc. Propeller Governors, Part...

Science.gov (United States)

2010-03-15

... propeller pitch by regulating oil pressure to the propeller pitch change mechanism. Changes in governor oil..., the propeller governor cannot control oil pressure to the propeller pitch control mechanism. This... proposing this AD to prevent loss of propeller pitch control, damage to the propeller governor, and internal...

Nonlinear Constrained Adaptive Backstepping Tracking Control for a Hypersonic Vehicle with Uncertainty

Directory of Open Access Journals (Sweden)

Qin Zou

2015-01-01

Full Text Available The control problem of a flexible hypersonic vehicle is presented, where input saturation and aerodynamic uncertainty are considered. A control-oriented model including aerodynamic uncertainty is derived for simple controller design due to the nonlinearity and complexity of hypersonic vehicle model. Then it is separated into velocity subsystem and altitude subsystem. On the basis of the integration of robust adaptive control and backstepping technique, respective controller is designed for each subsystem, where an auxiliary signal provided by an additional dynamic system is used to compensate for the control saturation effect. Then to deal with the “explosion of terms” problem inherent in backstepping control, a novel first-order filter is proposed. Simulation results are included to demonstrate the effectiveness of the adaptive backstepping control scheme.

Self-propelled oil droplets consuming "fuel" surfactant

DEFF Research Database (Denmark)

Toyota, Taro; Maru, Naoto; Hanczyc, Martin M

2009-01-01

A micrometer-sized oil droplet of 4-octylaniline containing 5 mol % of an amphiphilic catalyst exhibited a self-propelled motion, producing tiny oil droplets, in an aqueous dispersion of an amphiphilic precursor of 4-octylaniline. The tiny droplets on the surface of the self-propelled droplet wer...

Structural concepts and experimental considerations for a versatile high-speed research airplane

Science.gov (United States)

Jackson, L. R.; Kirkham, F. S.; Weidner, J. P.

1978-01-01

Future aircraft may be hydrogen fueled and fly at hypersonic speeds. The resulting environments will require new structural concepts to satisfy performance goals. Large representative structures will have to be flight tested prior to commitment to a costly vehicle fleet. To perform flight tests, a versatile, economical, high-speed research airplane is defined. Results of this study including experimental considerations for a hypersonic research airplane are reported.

MEMS-Based Solid Propellant Rocket Array Thruster

Science.gov (United States)

Tanaka, Shuji; Hosokawa, Ryuichiro; Tokudome, Shin-Ichiro; Hori, Keiichi; Saito, Hirobumi; Watanabe, Masashi; Esashi, Masayoshi

The prototype of a solid propellant rocket array thruster for simple attitude control of a 10 kg class micro-spacecraft was completed and tested. The prototype has 10×10 φ0.8 mm solid propellant micro-rockets arrayed at a pitch of 1.2 mm on a 20×22 mm substrate. To realize such a dense array of micro-rockets, each ignition heater is powered from the backside of the thruster through an electrical feedthrough which passes along a propellant cylinder wall. Boron/potassium nitrate propellant (NAB) is used with/without lead rhodanide/potassium chlorate/nitrocellulose ignition aid (RK). Impulse thrust was measured by a pendulum method in air. Ignition required electric power of at least 3 4 W with RK and 4 6 W without RK. Measured impulse thrusts were from 2×10-5 Ns to 3×10-4 Ns after the calculation of compensation for air dumping.

Nuclear thermal rockets using indigenous Martian propellants

International Nuclear Information System (INIS)

Zubrin, R.M.

1989-01-01

This paper considers a novel concept for a Martian descent and ascent vehicle, called NIMF (for nuclear rocket using indigenous Martian fuel), the propulsion for which will be provided by a nuclear thermal reactor which will heat an indigenous Martian propellant gas to form a high-thrust rocket exhaust. The performance of each of the candidate Martian propellants, which include CO2, H2O, CH4, N2, CO, and Ar, is assessed, and the methods of propellant acquisition are examined. Attention is also given to the issues of chemical compatibility between candidate propellants and reactor fuel and cladding materials, and the potential of winged Mars supersonic aircraft driven by this type of engine. It is shown that, by utilizing the nuclear landing craft in combination with a hydrogen-fueled nuclear thermal interplanetary vehicle and a heavy lift booster, it is possible to achieve a manned Mars mission in one launch. 6 refs

A Scramjet Compression System for Hypersonic Air Transportation Vehicle Combined Cycle Engines

Directory of Open Access Journals (Sweden)

Devendra Sen

2018-06-01

Full Text Available This paper proposes a compression system for a scramjet, to be used as part of a combined cycle engine on a hypersonic transport vehicle that can achieve sustained flight at 8 Mach 8. Initially research into scramjet compression system and shock wave interaction was conducted to establish the foundation of the scramjet inlet and isolator sections. A Computational Fluid Dynamics (CFD campaign was conducted, where the shock structure and flow characteristics was analysed between Mach 4.5–8. The compression system of a scramjet is of crucial importance in providing air at suitable Mach number, pressure and temperature to the combustion chamber. The use of turbojet engines in over-under configuration with the scramjet was investigated as well as the study of a combined cycle scramjet-ramjet configuration. It was identified that locating the scramjet in the centre with a rotated ramjet on either side, where its ramps make up the scramjet wall was the most optimal configuration, as it mitigated the effect of the oblique shocks propagating from the scramjet walls into the adjacent ramjet. Furthermore, this meant that the forebody of the vehicle could solely be used as the compression surface by the scramjet. In this paper, the sizing of the scramjet combustion chamber and nozzle were modified to match the flow properties of the oncoming flow with the purpose of producing the most optimum scramjet configuration for the cruise speed of Mach 8. CFD simulations showed that the scramjet inlet did not provide the levels of compression and stagnation pressure recovery initially required. However, it was found that the scramjet provided significantly more thrust than the drag of the aircraft at sustained Mach 8 flight, due to its utilisation of a very aerodynamic vehicle design.

Parametric Study of Cantilever Plates Exposed to Supersonic and Hypersonic Flows

Science.gov (United States)

Sri Harsha, A.; Rizwan, M.; Kuldeep, S.; Giridhara Prasad, A.; Akhil, J.; Nagaraja, S. R.

2017-08-01

Analysis of hypersonic flows associated with re-entry vehicles has gained a lot of significance due to the advancements in Aerospace Engineering. An area that is studied extensively by researchers is the simultaneous reduction aerodynamic drag and aero heating in re-entry vehicles. Out of the many strategies being studied, the use of aerospikes at the stagnation point of the vehicle is found to give favourable results. The structural stability of the aerospike becomes important as it is exposed to very high pressures and temperatures. Keeping this in view, the deflection and vibration of an inclined cantilever plate in hypersonic flow is carried out using ANSYS. Steady state pressure distribution obtained from Fluent is applied as load to the transient structural module for analysis. After due validation of the methods, the effects of parameters like flow Mach number, plate inclination and plate thickness on the deflection and vibration are studied.

Application of supersonic linear theory and hypersonic impact methods to three nonslender hypersonic airplane concepts at Mach numbers from 1.10 to 2.86

Science.gov (United States)

Pittman, J. L.

1979-01-01

Aerodynamic predictions from supersonic linear theory and hypersonic impact theory were compared with experimental data for three hypersonic research airplane concepts over a Mach number range from 1.10 to 2.86. The linear theory gave good lift prediction and fair to good pitching-moment prediction over the Mach number (M) range. The tangent-cone theory predictions were good for lift and fair to good for pitching moment for M more than or equal to 2.0. The combined tangent-cone theory predictions were good for lift and fair to good for pitching moment for M more than or equal to 2.0. The combined tangent-cone/tangent-wedge method gave the least accurate prediction of lift and pitching moment. The zero-lift drag was overestimated, especially for M less than 2.0. The linear theory drag prediction was generally poor, with areas of good agreement only for M less than or equal to 1.2. For M more than or equal to 2.), the tangent-cone method predicted the zero-lift drag most accurately.

Flight Flutter Testing of Rotary Wing Aircraft Using a Control System Oscillation Technique

Science.gov (United States)

Yen, J. G.; Viswanathan, S.; Matthys, C. G.

1976-01-01

A flight flutter testing technique is described in which the rotor controls are oscillated by series actuators to excite the rotor and airframe modes of interest, which are then allowed to decay. The moving block technique is then used to determine the damped frequency and damping variation with rotor speed. The method proved useful for tracking the stability of relatively well damped modes. The results of recently completed flight tests of an experimental soft-in-plane rotor are used to illustrate the technique. Included is a discussion of the application of this technique to investigation of the propeller whirl flutter stability characteristics of the NASA/Army XV-15 VTOL tilt rotor research aircraft.

Image based measurement techniques for aircraft propeller flow diagnostics : Propeller slipstream investigations at high-lift conditions and thrust reverse

NARCIS (Netherlands)

Roosenboom, E.W.M.

2011-01-01

The aim of the thesis is to measure the propeller slipstream properties (velocity and vorticity) and to assess the unsteady and instantaneous behavior of the propeller flow field at high disk loadings, zero thrust and thrust reverse using the image based measurement techniques. Along with its

Computational Tool for Coupled Simulation of Nonequilibrium Hypersonic Flows with Ablation, Phase II

Data.gov (United States)

National Aeronautics and Space Administration — The goal of this SBIR project is to develop a predictive computational tool for the aerothermal environment around ablation-cooled hypersonic atmospheric entry...

Trajectory optimization for A S.S.T.O. using in-flight LOX collection

Science.gov (United States)

Saint-Mard, M.; Hendrick, P.

A key point for a space mission (launch of a satellite, earth observation,…) is the optimization of the vehicle trajectory in order to burn the smallest quantity of propelant and then maximize the payload. This is true for evay space vehicle, but especially it is a crucial point for a Single-Stage-To-Orbit (SSTO) where the choice of a bad trajectory can result in an unrealizable vehicle due to the large airbreathing part of the flight In this study, we discuss the trajectory optimization for a Vertical Take-Off and Horizontal Landing (VTOHL) SSTO using supersonic in-flight atmospheric oxygen collection during a cruise phase (constant speed & constant altitude). This collected oxygen is stored in the LOX tanks and reused in the final rocket phase. This SSTO bas a Blended Body aerodynamic configuration as the one chosen by Lockheed Martin for its new space launcher (VentureStar and X-33). This SSTO uses rocket engines from take-off to Mach 1.7 and also for the exoatmospheric flight phase (that means for an altitude higher than 30km and a Mach number evolution from 6.8 to about 20). Between these two rocket phases, the SSTO is propelled by a subsonic ramjet. To perform this study, we use 2 computer programs (running on a home Computer): the first one allows to estimate the SSTO performances (TOGW, dry weight, hydrogen and oxygen consumptions) for a fixed payload mass and the second one permits the evaluation of the payload mass for a fixed TOGW.

Final Phase Flight Performance and Touchdown Time Assessment of TDV in RLV-TD HEX-01 Mission

Science.gov (United States)

Yadav, Sandeep; Jayakumar, M.; Nizin, Aziya; Kesavabrahmaji, K.; Shyam Mohan, N.

2017-12-01

RLV-TD HEX-01 mission was configured as a precursor flight to actual two stages to orbit vehicle. In this mission RLV-TD was designed as a two stage vehicle for demonstrating the hypersonic flight of a winged body vehicle at Mach No. 5. One of the main objectives of this mission was to generate data for better understanding of new technologies required to design the future vehicle. In this mission, the RLV-TD vehicle was heavily instrumented to get data related to performance of different subsystems. As per the mission design, RLV-TD will land in sea after flight duration of 700 s and travelling a distance of nearly 500 km in Bay of Bengal from the launch site for a nominal trajectory. The visibility studies for telemetry data of vehicle for the nominal and off nominal trajectories were carried out. Based on that, three ground stations were proposed for the telemetry data reception (including one in sea). Even with this scheme it was seen that during the final phase of the flight there will not be any ground station visible to the flight due to low elevation. To have the mission critical data during final phase of the flight, telemetry through INSAT scheme was introduced. During the end of the mission RLV-TD will be landing in the sea on a hypothetical runway. To know the exact time of touchdown for the flight in sea, there was no direct measurement available. Simultaneously there were all chances of losing ground station visibility just before touchdown, making it difficult to assess flight performance during that phase. In this work, telemetry and instrumentation scheme of RLV-TD HEX-01 mission is discussed with an objective to determine the flight performance during the final phase. Further, using various flight sensor data the touchdown time of TDV is assessed for this mission.

Performance Analysis Rim Driven Propeller as a Propulsor using Open Water Test

OpenAIRE

Agoes Santoso; Irfan Syarif Arief; Anggara Tio Kurniawan

2017-01-01

The use of duct in propeller is one of the breakthrough in the development of the propeller. Ducting not only claimed to be increasing efficiency of the propeller, but also capable to protect the propeller from impact therefore propeller lifespan is longer. From that idea then RDP is created. RDP propeller blade are designed to be fix at their housing called Rim, in the other word, the driving force came from it’s rim. On current RDP blade used is non-conventional blade. This thesis will disc...

Non-destructive testing of rocket propellant quality using -X-ray radiography

International Nuclear Information System (INIS)

Arayaprecha, W.

1979-01-01

Currently, X-rays radiography has been used extensively in various industries. In this thesis, X-rays has been used in the study of compaction of rocket propellant. For a rocket, to gain an accurate guidance result, the propellant used must be mixed and compacted thoroughly. The quality control of the production of propellant sticks must be carefully done. In this study of non-destructive quality testing of rocket propellant, at first the ultrasonic rays was used to test its homogeneity. However, because the density of the propellant was too low, the test equipment could not detect any reflected signals from the propellant being tested. Then the new procedure using X-rays radiography was tried. The variables in the test procedure were voltage, amperage and the focal-film distance. Also different types of films were used. The results of this experiment were then used to construct an exposure chart for testing the homogeneity of the rocket propellant. The advantage of this chart is that a tester can use this table with propellant sticks of different sizes if they have similar density to the density specified in the chart. Also, it is not necessary that the mixture of the testing propellant be the same as the ones used to construct this chart

Biofouling Removal and Protein Detection Using a Hypersonic Resonator.

Science.gov (United States)

Pan, Shuting; Zhang, Hongxiang; Liu, Wenpeng; Wang, Yanyan; Pang, Wei; Duan, Xuexin

2017-08-25

Nonspecific binding (NSB) is a general issue for surface based biosensors. Various approaches have been developed to prevent or remove the NSBs. However, these approaches either increased the background signals of the sensors or limited to specific transducers interface. In this work, we developed a hydrodynamic approach to selectively remove the NSBs using a microfabricated hypersonic resonator with 2.5 gigahertz (GHz) resonant frequency. The high frequency device facilitates generation of multiple controlled microvortexes which then create cleaning forces at the solid-liquid interfaces. The competitive adhesive and cleaning forces have been investigated using the finite element method (FEM) simulation, identifying the feasibility of the vortex-induced NSB removal. NSB proteins have been selectively removed experimentally both on the surface of the resonator and on other substrates which contact the vortexes. Thus, the developed hydrodynamic approach is believed to be a simple and versatile tool for NSB removal and compatible to many sensor systems. The unique feature of the hypersonic resonator is that it can be used as a gravimetric sensor as well; thus a combined NSB removal and protein detection dual functional biosensor system is developed.

Impulse Force Balance for Ultrashort Duration Hypersonic Test Facilities

Directory of Open Access Journals (Sweden)

P. Singh

2015-01-01

Full Text Available This paper presents the measurement of side force, pitching, and yawing moments on a model, using an accelerometer force balance, in a short duration hypersonic shock tunnel. The test model is a blunt-nosed, flapped delta wing, mounted on a support sting through a force balance. The flexible rubber bushes constituting the balance allow the model to float freely on the sting during the test. The accelerometers were located in the model to record accelerations in the directions of interest. The model was tested in shock tunnel at Mach 8 at different angles of incidence with the freestream. Dynamic calibration of the test assembly was carried out for the acquisition of impulse response functions for the above components of force and moments, using an impulse hammer. The convolution technique was applied to derive the impulse response functions. The accelerometer outputs from the model in the hypersonic freestream were processed using the respective impulse response functions to derive the unknown aerodynamic force and moments. The newly adopted convolution technique has been found very effective for data reduction from accelerometer force balances developed for shock tunnel applications.

Tile Surface Thermocouple Measurement Challenges from the Orbiter Boundary Layer Transition Flight Experiment

Science.gov (United States)

Campbell, Charles H.; Berger, Karen; Anderson, Brian

2012-01-01

Hypersonic entry flight testing motivated by efforts seeking to characterize boundary layer transition on the Space Shuttle Orbiters have identified challenges in our ability to acquire high quality quantitative surface temperature measurements versus time. Five missions near the end of the Space Shuttle Program implemented a tile surface protuberance as a boundary layer trip together with tile surface thermocouples to capture temperature measurements during entry. Similar engineering implementations of these measurements on Discovery and Endeavor demonstrated unexpected measurement voltage response during the high heating portion of the entry trajectory. An assessment has been performed to characterize possible causes of the issues experienced during STS-119, STS-128, STS-131, STS-133 and STS-134 as well as similar issues encountered during other orbiter entries.

Mode Transition Variable Geometry for High Speed Inlets for Hypersonic Aircraft, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Hypersonic propulsion research has been a focus of the NASA aeronautics program for years. Previous high-speed cruise and space access programs have examined the...

Simplified Thermo-Chemical Modelling For Hypersonic Flow

Science.gov (United States)

Sancho, Jorge; Alvarez, Paula; Gonzalez, Ezequiel; Rodriguez, Manuel

2011-05-01

Hypersonic flows are connected with high temperatures, generally associated with strong shock waves that appear in such flows. At high temperatures vibrational degrees of freedom of the molecules may become excited, the molecules may dissociate into atoms, the molecules or free atoms may ionize, and molecular or ionic species, unimportant at lower temperatures, may be formed. In order to take into account these effects, a chemical model is needed, but this model should be simplified in order to be handled by a CFD code, but with a sufficient precision to take into account the physics more important. This work is related to a chemical non-equilibrium model validation, implemented into a commercial CFD code, in order to obtain the flow field around bodies in hypersonic flow. The selected non-equilibrium model is composed of seven species and six direct reactions together with their inverse. The commercial CFD code where the non- equilibrium model has been implemented is FLUENT. For the validation, the X38/Sphynx Mach 20 case is rebuilt on a reduced geometry, including the 1/3 Lref forebody. This case has been run in laminar regime, non catalytic wall and with radiative equilibrium wall temperature. The validated non-equilibrium model is applied to the EXPERT (European Experimental Re-entry Test-bed) vehicle at a specified trajectory point (Mach number 14). This case has been run also in laminar regime, non catalytic wall and with radiative equilibrium wall temperature.

Computer aided design and development of mixed-propeller pumps

International Nuclear Information System (INIS)

Bhaoyal, B.C.

1994-01-01

This paper deals with the design principle of mixed propeller hydraulic aided by CADD software developed by author for generation of the hydraulic profile of the mixed propeller and diffuser geometry. The design methodology for plotting the vane profile of mixed propeller pump has been discussed in detail with special reference to conformal transformation in cylindrical as well as conical plane. (author). 10 refs., 11 figs

An Upgrade of the Imaging for Hypersonic Experimental Aeroheating Testing (IHEAT) Software

Science.gov (United States)

Mason, Michelle L.; Rufer, Shann J.

2015-01-01

The Imaging for Hypersonic Experimental Aeroheating Testing (IHEAT) code is used at NASA Langley Research Center to analyze global aeroheating data on wind tunnel models tested in the Langley Aerothermodynamics Laboratory. One-dimensional, semi-infinite heating data derived from IHEAT are used to design thermal protection systems to mitigate the risks due to the aeroheating loads on hypersonic vehicles, such as re-entry vehicles during descent and landing procedures. This code was originally written in the PV-WAVE programming language to analyze phosphor thermography data from the two-color, relativeintensity system developed at Langley. To increase the efficiency, functionality, and reliability of IHEAT, the code was migrated to MATLAB syntax and compiled as a stand-alone executable file labeled version 4.0. New features of IHEAT 4.0 include the options to batch process all of the data from a wind tunnel run, to map the two-dimensional heating distribution to a three-dimensional computer-aided design model of the vehicle to be viewed in Tecplot, and to extract data from a segmented line that follows an interesting feature in the data. Results from IHEAT 4.0 were compared on a pixel level to the output images from the legacy code to validate the program. The differences between the two codes were on the order of 10-5 to 10-7. IHEAT 4.0 replaces the PV-WAVE version as the production code for aeroheating experiments conducted in the hypersonic facilities at NASA Langley.

Hyper-X Program Status

Science.gov (United States)

McClinton, Charles R.; Reubush, David E.; Sitz, Joel; Reukauf, Paul

2001-01-01

This paper provides an overview of the objectives and status of the Hyper-X program, which is tailored to move hypersonic, airbreathing vehicle technology from the laboratory environment to the flight environment. The first Hyper-X research vehicle (HXRV), designated X-43, is being prepared at the Dryden Flight Research Center for flight at Mach 7. Extensive risk reduction activities for the first flight are completed, and non-recurring design activities for the Mach 10 X-43 (third flight) are nearing completion. The Mach 7 flight of the X-43, in the spring of 2001, will be the first flight of an airframe-integrated scramjet-powered vehicle. The Hyper-X program is continuing to plan follow-on activities to focus an orderly continuation of hypersonic technology development through flight research.