Launch Vehicles Based on Advanced Hybrid Rocket Motors: An Enabling Technology for the Commercial Small and Micro Satellite Planetary Science

Science.gov (United States)

Karabeyoglu, Arif; Tuncer, Onur; Inalhan, Gokhan

2016-07-01

Mankind is relient on chemical propulsion systems for space access. Nevertheless, this has been a stagnant area in terms of technological development and the technology base has not changed much almost for the past forty years. This poses a vicious circle for launch applications such that high launch costs constrain the demand and low launch freqencies drive costs higher. This also has been a key limiting factor for small and micro satellites that are geared towards planetary science. Rather this be because of the launch frequencies or the costs, the access of small and micro satellites to orbit has been limited. With today's technology it is not possible to escape this circle. However the emergence of cost effective and high performance propulsion systems such as advanced hybrid rockets can decrease launch costs by almost an order or magnitude. This paper briefly introduces the timeline and research challenges that were overcome during the development of advanced hybrid LOX/paraffin based rockets. Experimental studies demonstrated effectiveness of these advanced hybrid rockets which incorporate fast burning parafin based fuels, advanced yet simple internal balistic design and carbon composite winding/fuel casting technology that enables the rocket motor to be built from inside out. A feasibility scenario is studied using these rocket motors as building blocks for a modular launch vehicle capable of delivering micro satellites into low earth orbit. In addition, the building block rocket motor can be used further solar system missions providing the ability to do standalone small and micro satellite missions to planets within the solar system. This enabling technology therefore offers a viable alternative in order to escape the viscous that has plagued the space launch industry and that has limited the small and micro satellite delivery for planetary science.

Rocket Flight Path

Directory of Open Access Journals (Sweden)

Jamie Waters

2014-09-01

Full Text Available This project uses Newton’s Second Law of Motion, Euler’s method, basic physics, and basic calculus to model the flight path of a rocket. From this, one can find the height and velocity at any point from launch to the maximum altitude, or apogee. This can then be compared to the actual values to see if the method of estimation is a plausible. The rocket used for this project is modeled after Bullistic-1 which was launched by the Society of Aeronautics and Rocketry at the University of South Florida.

Modeling Powered Aerodynamics for the Orion Launch Abort Vehicle Aerodynamic Database

Science.gov (United States)

Chan, David T.; Walker, Eric L.; Robinson, Philip E.; Wilson, Thomas M.

2011-01-01

Modeling the aerodynamics of the Orion Launch Abort Vehicle (LAV) has presented many technical challenges to the developers of the Orion aerodynamic database. During a launch abort event, the aerodynamic environment around the LAV is very complex as multiple solid rocket plumes interact with each other and the vehicle. It is further complicated by vehicle separation events such as between the LAV and the launch vehicle stack or between the launch abort tower and the crew module. The aerodynamic database for the LAV was developed mainly from wind tunnel tests involving powered jet simulations of the rocket exhaust plumes, supported by computational fluid dynamic simulations. However, limitations in both methods have made it difficult to properly capture the aerodynamics of the LAV in experimental and numerical simulations. These limitations have also influenced decisions regarding the modeling and structure of the aerodynamic database for the LAV and led to compromises and creative solutions. Two database modeling approaches are presented in this paper (incremental aerodynamics and total aerodynamics), with examples showing strengths and weaknesses of each approach. In addition, the unique problems presented to the database developers by the large data space required for modeling a launch abort event illustrate the complexities of working with multi-dimensional data.

Thirteenth Workshop for Computational Fluid Dynamic Applications in Rocket Propulsion and Launch Vehicle Technology. Volume 1

Science.gov (United States)

Williams, R. W. (Compiler)

1996-01-01

The purpose of the workshop was to discuss experimental and computational fluid dynamic activities in rocket propulsion and launch vehicles. The workshop was an open meeting for government, industry, and academia. A broad number of topics were discussed including computational fluid dynamic methodology, liquid and solid rocket propulsion, turbomachinery, combustion, heat transfer, and grid generation.

Multidisciplinary design of a rocket-based combined cycle SSTO launch vehicle using Taguchi methods

Science.gov (United States)

Olds, John R.; Walberg, Gerald D.

1993-01-01

Results are presented from the optimization process of a winged-cone configuration SSTO launch vehicle that employs a rocket-based ejector/ramjet/scramjet/rocket operational mode variable-cycle engine. The Taguchi multidisciplinary parametric-design method was used to evaluate the effects of simultaneously changing a total of eight design variables, rather than changing them one at a time as in conventional tradeoff studies. A combination of design variables was in this way identified which yields very attractive vehicle dry and gross weights.

Vibration control of uncertain multiple launch rocket system using radial basis function neural network

Science.gov (United States)

Li, Bo; Rui, Xiaoting

2018-01-01

Poor dispersion characteristics of rockets due to the vibration of Multiple Launch Rocket System (MLRS) have always restricted the MLRS development for several decades. Vibration control is a key technique to improve the dispersion characteristics of rockets. For a mechanical system such as MLRS, the major difficulty in designing an appropriate control strategy that can achieve the desired vibration control performance is to guarantee the robustness and stability of the control system under the occurrence of uncertainties and nonlinearities. To approach this problem, a computed torque controller integrated with a radial basis function neural network is proposed to achieve the high-precision vibration control for MLRS. In this paper, the vibration response of a computed torque controlled MLRS is described. The azimuth and elevation mechanisms of the MLRS are driven by permanent magnet synchronous motors and supposed to be rigid. First, the dynamic model of motor-mechanism coupling system is established using Lagrange method and field-oriented control theory. Then, in order to deal with the nonlinearities, a computed torque controller is designed to control the vibration of the MLRS when it is firing a salvo of rockets. Furthermore, to compensate for the lumped uncertainty due to parametric variations and un-modeled dynamics in the design of the computed torque controller, a radial basis function neural network estimator is developed to adapt the uncertainty based on Lyapunov stability theory. Finally, the simulated results demonstrate the effectiveness of the proposed control system and show that the proposed controller is robust with regard to the uncertainty.

Modeling and validation of Ku-band signal attenuation through rocket plumes

NARCIS (Netherlands)

Veek, van der B.J.; Chintalapati, S.; Kirk, D.R.; Gutierrez, H.; Bun, R.F.

2013-01-01

Communications to and from a launch vehicle during ascent are of critical importance to the success of rocket-launch operations. During ascent, the rocket's exhaust plume causes significant interference in the radio communications between the vehicle and ground station. This paper presents an

Propulsion and launching analysis of variable-mass rockets by analytical methods

OpenAIRE

D.D. Ganji; M. Gorji; M. Hatami; A. Hasanpour; N. Khademzadeh

2013-01-01

In this study, applications of some analytical methods on nonlinear equation of the launching of a rocket with variable mass are investigated. Differential transformation method (DTM), homotopy perturbation method (HPM) and least square method (LSM) were applied and their results are compared with numerical solution. An excellent agreement with analytical methods and numerical ones is observed in the results and this reveals that analytical methods are effective and convenient. Also a paramet...

Infrared Imagery of Solid Rocket Exhaust Plumes

Science.gov (United States)

Moran, Robert P.; Houston, Janice D.

2011-01-01

The Ares I Scale Model Acoustic Test program consisted of a series of 18 solid rocket motor static firings, simulating the liftoff conditions of the Ares I five-segment Reusable Solid Rocket Motor Vehicle. Primary test objectives included acquiring acoustic and pressure data which will be used to validate analytical models for the prediction of Ares 1 liftoff acoustics and ignition overpressure environments. The test article consisted of a 5% scale Ares I vehicle and launch tower mounted on the Mobile Launch Pad. The testing also incorporated several Water Sound Suppression Systems. Infrared imagery was employed during the solid rocket testing to support the validation or improvement of analytical models, and identify corollaries between rocket plume size or shape and the accompanying measured level of noise suppression obtained by water sound suppression systems.

Infrared signature modelling of a rocket jet plume - comparison with flight measurements

International Nuclear Information System (INIS)

Rialland, V; Perez, P; Roblin, A; Guy, A; Gueyffier, D; Smithson, T

2016-01-01

The infrared signature modelling of rocket plumes is a challenging problem involving rocket geometry, propellant composition, combustion modelling, trajectory calculations, fluid mechanics, atmosphere modelling, calculation of gas and particles radiative properties and of radiative transfer through the atmosphere. This paper presents ONERA simulation tools chained together to achieve infrared signature prediction, and the comparison of the estimated and measured signatures of an in-flight rocket plume. We consider the case of a solid rocket motor with aluminized propellant, the Black Brant sounding rocket. The calculation case reproduces the conditions of an experimental rocket launch, performed at White Sands in 1997, for which we obtained high quality infrared signature data sets from DRDC Valcartier. The jet plume is calculated using an in-house CFD software called CEDRE. The plume infrared signature is then computed on the spectral interval 1900-5000 cm -1 with a step of 5 cm -1 . The models and their hypotheses are presented and discussed. Then the resulting plume properties, radiance and spectra are detailed. Finally, the estimated infrared signature is compared with the spectral imaging measurements. The discrepancies are analyzed and discussed. (paper)

Assessment of exposure-response functions for rocket-emission toxicants

National Research Council Canada - National Science Library

Subcommittee on Rocket-Emission Toxicants, National Research Council

... aborted launch that results in a rocket being destroyed near the ground. Assessment of Exposure-Response Functions for Rocket-Emmission Toxicants evaluates the model and the data used for three rocket emission toxicants...

NASA Lewis Launch Collision Probability Model Developed and Analyzed

Science.gov (United States)

Bollenbacher, Gary; Guptill, James D

1999-01-01

There are nearly 10,000 tracked objects orbiting the earth. These objects encompass manned objects, active and decommissioned satellites, spent rocket bodies, and debris. They range from a few centimeters across to the size of the MIR space station. Anytime a new satellite is launched, the launch vehicle with its payload attached passes through an area of space in which these objects orbit. Although the population density of these objects is low, there always is a small but finite probability of collision between the launch vehicle and one or more of these space objects. Even though the probability of collision is very low, for some payloads even this small risk is unacceptable. To mitigate the small risk of collision associated with launching at an arbitrary time within the daily launch window, NASA performs a prelaunch mission assurance Collision Avoidance Analysis (or COLA). For the COLA of the Cassini spacecraft, the NASA Lewis Research Center conducted an in-house development and analysis of a model for launch collision probability. The model allows a minimum clearance criteria to be used with the COLA analysis to ensure an acceptably low probability of collision. If, for any given liftoff time, the nominal launch vehicle trajectory would pass a space object with less than the minimum required clearance, launch would not be attempted at that time. The model assumes that the nominal positions of the orbiting objects and of the launch vehicle can be predicted as a function of time, and therefore, that any tracked object that comes within close proximity of the launch vehicle can be identified. For any such pair, these nominal positions can be used to calculate a nominal miss distance. The actual miss distances may differ substantially from the nominal miss distance, due, in part, to the statistical uncertainty of the knowledge of the objects positions. The model further assumes that these position uncertainties can be described with position covariance matrices

History of Solid Rockets

Science.gov (United States)

Green, Rebecca

2017-01-01

Solid rockets are of interest to the space program because they are commonly used as boosters that provide the additional thrust needed for the space launch vehicle to escape the gravitational pull of the Earth. Larger, more advanced solid rockets allow for space launch vehicles with larger payload capacities, enabling mankind to reach new depths of space. This presentation will discuss, in detail, the history of solid rockets. The history begins with the invention and origin of the solid rocket, and then goes into the early uses and design of the solid rocket. The evolution of solid rockets is depicted by a description of how solid rockets changed and improved and how they were used throughout the 16th, 17th, 18th, and 19th centuries. Modern uses of the solid rocket include the Solid Rocket Boosters (SRBs) on the Space Shuttle and the solid rockets used on current space launch vehicles. The functions and design of the SRB and the advancements in solid rocket technology since the use of the SRB are discussed as well. Common failure modes and design difficulties are discussed as well.

Cryogenic rocket engine development at Delft aerospace rocket engineering

NARCIS (Netherlands)

Wink, J; Hermsen, R.; Huijsman, R; Akkermans, C.; Denies, L.; Barreiro, F.; Schutte, A.; Cervone, A.; Zandbergen, B.T.C.

2016-01-01

This paper describes the current developments regarding cryogenic rocket engine technology at Delft Aerospace Rocket Engineering (DARE). DARE is a student society based at Delft University of Technology with the goal of being the first student group in the world to launch a rocket into space. After

The Space Launch System -The Biggest, Most Capable Rocket Ever Built, for Entirely New Human Exploration Missions Beyond Earth's Orbit

Science.gov (United States)

Shivers, C. Herb

2012-01-01

NASA is developing the Space Launch System -- an advanced heavy-lift launch vehicle that will provide an entirely new capability for human exploration beyond Earth's orbit. The Space Launch System will provide a safe, affordable and sustainable means of reaching beyond our current limits and opening up new discoveries from the unique vantage point of space. The first developmental flight, or mission, is targeted for the end of 2017. The Space Launch System, or SLS, will be designed to carry the Orion Multi-Purpose Crew Vehicle, as well as important cargo, equipment and science experiments to Earth's orbit and destinations beyond. Additionally, the SLS will serve as a backup for commercial and international partner transportation services to the International Space Station. The SLS rocket will incorporate technological investments from the Space Shuttle Program and the Constellation Program in order to take advantage of proven hardware and cutting-edge tooling and manufacturing technology that will significantly reduce development and operations costs. The rocket will use a liquid hydrogen and liquid oxygen propulsion system, which will include the RS-25D/E from the Space Shuttle Program for the core stage and the J-2X engine for the upper stage. SLS will also use solid rocket boosters for the initial development flights, while follow-on boosters will be competed based on performance requirements and affordability considerations.

How High? How Fast? How Long? Modeling Water Rocket Flight with Calculus

Science.gov (United States)

Ashline, George; Ellis-Monaghan, Joanna

2006-01-01

We describe an easy and fun project using water rockets to demonstrate applications of single variable calculus concepts. We provide procedures and a supplies list for launching and videotaping a water rocket flight to provide the experimental data. Because of factors such as fuel expulsion and wind effects, the water rocket does not follow the…

Pegasus Rocket Model

Science.gov (United States)

1996-01-01

A small, desk-top model of Orbital Sciences Corporation's Pegasus winged rocket booster. Pegasus is an air-launched space booster produced by Orbital Sciences Corporation and Hercules Aerospace Company (initially; later, Alliant Tech Systems) to provide small satellite users with a cost-effective, flexible, and reliable method for placing payloads into low earth orbit. Pegasus has been used to launch a number of satellites and the PHYSX experiment. That experiment consisted of a smooth glove installed on the first-stage delta wing of the Pegasus. The glove was used to gather data at speeds of up to Mach 8 and at altitudes approaching 200,000 feet. The flight took place on October 22, 1998. The PHYSX experiment focused on determining where boundary-layer transition occurs on the glove and on identifying the flow mechanism causing transition over the glove. Data from this flight-research effort included temperature, heat transfer, pressure measurements, airflow, and trajectory reconstruction. Hypersonic flight-research programs are an approach to validate design methods for hypersonic vehicles (those that fly more than five times the speed of sound, or Mach 5). Dryden Flight Research Center, Edwards, California, provided overall management of the glove experiment, glove design, and buildup. Dryden also was responsible for conducting the flight tests. Langley Research Center, Hampton, Virginia, was responsible for the design of the aerodynamic glove as well as development of sensor and instrumentation systems for the glove. Other participating NASA centers included Ames Research Center, Mountain View, California; Goddard Space Flight Center, Greenbelt, Maryland; and Kennedy Space Center, Florida. Orbital Sciences Corporation, Dulles, Virginia, is the manufacturer of the Pegasus vehicle, while Vandenberg Air Force Base served as a pre-launch assembly facility for the launch that included the PHYSX experiment. NASA used data from Pegasus launches to obtain considerable

Rocket Flight.

Science.gov (United States)

Van Evera, Bill; Sterling, Donna R.

2002-01-01

Describes an activity for designing, building, and launching rockets that provides students with an intrinsically motivating and real-life application of what could have been classroom-only concepts. Includes rocket design guidelines and a sample grading rubric. (KHR)

Ramjet Application Possibilities for Increasing Fire Range of the Multiple Launch Rocket Systems Ammunition

Directory of Open Access Journals (Sweden)

V. N. Zubov

2015-01-01

Full Text Available The article considers a possibility to increase a flying range of the perspective rockets equipped with the control unit with aerodynamic controllers for the multiple launch rocket systems “Smerch”.To increase a flying range and reduce a starting mass of the rocket, the paper studies a possibility to replace the single-mode rocket engine used in the solid-fuel rocket motor for the direct-flow propulsion jet engine (DFPJE with not head sector air intakes. The DFPJE is implemented according to the classical scheme with a fuel charged in the combustion chamber. A separated solid propellant starting accelerator provides the rocket acceleration to reach a speed necessary for the DFPJE to run.When designing the DFPJE a proper choice of not head air intake parameters is one of the most difficult points. For this purpose a COSMOS Flow Simulation software package and analytical dependences were used to define the following: a boundary layer thickness where an air intake is set, maximum permissible and appropriate angles of attack and deviation angles of controllers at the section where the DFPJE works, and some other parameters as well.Calculation of DFPJE characteristics consisted in determining parameters of an air-gas path of the propulsion system, geometrical sizes of the pipeline flow area, sizes of a fuel charge, and dependence of the propulsion system impulse on the flight height and speed. Calculations were performed both in thermodynamic statement of problem and in using software package of COSMOS Flow Simulation.As a result of calculations and design engineering activities the air intake profile is created and mass-dimensional characteristics of DFPJE are defined. Besides, calculations of the starting solid fuel accelerator were carried out. Further design allowed us to create the rocket shape, estimate its mass-dimensional characteristics, and perform ballistic calculations, which proved that achieving a range of 120 km for the rocket is

Launch Environmental Test for KITSAT-3 FM

Directory of Open Access Journals (Sweden)

Sang-Hyun Lee

1999-06-01

Full Text Available The satellite experiences the severe launch environment such as vibration, acceleration, shock, and acoustics induced by rocket. Therefore, the satellite should be designed and manufactured to endure such severe launch environments. In this paper, we describe the structure of the KITSAT-3 FM(Flight Model and the processes and results of the launch environmental test to ensure the reliability during launch period.

The second stage of a Titan II rocket is lifted for mating at the launch tower, Vandenberg AFB

Science.gov (United States)

2000-01-01

At the launch tower, Vandenberg Air Force Base, Calif., the second stage of a Titan II rocket is lifted to vertical. The Titan will power the launch of a National Oceanic and Atmospheric Administration (NOAA-L) satellite scheduled no earlier than Sept. 12. NOAA-L is part of the Polar-Orbiting Operational Environmental Satellite (POES) program that provides atmospheric measurements of temperature, humidity, ozone and cloud images, tracking weather patterns that affect the global weather and climate. Pressure-Equalizing Cradle for Booster Rocket Mounting

Science.gov (United States)

Rutan, Elbert L. (Inventor)

2015-01-01

A launch system and method improve the launch efficiency of a booster rocket and payload. A launch aircraft atop which the booster rocket is mounted in a cradle, is flown or towed to an elevation at which the booster rocket is released. The cradle provides for reduced structural requirements for the booster rocket by including a compressible layer, that may be provided by a plurality of gas or liquid-filled flexible chambers. The compressible layer contacts the booster rocket along most of the length of the booster rocket to distribute applied pressure, nearly eliminating bending loads. Distributing the pressure eliminates point loading conditions and bending moments that would otherwise be generated in the booster rocket structure during carrying. The chambers may be balloons distributed in rows and columns within the cradle or cylindrical chambers extending along a length of the cradle. The cradle may include a manifold communicating gas between chambers.

Overview of GX launch services by GALEX

Science.gov (United States)

Sato, Koji; Kondou, Yoshirou

2006-07-01

Galaxy Express Corporation (GALEX) is a launch service company in Japan to develop a medium size rocket, GX rocket and to provide commercial launch services for medium/small low Earth orbit (LEO) and Sun synchronous orbit (SSO) payloads with a future potential for small geo-stationary transfer orbit (GTO). It is GALEX's view that small/medium LEO/SSO payloads compose of medium scaled but stable launch market due to the nature of the missions. GX rocket is a two-stage rocket of well flight proven liquid oxygen (LOX)/kerosene booster and LOX/liquid natural gas (LNG) upper stage. This LOX/LNG propulsion under development by Japan's Aerospace Exploration Agency (JAXA), is robust with comparable performance as other propulsions and have future potential for wider application such as exploration programs. GX rocket is being developed through a joint work between the industries and GX rocket is applying a business oriented approach in order to realize competitive launch services for which well flight proven hardware and necessary new technology are to be introduced as much as possible. It is GALEX's goal to offer “Easy Access to Space”, a highly reliable and user-friendly launch services with a competitive price. GX commercial launch will start in Japanese fiscal year (JFY) 2007 2008.

Rocket measurements of electron density irregularities during MAC/SINE

Science.gov (United States)

Ulwick, J. C.

1989-01-01

Four Super Arcas rockets were launched at the Andoya Rocket Range, Norway, as part of the MAC/SINE campaign to measure electron density irregularities with high spatial resolution in the cold summer polar mesosphere. They were launched as part of two salvos: the turbulent/gravity wave salvo (3 rockets) and the EISCAT/SOUSY radar salvo (one rocket). In both salvos meteorological rockets, measuring temperature and winds, were also launched and the SOUSY radar, located near the launch site, measured mesospheric turbulence. Electron density irregularities and strong gradients were measured by the rocket probes in the region of most intense backscatter observed by the radar. The electron density profiles (8 to 4 on ascent and 4 on descent) show very different characteristics in the peak scattering region and show marked spatial and temporal variability. These data are intercompared and discussed.

Space Launch System Base Heating Test: Sub-Scale Rocket Engine/Motor Design, Development and Performance Analysis

Science.gov (United States)

Mehta, Manish; Seaford, Mark; Kovarik, Brian; Dufrene, Aaron; Solly, Nathan; Kirchner, Robert; Engel, Carl D.

2014-01-01

The Space Launch System (SLS) base heating test is broken down into two test programs: (1) Pathfinder and (2) Main Test. The Pathfinder Test Program focuses on the design, development, hot-fire test and performance analyses of the 2% sub-scale SLS core-stage and booster element propulsion systems. The core-stage propulsion system is composed of four gaseous oxygen/hydrogen RS-25D model engines and the booster element is composed of two aluminum-based model solid rocket motors (SRMs). The first section of the paper discusses the motivation and test facility specifications for the test program. The second section briefly investigates the internal flow path of the design. The third section briefly shows the performance of the model RS-25D engines and SRMs for the conducted short duration hot-fire tests. Good agreement is observed based on design prediction analysis and test data. This program is a challenging research and development effort that has not been attempted in 40+ years for a NASA vehicle.

Using Virtual Reality in K-12 Education: A Simulation of Shooting Bottle Rockets for Distance

Directory of Open Access Journals (Sweden)

Charles Nippert

2012-10-01

Full Text Available Typically, it is often more challenging to shoot bottle rockets for distance instead of shooting them straight up and measuring altitude, as is often done.Using a device made from pipe and wood to launch bottle rockets and control the launch angle creates a much more interesting problem for students who are attempting to optimize launch conditions.Plans are presented for a launcher that allow students to adjust the launch angle. To help embellish the exercise, we supplement the bottle rocket with a model using virtual reality and a photorealistic simulation of the launch that allows the students to appreciate the optimization problems associated with water and air pressure and launch angle. Our usage data indicates that students easily adapt to the virtual reality simulation and use our simulation for intuitive experiments on their own to optimize launch conditions.

The Spanish national programme of balloons and sounding rockets

International Nuclear Information System (INIS)

Casas, J.; Pueyo, L.

1978-01-01

The main points of the Spanish scientific programme are briefly described: CONIE/NASA cooperative project on meteorological sounding rocket launchings; ozonospheric programme; CONIE/NASA/CNES cooperative ionospheric sounding rocket project; D-layer research; rocket infrared dayglow measurements; ultraviolet astronomy research; cosmic ray research. The schedule of sounding rocket launchings at El Arenosillo station during 1977 is given

A rapid compensation method for launch data of long-range rockets under influence of the Earth's disturbing gravity field

Directory of Open Access Journals (Sweden)

Baolin MA

2017-06-01

Full Text Available Regarding the rapid compensation of the influence of the Earth’ s disturbing gravity field upon trajectory calculation, the key point lies in how to derive the analytical solutions to the partial derivatives of the state of burnout point with respect to the launch data. In view of this, this paper mainly expounds on two issues: one is based on the approximate analytical solution to the motion equation for the vacuum flight section of a long-range rocket, deriving the analytical solutions to the partial derivatives of the state of burnout point with respect to the changing rate of the final-stage pitch program; the other is based on the initial positioning and orientation error propagation mechanism, proposing the analytical calculation formula for the partial derivatives of the state of burnout point with respect to the launch azimuth. The calculation results of correction data are simulated and verified under different circumstances. The simulation results are as follows: (1 the accuracy of approximation between the analytical solutions and the results attained via the difference method is higher than 90%, and the ratio of calculation time between them is lower than 0.2%, thus demonstrating the accuracy of calculation of data corrections and advantages in calculation speed; (2 after the analytical solutions are compensated, the longitudinal landing deviation of the rocket is less than 20 m and the lateral landing deviation of the rocket is less than 10 m, demonstrating that the corrected data can meet the requirements for the hit accuracy of a long-range rocket.

Preliminary Sizing Completed for Single- Stage-To-Orbit Launch Vehicles Powered By Rocket-Based Combined Cycle Technology

Science.gov (United States)

Roche, Joseph M.

2002-01-01

Single-stage-to-orbit (SSTO) propulsion remains an elusive goal for launch vehicles. The physics of the problem is leading developers to a search for higher propulsion performance than is available with all-rocket power. Rocket-based combined cycle (RBCC) technology provides additional propulsion performance that may enable SSTO flight. Structural efficiency is also a major driving force in enabling SSTO flight. Increases in performance with RBCC propulsion are offset with the added size of the propulsion system. Geometrical considerations must be exploited to minimize the weight. Integration of the propulsion system with the vehicle must be carefully planned such that aeroperformance is not degraded and the air-breathing performance is enhanced. Consequently, the vehicle's structural architecture becomes one with the propulsion system architecture. Geometrical considerations applied to the integrated vehicle lead to low drag and high structural and volumetric efficiency. Sizing of the SSTO launch vehicle (GTX) is itself an elusive task. The weight of the vehicle depends strongly on the propellant required to meet the mission requirements. Changes in propellant requirements result in changes in the size of the vehicle, which in turn, affect the weight of the vehicle and change the propellant requirements. An iterative approach is necessary to size the vehicle to meet the flight requirements. GTX Sizer was developed to do exactly this. The governing geometry was built into a spreadsheet model along with scaling relationships. The scaling laws attempt to maintain structural integrity as the vehicle size is changed. Key aerodynamic relationships are maintained as the vehicle size is changed. The closed weight and center of gravity are displayed graphically on a plot of the synthesized vehicle. In addition, comprehensive tabular data of the subsystem weights and centers of gravity are generated. The model has been verified for accuracy with finite element analysis. The

Launch Excitement with Water Rockets

Science.gov (United States)

Sanchez, Juan Carlos; Penick, John

2007-01-01

Explosions and fires--these are what many students are waiting for in science classes. And when they do occur, students pay attention. While we can't entertain our students with continual mayhem, we can catch their attention and cater to their desires for excitement by saying, "Let's make rockets." In this activity, students make simple, reusable…

Modeling of "Stripe" Wave Phenomena Seen by the CHARM II and ACES Sounding Rockets

Science.gov (United States)

Dombrowski, M. P.; Labelle, J. W.

2010-12-01

Two recent sounding-rocket missions—CHARM II and ACES—have been launched from Poker Flat Research Range, carrying the Dartmouth High-Frequency Experiment (HFE) among their primary instruments. The HFE is a receiver system which effectively yields continuous (100% duty cycle) E-field waveform measurements up to 5 MHz. The CHARM II sounding rocket was launched 9:49 UT on 15 February 2010 into a substorm, while the ACES mission consisted of two rockets, launched into quiet aurora at 9:49 and 9:50 UT on 29 January 2009. At approximately 350 km on CHARM II and the ACES High-Flyer, the HFE detected short (~2s) bursts of broadband (200-500 kHz) noise with a 'stripe' pattern of nulls imposed on it. These nulls have 10 to 20 kHz width and spacing, and many show a regular, non-linear frequency-time relation. These events are different from the 'stripes' discussed by Samara and LaBelle [2006] and Colpitts et al. [2010], because of the density of the stripes, the non-linearity, and the appearance of being an absorptive rather than emissive phenomenon. These events are similar to 'stripe' features reported by Brittain et al. [1983] in the VLF range, explained as an interference pattern between a downward-traveling whistler-mode wave and its reflection off the bottom of the ionosphere. Following their analysis method, we modeled our stripes as higher-frequency interfering whistlers reflecting off of a density gradient. This model predicts the near-hyperbolic frequency-time curves and high density of the nulls, and therefore shows promise at explaining the new observations.

Field of infrasound wave on the earth from blast wave, produced by supersonic flight of a rocket

International Nuclear Information System (INIS)

Drobzheva, Ya.V.; Krasnov, V.M.

2006-01-01

It was developed a physical model, which allowed calculating a field of infrasound wave on the earth from blast wave, produced by supersonic flight of a rocket. For space launching site Baikonur it is shown that the nearest horizontal distance from launching site of rocket up to which arrive infrasound waves, produced by supersonic flight of a rocket, is 56 km. Amplitude of acoustic impulse decreases in 5 times on distance of 600 km. Duration of acoustic impulse increases from 1.5 to 3 s on the same distance. Values of acoustic field parameters on the earth surface, practically, do not depend from season of launching of rocket. (author)

Observation and simulation of the ionosphere disturbance waves triggered by rocket exhausts

Science.gov (United States)

Lin, Charles C. H.; Chen, Chia-Hung; Matsumura, Mitsuru; Lin, Jia-Ting; Kakinami, Yoshihiro

2017-08-01

Observations and theoretical modeling of the ionospheric disturbance waves generated by rocket launches are investigated. During the rocket passage, time rate change of total electron content (rTEC) enhancement with the V-shape shock wave signature is commonly observed, followed by acoustic wave disturbances and region of negative rTEC centered along the trajectory. Ten to fifteen min after the rocket passage, delayed disturbance waves appeared and propagated along direction normal to the V-shape wavefronts. These observation features appeared most prominently in the 2016 North Korea rocket launch showing a very distinct V-shape rTEC enhancement over enormous areas along the southeast flight trajectory despite that it was also appeared in the 2009 North Korea rocket launch with the eastward flight trajectory. Numerical simulations using the physical-based nonlinear and nonhydrostatic coupled model of neutral atmosphere and ionosphere reproduce promised results in qualitative agreement with the characteristics of ionospheric disturbance waves observed in the 2009 event by considering the released energy of the rocket exhaust as the disturbance source. Simulations reproduce the shock wave signature of electron density enhancement, acoustic wave disturbances, the electron density depletion due to the rocket-induced pressure bulge, and the delayed disturbance waves. The pressure bulge results in outward neutral wind flows carrying neutrals and plasma away from it and leading to electron density depletions. Simulations further show, for the first time, that the delayed disturbance waves are produced by the surface reflection of the earlier arrival acoustic wave disturbances.

Ionospheric hole made by a North Korean rocket launched in 2012 December: Observation with the Russian GNSS

Science.gov (United States)

Nakashima, Y.; Heki, K.

2013-12-01

The Unha-3 rocket was launched due southward at 00:49:46UT on Dec. 12, 2012, from the Tongchang-ri.launch pad on the Yellow Sea side of North Korea. We converted the RINEX format GPS data of the launch day to TEC, and looked for the ionospheric hole signatures. We could not find clear electron depletion signals simply because no GPS satellites were available in the northwestern skies. GPS is the American GNSS system, and other systems are becoming operational. GEONET receivers have been replaced with the new models capable of receiving multiple GNSS, and about 10 percent of them could observe GLONASS and QZSS, the Russian and the Japanese GNSS, respectively, at the time of the Unha-3 launch. More than 20 GLONASS satellites are already in operation, and we used the number 13 satellite to detect the ionospheric hole formation above the Yellow Sea (see Figure). We modified the software to convert RINEX file into TEC time series [Heki et al., JGSJ 2010] in order to handle RINEX v.2.12 files including GLONASS/QZSS data. The broadcast orbits of the GLONASS satellites are given in the geocentric Cartesian coordinates instead of the Keplerian elements like GPS and QZSS. GLONASS uses different microwave frequencies for different satellites, which also required the modification for the original software to calculate TEC. Ozeki & Heki [2010] compared the thrust of the 1998 and 2009 Taepodong missiles by comparing the sizes/depths of the ionospheric holes, and here we compare the hole made by the 2012 December Unha-3 launch with the past cases. The onset times of the depletion are the same, suggesting similar ascending speeds of the three rockets (missiles). Depth of the hole depends both on the amount of water vapor in the exhaust and the background TEC. The hole of the Unha-3 is similar to the 2009 case (or somewhat deeper/larger), which would reflect the vertical TEC in the 2012 case about 1/3 larger than that in 2009. The hole seems to last longer in the 2012 case possibly

Rockets two classic papers

CERN Document Server

Goddard, Robert

2002-01-01

Rockets, in the primitive form of fireworks, have existed since the Chinese invented them around the thirteenth century. But it was the work of American Robert Hutchings Goddard (1882-1945) and his development of liquid-fueled rockets that first produced a controlled rocket flight. Fascinated by rocketry since boyhood, Goddard designed, built, and launched the world's first liquid-fueled rocket in 1926. Ridiculed by the press for suggesting that rockets could be flown to the moon, he continued his experiments, supported partly by the Smithsonian Institution and defended by Charles Lindbergh. T

Safety Practices Followed in ISRO Launch Complex- An Overview

Science.gov (United States)

Krishnamurty, V.; Srivastava, V. K.; Ramesh, M.

2005-12-01

The spaceport of India, Satish Dhawan Space Centre (SDSC) SHAR of Indian Space Research Organisation (ISRO), is located at Sriharikota, a spindle shaped island on the east coast of southern India.SDSC SHAR has a unique combination of facilities, such as a solid propellant production plant, a rocket motor static test facility, launch complexes for different types of rockets, telemetry, telecommand, tracking, data acquisition and processing facilities and other support services.The Solid Propellant Space Booster Plant (SPROB) located at SDSC SHAR produces composite solid propellant for rocket motors of ISRO. The main ingredients of the propellant produced here are ammonium perchlorate (oxidizer), fine aluminium powder (fuel) and hydroxyl terminated polybutadiene (binder).SDSC SHAR has facilities for testing solid rocket motors, both at ambient conditions and at simulated high altitude conditions. Other test facilities for the environmental testing of rocket motors and their subsystems include Vibration, Shock, Constant Acceleration and Thermal / Humidity.SDSC SHAR has the necessary infrastructure for launching satellites into low earth orbit, polar orbit and geo-stationary transfer orbit. The launch complexes provide complete support for vehicle assembly, fuelling with both earth storable and cryogenic propellants, checkout and launch operations. Apart from these, it has facilities for launching sounding rockets for studying the Earth's upper atmosphere and for controlled reentry and recovery of ISRO's space capsule reentry missions.Safety plays a major role at SDSC SHAR right from the mission / facility design phase to post launch operations. This paper presents briefly the infrastructure available at SDSC SHAR of ISRO for launching sounding rockets, satellite launch vehicles, controlled reentry missions and the built in safety systems. The range safety methodology followed as a part of the real time mission monitoring is presented. The built in safety systems

Low-Cost Propellant Launch to Earth Orbit from a Tethered Balloon

Science.gov (United States)

Wilcox, Brian H.

2006-01-01

Propellant will be more than 85% of the mass that needs to be lofted into Low Earth Orbit (LEO) in the planned program of Exploration of the Moon, Mars, and beyond. This paper describes a possible means for launching thousands of tons of propellant per year into LEO at a cost 15 to 30 times less than the current launch cost per kilogram. The basic idea is to mass-produce very simple, small and relatively low-performance rockets at a cost per kilogram comparable to automobiles, instead of the 25X greater cost that is customary for current launch vehicles that are produced in small quantities and which are manufactured with performance near the limits of what is possible. These small, simple rockets can reach orbit because they are launched above 95% of the atmosphere, where the drag losses even on a small rocket are acceptable, and because they can be launched nearly horizontally with very simple guidance based primarily on spin-stabilization. Launching above most of the atmosphere is accomplished by winching the rocket up a tether to a balloon. A fuel depot in equatorial orbit passes over the launch site on every orbit (approximately every 90 minutes). One or more rockets can be launched each time the fuel depot passes overhead, so the launch rate can be any multiple of 6000 small rockets per year, a number that is sufficient to reap the benefits of mass production.

Investigation Regarding Assertions Made by Former United Launch Alliance Executive (REDACTED)

Science.gov (United States)

2016-12-05

reviewed documents and contract awards. First, did ULA improperly transfer five RD-180 rocket engines from NSS launch missions to commercial ...distinguish between commercial and government missions and uses a “first in, first out” asset management method for its RD-180 rocket engines. The...improperly transfer five RD-180 rocket engines from NSS launch missions to commercial launch missions to influence congressional legislation? 2. Did

Hybrid rocket engine, theoretical model and experiment

Science.gov (United States)

Chelaru, Teodor-Viorel; Mingireanu, Florin

2011-06-01

The purpose of this paper is to build a theoretical model for the hybrid rocket engine/motor and to validate it using experimental results. The work approaches the main problems of the hybrid motor: the scalability, the stability/controllability of the operating parameters and the increasing of the solid fuel regression rate. At first, we focus on theoretical models for hybrid rocket motor and compare the results with already available experimental data from various research groups. A primary computation model is presented together with results from a numerical algorithm based on a computational model. We present theoretical predictions for several commercial hybrid rocket motors, having different scales and compare them with experimental measurements of those hybrid rocket motors. Next the paper focuses on tribrid rocket motor concept, which by supplementary liquid fuel injection can improve the thrust controllability. A complementary computation model is also presented to estimate regression rate increase of solid fuel doped with oxidizer. Finally, the stability of the hybrid rocket motor is investigated using Liapunov theory. Stability coefficients obtained are dependent on burning parameters while the stability and command matrixes are identified. The paper presents thoroughly the input data of the model, which ensures the reproducibility of the numerical results by independent researchers.

Improving of technical characteristics of launch vehicles with liquid rocket engines using active onboard de-orbiting systems

Science.gov (United States)

Trushlyakov, V.; Shatrov, Ya.

2017-09-01

In this paper, the analysis of technical requirements (TR) for the development of modern space launch vehicles (LV) with main liquid rocket engines (LRE) is fulfilled in relation to the anthropogenic impact decreasing. Factual technical characteristics on the example of a promising type of rocket ;Soyuz-2.1.v.; are analyzed. Meeting the TR in relation to anthropogenic impact decrease based on the conventional design approach and the content of the onboard system does not prove to be efficient and leads to depreciation of the initial technical characteristics obtained at the first design stage if these requirements are not included. In this concern, it is shown that the implementation of additional active onboard de-orbiting system (AODS) of worked-off stages (WS) into the onboard LV stages systems allows to meet the TR related to the LV environmental characteristics, including fire-explosion safety. In some cases, the orbital payload mass increases.

Development and Performance of the 10 kN Hybrid Rocket Motor for the Stratos II Sounding Rocket

NARCIS (Netherlands)

Werner, R.M.; Knop, T.R.; Wink, J; Ehlen, J; Huijsman, R; Powell, S; Florea, R.; Wieling, W; Cervone, A.; Zandbergen, B.T.C.

2016-01-01

This paper presents the development work of the 10 kN hybrid rocket motor DHX-200 Aurora. The DHX-200 Aurora was developed by Delft Aerospace Rocket Engineering (DARE) to power the Stratos II and Stratos II+ sounding rocket, with the later one being launched in October 2015. Stratos II and Stratos

Hyper-X Research Vehicle - Artist Concept Mounted on Pegasus Rocket Attached to B-52 Launch Aircraft

Science.gov (United States)

1997-01-01

This artist's concept depicts the Hyper-X research vehicle riding on a booster rocket prior to being launched by the Dryden Flight Research Center's B-52 at about 40,000 feet. 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

Measuring Model Rocket Engine Thrust Curves

Science.gov (United States)

Penn, Kim; Slaton, William V.

2010-01-01

This paper describes a method and setup to quickly and easily measure a model rocket engine's thrust curve using a computer data logger and force probe. Horst describes using Vernier's LabPro and force probe to measure the rocket engine's thrust curve; however, the method of attaching the rocket to the force probe is not discussed. We show how a…

Ionospheric shock waves triggered by rockets

Directory of Open Access Journals (Sweden)

C. H. Lin

2014-09-01

Full Text Available This paper presents a two-dimensional structure of the shock wave signatures in ionospheric electron density resulting from a rocket transit using the rate of change of the total electron content (TEC derived from ground-based GPS receivers around Japan and Taiwan for the first time. From the TEC maps constructed for the 2009 North Korea (NK Taepodong-2 and 2013 South Korea (SK Korea Space Launch Vehicle-II (KSLV-II rocket launches, features of the V-shaped shock wave fronts in TEC perturbations are prominently seen. These fronts, with periods of 100–600 s, produced by the propulsive blasts of the rockets appear immediately and then propagate perpendicularly outward from the rocket trajectory with supersonic velocities between 800–1200 m s−1 for both events. Additionally, clear rocket exhaust depletions of TECs are seen along the trajectory and are deflected by the background thermospheric neutral wind. Twenty minutes after the rocket transits, delayed electron density perturbation waves propagating along the bow wave direction appear with phase velocities of 800–1200 m s−1. According to their propagation character, these delayed waves may be generated by rocket exhaust plumes at earlier rocket locations at lower altitudes.

Liquid Rocket Engine Testing Overview

Science.gov (United States)

Rahman, Shamim

2005-01-01

Contents include the following: Objectives and motivation for testing. Technology, Research and Development Test and Evaluation (RDT&E), evolutionary. Representative Liquid Rocket Engine (LRE) test compaigns. Apollo, shuttle, Expandable Launch Vehicles (ELV) propulsion. Overview of test facilities for liquid rocket engines. Boost, upper stage (sea-level and altitude). Statistics (historical) of Liquid Rocket Engine Testing. LOX/LH, LOX/RP, other development. Test project enablers: engineering tools, operations, processes, infrastructure.

Reusable Military Launch Systems (RMLS)

Science.gov (United States)

2008-02-01

shown in Figure 11. The second configuration is an axisymmetric, rocket-based combined cycle (RBCC) powered, SSTO vehicle, similar to the GTX...McCormick, D., and Sorensen, K., “Hyperion: An SSTO Vision Vehicle Concept Utilizing Rocket-Based Combined Cycle Propulsion”, AIAA paper 99-4944...there have been several failedattempts at the development of reusable rocket or air-breathing launch vehicle systems. Single-stage-to-orbit ( SSTO

Potential climate impact of black carbon emitted by rockets

Science.gov (United States)

Ross, Martin; Mills, Michael; Toohey, Darin

2010-12-01

A new type of hydrocarbon rocket engine is expected to power a fleet of suborbital rockets for commercial and scientific purposes in coming decades. A global climate model predicts that emissions from a fleet of 1000 launches per year of suborbital rockets would create a persistent layer of black carbon particles in the northern stratosphere that could cause potentially significant changes in the global atmospheric circulation and distributions of ozone and temperature. Tropical stratospheric ozone abundances are predicted to change as much as 1%, while polar ozone changes by up to 6%. Polar surface temperatures change as much as one degree K regionally with significant impacts on polar sea ice fractions. After one decade of continuous launches, globally averaged radiative forcing from the black carbon would exceed the forcing from the emitted CO2 by a factor of about 105 and would be comparable to the radiative forcing estimated from current subsonic aviation.

Two-dimensional motions of rockets

International Nuclear Information System (INIS)

Kang, Yoonhwan; Bae, Saebyok

2007-01-01

We analyse the two-dimensional motions of the rockets for various types of rocket thrusts, the air friction and the gravitation by using a suitable representation of the rocket equation and the numerical calculation. The slope shapes of the rocket trajectories are discussed for the three types of rocket engines. Unlike the projectile motions, the descending parts of the trajectories tend to be gentler and straighter slopes than the ascending parts for relatively large launching angles due to the non-vanishing thrusts. We discuss the ranges, the maximum altitudes and the engine performances of the rockets. It seems that the exponential fuel exhaustion can be the most potent engine for the longest and highest flights

MSFC Advanced Concepts Office and the Iterative Launch Vehicle Concept Method

Science.gov (United States)

Creech, Dennis

2011-01-01

This slide presentation reviews the work of the Advanced Concepts Office (ACO) at Marshall Space Flight Center (MSFC) with particular emphasis on the method used to model launch vehicles using INTegrated ROcket Sizing (INTROS), a modeling system that assists in establishing the launch concept design, and stage sizing, and facilitates the integration of exterior analytic efforts, vehicle architecture studies, and technology and system trades and parameter sensitivities.

Rocket + Science = Dialogue

Science.gov (United States)

Morris,Bruce; Sullivan, Greg; Burkey, Martin

2010-01-01

It's a cliche that rocket engineers and space scientists don t see eye-to-eye. That goes double for rocket engineers working on human spaceflight and scientists working on space telescopes and planetary probes. They work fundamentally different problems but often feel that they are competing for the same pot of money. Put the two groups together for a weekend, and the results could be unscientific or perhaps combustible. Fortunately, that wasn't the case when NASA put heavy lift launch vehicle designers together with astronomers and planetary scientists for two weekend workshops in 2008. The goal was to bring the top people from both groups together to see how the mass and volume capabilities of NASA's Ares V heavy lift launch vehicle could benefit the science community. Ares V is part of NASA's Constellation Program for resuming human exploration beyond low Earth orbit, starting with missions to the Moon. In the current mission scenario, Ares V launches a lunar lander into Earth orbit. A smaller Ares I rocket launches the Orion crew vehicle with up to four astronauts. Orion docks with the lander, attached to the Ares V Earth departure stage. The stage fires its engine to send the mated spacecraft to the Moon. Standing 360 feet high and weighing 7.4 million pounds, NASA's new heavy lifter will be bigger than the 1960s-era Saturn V. It can launch almost 60 percent more payload to translunar insertion together with the Ares I and 35 percent more mass to low Earth orbit than the Saturn V. This super-sized capability is, in short, designed to send more people to more places to do more things than the six Apollo missions.

Initial risk assessment for a single stage to orbit nuclear thermal rocket

Energy Technology Data Exchange (ETDEWEB)

Labib, Satira, E-mail: Satira.Labib@duke-energy.com; King, Jeffrey, E-mail: kingjc@mines.edu

2015-06-15

Highlights: • The risks posed by the surface launch of a nuclear thermal rocket are considered. • Radiation exposure at the public viewing distance is insignificant. • Production of fission products and actinides during launch is limited. • The production of activated argon around the rocket may be a significant concern. - Abstract: In order to consider the possibility of a nuclear thermal rocket (NTR) ground launch, it is necessary to evaluate the risks from such a launch. This includes analysis of the radiation dose rate around the rocket, determining the rate of activation of the materials near the launch, and considering the radionuclides present in the core after the launch. This paper evaluates the potential risk of the NTR ground launch for a range of payloads from 1 to 15 metric tons (MT) using three NTR reactor cores (40, 80, and 120 cm in length) designed in a previous study, based on data produced by MCNP5 and MCNPX models. At the same power level, the 40 cm core length reactor results in the lowest radiation dose rate of the three reactors. Radiation dose rates decrease to background levels 3.5 km from the launch site. After a 1-year decay time, all of the activated materials produced by an NTR launch would be classified as Class A low-level waste. The activation of air produces significant amounts of argon-41 and nitrogen-16 within 100 m of the launch. The derived air concentration (DAC) ratio of the activation products decays to less than unity within 2 days, with only argon-41 remaining. After 10 min of full power operation, the 120 cm core for a 15 MT payload contains 2.5 × 10{sup 13}, 1.4 × 10{sup 12} and 1.5 × 10{sup 12} Bq of {sup 131}I, {sup 137}Cs, and {sup 90}Sr, respectively. The decay heat after shutdown increases with increasing reactor power with a maximum decay heat of 108 kW immediately after shutdown for the 15 MT payload.

The NASA Sounding Rocket Program and space sciences

Science.gov (United States)

Gurkin, L. W.

1992-01-01

High altitude suborbital rockets (sounding rockets) have been extensively used for space science research in the post-World War II period; the NASA Sounding Rocket Program has been on-going since the inception of the Agency and supports all space science disciplines. In recent years, sounding rockets have been utilized to provide a low gravity environment for materials processing research, particularly in the commercial sector. Sounding rockets offer unique features as a low gravity flight platform. Quick response and low cost combine to provide more frequent spaceflight opportunities. Suborbital spacecraft design practice has achieved a high level of sophistication which optimizes the limited available flight times. High data-rate telemetry, real-time ground up-link command and down-link video data are routinely used in sounding rocket payloads. Standard, off-the-shelf, active control systems are available which limit payload body rates such that the gravitational environment remains less than 10(-4) g during the control period. Operational launch vehicles are available which can provide up to 7 minutes of experiment time for experiment weights up to 270 kg. Standard payload recovery systems allow soft impact retrieval of payloads. When launched from White Sands Missile Range, New Mexico, payloads can be retrieved and returned to the launch site within hours.

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.

Modeling of Mutiscale Electromagnetic Magnetosphere-Ionosphere Interactions near Discrete Auroral Arcs Observed by the MICA Sounding Rocket

Science.gov (United States)

Streltsov, A. V.; Lynch, K. A.; Fernandes, P. A.; Miceli, R.; Hampton, D. L.; Michell, R. G.; Samara, M.

2012-12-01

The MICA (Magnetosphere-Ionosphere Coupling in the Alfvén Resonator) sounding rocket was launched from Poker Flat on February 19, 2012. The rocket was aimed into the system of discrete auroral arcs and during its flight it detected small-scale electromagnetic disturbances with characteristic features of dispersive Alfvén waves. We report results from numerical modeling of these observations. Our simulations are based on a two-fluid MHD model describing multi-scale interactions between magnetic field-aligned currents carried by shear Alfven waves and the ionosphere. The results from our simulations suggest that the small-scale electromagnetic structures measured by MICA indeed can be interpreted as dispersive Alfvén waves generated by the active ionospheric response (ionopspheric feedback instability) inside the large-scale downward magnetic field-aligned current interacting with the ionosphere.

South Pole rockets, (1)

International Nuclear Information System (INIS)

Kimura, Iwane

1977-01-01

Wave-particle interaction was observed, using three rockets, S-210 JA-20, -21 and S-310 JA-2, launched from the South Pole into aurora. Electron density and temperature were measured with these rockets. Simultaneous observations of waves were also made from a satellite (ISIS-II) and at two ground bases (Showa base and Mizuho base). Observed data are presented in this paper. These include electron density and temperature in relation to altitude; variation of electron (60 - 80 keV) count rate with altitude; VLF spectra measured by the PWL of S-210 JA-20 and -21 rockets and the corresponding VLF spectra at the ground bases; low-energy (<10 keV) electron flux measured by S-310 JA-2 rocket; and VLF spectrum measured with S-310 JA-2 rocket. Scheduled measurements for the next project are also briefly described. (Aoki, K.)

Focused RBCC Experiments: Two-Rocket Configuration Experiments and Hydrocarbon/Oxygen Rocket Ejector Experiments

Science.gov (United States)

Santoro, Robert J.; Pal, Sibtosh

2003-01-01

This addendum report documents the results of two additional efforts for the Rocket Based Combined Cycle (RBCC) rocket-ejector mode research work carried out at the Penn State Propulsion Engineering Research Center in support of NASA s technology development efforts for enabling 3 d generation Reusable Launch Vehicles (RLV). The tasks reported here build on an earlier NASA MSFC funded research program on rocket ejector investigations. The first task investigated the improvements of a gaseous hydrogen/oxygen twin thruster RBCC rocket ejector system over a single rocket system. The second task investigated the performance of a hydrocarbon (liquid JP-7)/gaseous oxygen single thruster rocket-ejector system. To gain a systematic understanding of the rocket-ejector s internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static diffusion and afterburning (DAB) configurations for a range of rocket operating conditions. For all experimental conditions, overall system performance was obtained through global measurements of wall static pressure profiles, heat flux profiles and engine thrust. Detailed mixing and combustion information was obtained through Raman spectroscopy measurements of major species (gaseous oxygen, hydrogen, nitrogen and water vapor) for the gaseous hydrogen/oxygen rocket ejector experiments.

Rocket observations

Science.gov (United States)

1984-05-01

The Institute of Space and Astronautical Science (ISAS) sounding rocket experiments were carried out during the periods of August to September, 1982, January to February and August to September, 1983 and January to February, 1984 with sounding rockets. Among 9 rockets, 3 were K-9M, 1 was S-210, 3 were S-310 and 2 were S-520. Two scientific satellites were launched on February 20, 1983 for solar physics and on February 14, 1984 for X-ray astronomy. These satellites were named as TENMA and OHZORA and designated as 1983-011A and 1984-015A, respectively. Their initial orbital elements are also described. A payload recovery was successfully carried out by S-520-6 rocket as a part of MINIX (Microwave Ionosphere Non-linear Interaction Experiment) which is a scientific study of nonlinear plasma phenomena in conjunction with the environmental assessment study for the future SPS project. Near IR observation of the background sky shows a more intense flux than expected possibly coming from some extragalactic origin and this may be related to the evolution of the universe. US-Japan cooperative program of Tether Experiment was done on board US rocket.

Technology Innovations from NASA's Next Generation Launch Technology Program

Science.gov (United States)

Cook, Stephen A.; Morris, Charles E. K., Jr.; Tyson, Richard W.

2004-01-01

NASA's Next Generation Launch Technology Program has been on the cutting edge of technology, improving the safety, affordability, and reliability of future space-launch-transportation systems. The array of projects focused on propulsion, airframe, and other vehicle systems. Achievements range from building miniature fuel/oxygen sensors to hot-firings of major rocket-engine systems as well as extreme thermo-mechanical testing of large-scale structures. Results to date have significantly advanced technology readiness for future space-launch systems using either airbreathing or rocket propulsion.

Study on Alternative Cargo Launch Options from the Lunar Surface

Energy Technology Data Exchange (ETDEWEB)

Cheryl A. Blomberg; Zamir A. Zulkefli; Spencer W. Rich; Steven D. Howe

2013-07-01

In the future, there will be a need for constant cargo launches from Earth to Mars in order to build, and then sustain, a Martian base. Currently, chemical rockets are used for space launches. These are expensive and heavy due to the amount of necessary propellant. Nuclear thermal rockets (NTRs) are the next step in rocket design. Another alternative is to create a launcher on the lunar surface that uses magnetic levitation to launch cargo to Mars in order to minimize the amount of necessary propellant per mission. This paper investigates using nuclear power for six different cargo launching alternatives, as well as the orbital mechanics involved in launching cargo to a Martian base from the moon. Each alternative is compared to the other alternative launchers, as well as compared to using an NTR instead. This comparison is done on the basis of mass that must be shipped from Earth, the amount of necessary propellant, and the number of equivalent NTR launches. Of the options, a lunar coil launcher had a ship mass that is 12.7% less than the next best option and 17 NTR equivalent launches, making it the best of the presented six options.

14 CFR 437.95 - Inspection of additional reusable suborbital rockets.

Science.gov (United States)

2010-01-01

... suborbital rockets. 437.95 Section 437.95 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL... of an Experimental Permit § 437.95 Inspection of additional reusable suborbital rockets. A permittee may launch or reenter additional reusable suborbital rockets of the same design under the permit after...

Significant Climate Changes Caused by Soot Emitted From Rockets in the Stratosphere

Science.gov (United States)

Mills, M. J.; Ross, M.; Toohey, D. W.

2010-12-01

A new type of hydrocarbon rocket engine with a larger soot emission index than current kerosene rockets is expected to power a fleet of suborbital rockets for commercial and scientific purposes in coming decades. At projected launch rates, emissions from these rockets will create a persistent soot layer in the northern middle stratosphere that would disproportionally affect the Earth’s atmosphere and cryosphere. A global climate model predicts that thermal forcing in the rocket soot layer will cause significant changes in the global atmospheric circulation and distributions of ozone and temperature. Tropical ozone columns decline as much as 1%, while polar ozone columns increase by up to 6%. Polar surface temperatures rise one Kelvin regionally and polar summer sea ice fractions shrink between 5 - 15%. After 20 years of suborbital rocket fleet operation, globally averaged radiative forcing (RF) from rocket soot exceeds the RF from rocket CO_{2} by six orders of magnitude, but remains small, comparable to the global RF from aviation. The response of the climate system is surprising given the small forcing, and should be investigated further with different climate models.

The Norwegian sounding rocket programme 1978-81

International Nuclear Information System (INIS)

Landmark, B.

1978-01-01

The Norwegian sounding rocket programme is reasonably well defined up to and including the winter of 1981/82. All the projects have been planned and will be carried out in international cooperation. Norwegian scientists so far plan to participate in a number of 24 rocket payloads over the period. Out of these 18 will be launched from the Andoya rocket range, 3 from Esrange and 3 from the siple station in the antarctic. (author)

Ozone Depletion Caused by Rocket Engine Emissions: A Fundamental Limit on the Scale and Viability of Space-Based Geoengineering Schemes

Science.gov (United States)

Ross, M. N.; Toohey, D.

2008-12-01

Emissions from solid and liquid propellant rocket engines reduce global stratospheric ozone levels. Currently ~ one kiloton of payloads are launched into earth orbit annually by the global space industry. Stratospheric ozone depletion from present day launches is a small fraction of the ~ 4% globally averaged ozone loss caused by halogen gases. Thus rocket engine emissions are currently considered a minor, if poorly understood, contributor to ozone depletion. Proposed space-based geoengineering projects designed to mitigate climate change would require order of magnitude increases in the amount of material launched into earth orbit. The increased launches would result in comparable increases in the global ozone depletion caused by rocket emissions. We estimate global ozone loss caused by three space-based geoengineering proposals to mitigate climate change: (1) mirrors, (2) sunshade, and (3) space-based solar power (SSP). The SSP concept does not directly engineer climate, but is touted as a mitigation strategy in that SSP would reduce CO2 emissions. We show that launching the mirrors or sunshade would cause global ozone loss between 2% and 20%. Ozone loss associated with an economically viable SSP system would be at least 0.4% and possibly as large as 3%. It is not clear which, if any, of these levels of ozone loss would be acceptable under the Montreal Protocol. The large uncertainties are mainly caused by a lack of data or validated models regarding liquid propellant rocket engine emissions. Our results offer four main conclusions. (1) The viability of space-based geoengineering schemes could well be undermined by the relatively large ozone depletion that would be caused by the required rocket launches. (2) Analysis of space- based geoengineering schemes should include the difficult tradeoff between the gain of long-term (~ decades) climate control and the loss of short-term (~ years) deep ozone loss. (3) The trade can be properly evaluated only if our

Subsonic Glideback Rocket Demonstrator Flight Testing

Science.gov (United States)

DeTurris, Dianne J.; Foster, Trevor J.; Barthel, Paul E.; Macy, Daniel J.; Droney, Christopher K.; Talay, Theodore A. (Technical Monitor)

2001-01-01

For the past two years, Cal Poly's rocket program has been aggressively exploring the concept of remotely controlled, fixed wing, flyable rocket boosters. This program, embodied by a group of student engineers known as Cal Poly Space Systems, has successfully demonstrated the idea of a rocket design that incorporates a vertical launch pattern followed by a horizontal return flight and landing. Though the design is meant for supersonic flight, CPSS demonstrators are deployed at a subsonic speed. Many steps have been taken by the club that allowed the evolution of the StarBooster prototype to reach its current size: a ten-foot tall, one-foot diameter, composite material rocket. Progress is currently being made that involves multiple boosters along with a second stage, third rocket.

Experimental validation of solid rocket motor damping models

Science.gov (United States)

Riso, Cristina; Fransen, Sebastiaan; Mastroddi, Franco; Coppotelli, Giuliano; Trequattrini, Francesco; De Vivo, Alessio

2017-12-01

In design and certification of spacecraft, payload/launcher coupled load analyses are performed to simulate the satellite dynamic environment. To obtain accurate predictions, the system damping properties must be properly taken into account in the finite element model used for coupled load analysis. This is typically done using a structural damping characterization in the frequency domain, which is not applicable in the time domain. Therefore, the structural damping matrix of the system must be converted into an equivalent viscous damping matrix when a transient coupled load analysis is performed. This paper focuses on the validation of equivalent viscous damping methods for dynamically condensed finite element models via correlation with experimental data for a realistic structure representative of a slender launch vehicle with solid rocket motors. A second scope of the paper is to investigate how to conveniently choose a single combination of Young's modulus and structural damping coefficient—complex Young's modulus—to approximate the viscoelastic behavior of a solid propellant material in the frequency band of interest for coupled load analysis. A scaled-down test article inspired to the Z9-ignition Vega launcher configuration is designed, manufactured, and experimentally tested to obtain data for validation of the equivalent viscous damping methods. The Z9-like component of the test article is filled with a viscoelastic material representative of the Z9 solid propellant that is also preliminarily tested to investigate the dependency of the complex Young's modulus on the excitation frequency and provide data for the test article finite element model. Experimental results from seismic and shock tests performed on the test configuration are correlated with numerical results from frequency and time domain analyses carried out on its dynamically condensed finite element model to assess the applicability of different equivalent viscous damping methods to describe

Experimental validation of solid rocket motor damping models

Science.gov (United States)

Riso, Cristina; Fransen, Sebastiaan; Mastroddi, Franco; Coppotelli, Giuliano; Trequattrini, Francesco; De Vivo, Alessio

2018-06-01

In design and certification of spacecraft, payload/launcher coupled load analyses are performed to simulate the satellite dynamic environment. To obtain accurate predictions, the system damping properties must be properly taken into account in the finite element model used for coupled load analysis. This is typically done using a structural damping characterization in the frequency domain, which is not applicable in the time domain. Therefore, the structural damping matrix of the system must be converted into an equivalent viscous damping matrix when a transient coupled load analysis is performed. This paper focuses on the validation of equivalent viscous damping methods for dynamically condensed finite element models via correlation with experimental data for a realistic structure representative of a slender launch vehicle with solid rocket motors. A second scope of the paper is to investigate how to conveniently choose a single combination of Young's modulus and structural damping coefficient—complex Young's modulus—to approximate the viscoelastic behavior of a solid propellant material in the frequency band of interest for coupled load analysis. A scaled-down test article inspired to the Z9-ignition Vega launcher configuration is designed, manufactured, and experimentally tested to obtain data for validation of the equivalent viscous damping methods. The Z9-like component of the test article is filled with a viscoelastic material representative of the Z9 solid propellant that is also preliminarily tested to investigate the dependency of the complex Young's modulus on the excitation frequency and provide data for the test article finite element model. Experimental results from seismic and shock tests performed on the test configuration are correlated with numerical results from frequency and time domain analyses carried out on its dynamically condensed finite element model to assess the applicability of different equivalent viscous damping methods to describe

Internet Based Simulations of Debris Dispersion of Shuttle Launch

Science.gov (United States)

Bardina, Jorge; Thirumalainambi, Rajkumar

2004-01-01

The debris dispersion model (which dispersion model?) is so heterogeneous and interrelated with various factors, 3D graphics combined with physical models are useful in understanding the complexity of launch and range operations. Modeling and simulation in this area mainly focuses on orbital dynamics and range safety concepts, including destruct limits, telemetry and tracking, and population risk. Particle explosion modeling is the process of simulating an explosion by breaking the rocket into many pieces. The particles are scattered throughout their motion using the laws of physics eventually coming to rest. The size of the foot print explains the type of explosion and distribution of the particles. The shuttle launch and range operations in this paper are discussed based on the operations of the Kennedy Space Center, Florida, USA. Java 3D graphics provides geometric and visual content with suitable modeling behaviors of Shuttle launches.

The Application of the NASA Advanced Concepts Office, Launch Vehicle Team Design Process and Tools for Modeling Small Responsive Launch Vehicles

Science.gov (United States)

Threet, Grady E.; Waters, Eric D.; Creech, Dennis M.

2012-01-01

The Advanced Concepts Office (ACO) Launch Vehicle Team at the NASA Marshall Space Flight Center (MSFC) is recognized throughout NASA for launch vehicle conceptual definition and pre-phase A concept design evaluation. The Launch Vehicle Team has been instrumental in defining the vehicle trade space for many of NASA s high level launch system studies from the Exploration Systems Architecture Study (ESAS) through the Augustine Report, Constellation, and now Space Launch System (SLS). The Launch Vehicle Team s approach to rapid turn-around and comparative analysis of multiple launch vehicle architectures has played a large role in narrowing the design options for future vehicle development. Recently the Launch Vehicle Team has been developing versions of their vetted tools used on large launch vehicles and repackaged the process and capability to apply to smaller more responsive launch vehicles. Along this development path the LV Team has evaluated trajectory tools and assumptions against sounding rocket trajectories and air launch systems, begun altering subsystem mass estimating relationships to handle smaller vehicle components, and as an additional development driver, have begun an in-house small launch vehicle study. With the recent interest in small responsive launch systems and the known capability and response time of the ACO LV Team, ACO s launch vehicle assessment capability can be utilized to rapidly evaluate the vast and opportune trade space that small launch vehicles currently encompass. This would provide a great benefit to the customer in order to reduce that large trade space to a select few alternatives that should best fit the customer s payload needs.

National Report on the NASA Sounding Rocket and Balloon Programs

Science.gov (United States)

Eberspeaker, Philip; Fairbrother, Debora

2013-01-01

The U. S. National Aeronautics and Space Administration (NASA) Sounding Rockets and Balloon Programs conduct a total of 30 to 40 missions per year in support of the NASA scientific community and other users. The NASA Sounding Rockets Program supports the science community by integrating their experiments into the sounding rocket payloads, and providing both the rocket vehicle and launch operations services. Activities since 2011 have included two flights from Andoya Rocket Range, more than eight flights from White Sands Missile Range, approximately sixteen flights from Wallops Flight Facility, two flights from Poker Flat Research Range, and four flights from Kwajalein Atoll. Other activities included the final developmental flight of the Terrier-Improved Malemute launch vehicle, a test flight of the Talos-Terrier-Oriole launch vehicle, and a host of smaller activities to improve program support capabilities. Several operational missions have utilized the new Terrier-Malemute vehicle. The NASA Sounding Rockets Program is currently engaged in the development of a new sustainer motor known as the Peregrine. The Peregrine development effort will involve one static firing and three flight tests with a target completion data of August 2014. The NASA Balloon Program supported numerous scientific and developmental missions since its last report. The program conducted flights from the U.S., Sweden, Australia, and Antarctica utilizing standard and experimental vehicles. Of particular note are the successful test flights of the Wallops Arc Second Pointer (WASP), the successful demonstration of a medium-size Super Pressure Balloon (SPB), and most recently, three simultaneous missions aloft over Antarctica. NASA continues its successful incremental design qualification program and will support a science mission aboard WASP in late 2013 and a science mission aboard the SPB in early 2015. NASA has also embarked on an intra-agency collaboration to launch a rocket from a balloon to

National Security Space Launch at a Crossroads

Science.gov (United States)

2016-05-13

appropriations measure expire (i.e., at the start of FY2017). Rocket Engines: Goods or Services37 In the Commercial Space Act of 1998 (CSA),38...procured from commercial sources or whether the government may independently develop and manufacture rocket engines. The resolution to this question may...Space Act of 1998 ...) the DoD procures commercial launch services rather than rockets or engines used in those services.”43 Notably, the view

Space Launch System Base Heating Test: Environments and Base Flow Physics

Science.gov (United States)

Mehta, Manish; Knox, Kyle S.; Seaford, C. Mark; Dufrene, Aaron T.

2016-01-01

The NASA Space Launch System (SLS) vehicle is composed of four RS-25 liquid oxygen- hydrogen rocket engines in the core-stage and two 5-segment solid rocket boosters and as a result six hot supersonic plumes interact within the aft section of the vehicle during ight. Due to the complex nature of rocket plume-induced ows within the launch vehicle base during ascent and a new vehicle con guration, sub-scale wind tunnel testing is required to reduce SLS base convective environment uncertainty and design risk levels. This hot- re test program was conducted at the CUBRC Large Energy National Shock (LENS) II short-duration test facility to simulate ight from altitudes of 50 kft to 210 kft. The test program is a challenging and innovative e ort that has not been attempted in 40+ years for a NASA vehicle. This presentation discusses the various trends of base convective heat ux and pressure as a function of altitude at various locations within the core-stage and booster base regions of the two-percent SLS wind tunnel model. In-depth understanding of the base ow physics is presented using the test data, infrared high-speed imaging and theory. The normalized test design environments are compared to various NASA semi- empirical numerical models to determine exceedance and conservatism of the ight scaled test-derived base design environments. Brief discussion of thermal impact to the launch vehicle base components is also presented.

Optimal Reference Strain Structure for Studying Dynamic Responses of Flexible Rockets

Science.gov (United States)

Tsushima, Natsuki; Su, Weihua; Wolf, Michael G.; Griffin, Edwin D.; Dumoulin, Marie P.

2017-01-01

In the proposed paper, the optimal design of reference strain structures (RSS) will be performed targeting for the accurate observation of the dynamic bending and torsion deformation of a flexible rocket. It will provide the detailed description of the finite-element (FE) model of a notional flexible rocket created in MSC.Patran. The RSS will be attached longitudinally along the side of the rocket and to track the deformation of the thin-walled structure under external loads. An integrated surrogate-based multi-objective optimization approach will be developed to find the optimal design of the RSS using the FE model. The Kriging method will be used to construct the surrogate model. For the data sampling and the performance evaluation, static/transient analyses will be performed with MSC.Natran/Patran. The multi-objective optimization will be solved with NSGA-II to minimize the difference between the strains of the launch vehicle and RSS. Finally, the performance of the optimal RSS will be evaluated by checking its strain-tracking capability in different numerical simulations of the flexible rocket.

Simple-1: Development stage of the data transmission system for a solid propellant mid-power rocket model

Science.gov (United States)

Yarce, Andrés; Sebastián Rodríguez, Juan; Galvez, Julián; Gómez, Alejandro; García, Manuel J.

2017-06-01

This paper presents the development stage of a communication module for a solid propellant mid-power rocket model. The communication module was named. Simple-1 and this work considers its design, construction and testing. A rocket model Estes Ventris Series Pro II® was modified to introduce, on the top of the payload, several sensors in a CanSat form factor. The Printed Circuit Board (PCB) was designed and fabricated from Commercial Off The Shelf (COTS) components and assembled in a cylindrical rack structure similar to this small format satellite concept. The sensors data was processed using one Arduino Mini and transmitted using a radio module to a Software Defined Radio (SDR) HackRF based platform on the ground station. The Simple-1 was tested using a drone in successive releases, reaching altitudes from 200 to 300 meters. Different kind of data, in terms of altitude, position, atmospheric pressure and vehicle temperature were successfully measured, making possible the progress to a next stage of launching and analysis.

Performance of a RBCC Engine in Rocket-Operation

Science.gov (United States)

Tomioka, Sadatake; Kubo, Takahiro; Noboru Sakuranaka; Tani, Koichiro

Combination of a scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines (the combined system termed as Rocket-based Combined Cycle engine) are expected to be the most effective propulsion system for space launch vehicles. Either SSTO (Single Stage To Orbit) system or TSTO (Two Stage To Orbit) system with separation at high altitude needs final stage acceleration in space, so that the RBCC (Rocket Based Combined Cycle) engine should be operated as rocket engines. Performance of the scramjet combustor as the extension to the rocket nozzle, was experimentally evaluated by injecting inert gas at various pressure through the embedded rocket chamber while the whole sub-scaled model was placed in a low pressure chamber connected to an air-driven ejector system. The results showed that the thrust coefficient was about 1.2, the low value being found to mainly due to the friction force on the scramjet combustor wall, while blocking the scramjet flow pass’s opening to increase nozzle extension thrust surface, was found to have little effects on the thrust performance. The combustor was shortened to reduce the friction loss, however, degree of reduction was limited as friction decreased rapidly with distance from the onset of the scramjet combustor.

Development and Short-Range Testing of a 100 kW Side-Illuminated Millimeter-Wave Thermal Rocket

Science.gov (United States)

Bruccoleri, Alexander; Eilers, James A.; Lambot, Thomas; Parkin, Kevin

2015-01-01

The objective of the phase described here of the Millimeter-Wave Thermal Launch System (MTLS) Project was to launch a small thermal rocket into the air using millimeter waves. The preliminary results of the first MTLS flight vehicle launches are presented in this work. The design and construction of a small thermal rocket with a planar ceramic heat exchanger mounted along the axis of the rocket is described. The heat exchanger was illuminated from the side by a millimeter-wave beam and fed propellant from above via a small tank containing high pressure argon or nitrogen. Short-range tests where the rocket was launched, tracked, and heated with the beam are described. The rockets were approximately 1.5 meters in length and 65 millimeters in diameter, with a liftoff mass of 1.8 kilograms. The rocket airframes were coated in aluminum and had a parachute recovery system activated via a timer and Pyrodex. At the rocket heat exchanger, the beam distance was 40 meters with a peak power intensity of 77 watts per square centimeter. and a total power of 32 kilowatts in a 30 centimeter diameter circle. An altitude of approximately 10 meters was achieved. Recommendations for improvements are discussed.

NASA Space Launch System Operations Outlook

Science.gov (United States)

Hefner, William Keith; Matisak, Brian P.; McElyea, Mark; Kunz, Jennifer; Weber, Philip; Cummings, Nicholas; Parsons, Jeremy

2014-01-01

The National Aeronautics and Space Administration's (NASA) Space Launch System (SLS) Program, managed at the Marshall Space Flight Center (MSFC), is working with the Ground Systems Development and Operations (GSDO) Program, based at the Kennedy Space Center (KSC), to deliver a new safe, affordable, and sustainable capability for human and scientific exploration beyond Earth's orbit (BEO). Larger than the Saturn V Moon rocket, SLS will provide 10 percent more thrust at liftoff in its initial 70 metric ton (t) configuration and 20 percent more in its evolved 130-t configuration. The primary mission of the SLS rocket will be to launch astronauts to deep space destinations in the Orion Multi- Purpose Crew Vehicle (MPCV), also in development and managed by the Johnson Space Center. Several high-priority science missions also may benefit from the increased payload volume and reduced trip times offered by this powerful, versatile rocket. Reducing the lifecycle costs for NASA's space transportation flagship will maximize the exploration and scientific discovery returned from the taxpayer's investment. To that end, decisions made during development of SLS and associated systems will impact the nation's space exploration capabilities for decades. This paper will provide an update to the operations strategy presented at SpaceOps 2012. It will focus on: 1) Preparations to streamline the processing flow and infrastructure needed to produce and launch the world's largest rocket (i.e., through incorporation and modification of proven, heritage systems into the vehicle and ground systems); 2) Implementation of a lean approach to reach-back support of hardware manufacturing, green-run testing, and launch site processing and activities; and 3) Partnering between the vehicle design and operations communities on state-of-the-art predictive operations analysis techniques. An example of innovation is testing the integrated vehicle at the processing facility in parallel, rather than

COSMOS Launch Services

Science.gov (United States)

Kalnins, Indulis

2002-01-01

COSMOS-3M is a two stage launcher with liquid propellant rocket engines. Since 1960's COSMOS has launched satellites of up to 1.500kg in both circular low Earth and elliptical orbits with high inclination. The direct SSO ascent is available from Plesetsk launch site. The very high number of 759 launches and the achieved success rate of 97,4% makes this space transportation system one of the most reliable and successful launchers in the world. The German small satellite company OHB System co-operates since 1994 with the COSMOS manufacturer POLYOT, Omsk, in Russia. They have created the joint venture COSMOS International and successfully launched five German and Italian satellites in 1999 and 2000. The next commercial launches are contracted for 2002 and 2003. In 2005 -2007 COSMOS will be also used for the new German reconnaissance satellite launches. This paper provides an overview of COSMOS-3M launcher: its heritage and performance, examples of scientific and commercial primary and piggyback payload launches, the launch service organization and international cooperation. The COSMOS launch service business strategy main points are depicted. The current and future position of COSMOS in the worldwide market of launch services is outlined.

NASA Sounding Rocket Program Educational Outreach

Science.gov (United States)

Rosanova, G.

2013-01-01

Educational and public outreach is a major focus area for the National Aeronautics and Space Administration (NASA). The NASA Sounding Rocket Program (NSRP) shares in the belief that NASA plays a unique and vital role in inspiring future generations to pursue careers in science, mathematics, and technology. To fulfill this vision, the NSRP engages in a variety of educator training workshops and student flight projects that provide unique and exciting hands-on rocketry and space flight experiences. Specifically, the Wallops Rocket Academy for Teachers and Students (WRATS) is a one-week tutorial laboratory experience for high school teachers to learn the basics of rocketry, as well as build an instrumented model rocket for launch and data processing. The teachers are thus armed with the knowledge and experience to subsequently inspire the students at their home institution. Additionally, the NSRP has partnered with the Colorado Space Grant Consortium (COSGC) to provide a "pipeline" of space flight opportunities to university students and professors. Participants begin by enrolling in the RockOn! Workshop, which guides fledgling rocketeers through the construction and functional testing of an instrumentation kit. This is then integrated into a sealed canister and flown on a sounding rocket payload, which is recovered for the students to retrieve and process their data post flight. The next step in the "pipeline" involves unique, user-defined RockSat-C experiments in a sealed canister that allow participants more independence in developing, constructing, and testing spaceflight hardware. These experiments are flown and recovered on the same payload as the RockOn! Workshop kits. Ultimately, the "pipeline" culminates in the development of an advanced, user-defined RockSat-X experiment that is flown on a payload which provides full exposure to the space environment (not in a sealed canister), and includes telemetry and attitude control capability. The RockOn! and Rock

Sounding rocket study of auroral electron precipitation

International Nuclear Information System (INIS)

McFadden, J.P.

1985-01-01

Measurement of energetic electrons in the auroral zone have proved to be one of the most useful tools in investigating the phenomena of auroral arc formation. This dissertation presents a detailed analysis of the electron data from two sounding rocket campaigns and interprets the measurements in terms of existing auroral models. The Polar Cusp campaign consisted of a single rocket launched from Cape Parry, Canada into the afternoon auroral zone at 1:31:13 UT on January 21, 1982. The results include the measurement of a narrow, magnetic field aligned electron flux at the edge of an arc. This electron precipitation was found to have a remarkably constant 1.2 eV temperature perpendicular to the magnetic field over a 200 to 900 eV energy range. The payload also made simultaneous measurements of both energetic electrons and 3-MHz plasma waves in an auroral arc. Analysis has shown that the waves are propagating in the upper hybrid band and should be generated by a positive slope in the parallel electron distribution. A correlation was found between the 3-MHz waves and small positive slopes in the parallel electron distribution but experimental uncertainties in the electron measurement were large enough to influence the analysis. The BIDARCA campaign consisted of two sounding rockets launched from Poker Flat and Fort Yukon, Alaska at 9:09:00 UT and 9:10:40 UT on February 7, 1984

Game Changing: NASA's Space Launch System and Science Mission Design

Science.gov (United States)

Creech, Stephen D.

2013-01-01

NASA s Marshall Space Flight Center (MSFC) is directing efforts to build the Space Launch System (SLS), a heavy-lift rocket that will carry the Orion Multi-Purpose Crew Vehicle (MPCV) and other important payloads far beyond Earth orbit (BEO). Its evolvable architecture will allow NASA to begin with Moon fly-bys and then go on to transport humans or robots to distant places such as asteroids and Mars. Designed to simplify spacecraft complexity, the SLS rocket will provide improved mass margins and radiation mitigation, and reduced mission durations. These capabilities offer attractive advantages for ambitious missions such as a Mars sample return, by reducing infrastructure requirements, cost, and schedule. For example, if an evolved expendable launch vehicle (EELV) were used for a proposed mission to investigate the Saturn system, a complicated trajectory would be required - with several gravity-assist planetary fly-bys - to achieve the necessary outbound velocity. The SLS rocket, using significantly higher C3 energies, can more quickly and effectively take the mission directly to its destination, reducing trip time and cost. As this paper will report, the SLS rocket will launch payloads of unprecedented mass and volume, such as "monolithic" telescopes and in-space infrastructure. Thanks to its ability to co-manifest large payloads, it also can accomplish complex missions in fewer launches. Future analyses will include reviews of alternate mission concepts and detailed evaluations of SLS figures of merit, helping the new rocket revolutionize science mission planning and design for years to come.

National Security Space Launch Report

Science.gov (United States)

2006-01-01

Company Clayton Mowry, President, Arianespace Inc., North American—“Launch Solutions” Elon Musk , CEO and CTO, Space Exploration Technologies (SpaceX...technologies to the NASA Exploration Initiative (“…Moon, Mars and Beyond.”).1 EELV Technology Needs The Atlas V and Delta IV vehicles incorporate current... Mars and other destinations.” 46 National Security Space Launch Report Figure 6.1 U.S. Government Liquid Propulsion Rocket Investment, 1991–2005

Unsteady Aerodynamic Investigation of the Propeller-Wing Interaction for a Rocket Launched Unmanned Air Vehicle

Directory of Open Access Journals (Sweden)

G. Q. Zhang

2013-01-01

Full Text Available The aerodynamic characteristics of propeller-wing interaction for the rocket launched UAV have been investigated numerically by means of sliding mesh technology. The corresponding forces and moments have been collected for axial wing placements ranging from 0.056 to 0.5D and varied rotating speeds. The slipstream generated by the rotating propeller has little effects on the lift characteristics of the whole UAV. The drag can be seen to remain unchanged as the wing's location moves progressively closer to the propeller until 0.056D away from the propeller, where a nearly 20% increase occurred sharply. The propeller position has a negligible effect on the overall thrust and torque of the propeller. The efficiency affected by the installation angle of the propeller blade has also been analyzed. Based on the pressure cloud and streamlines, the vortices generated by propeller, propeller-wing interaction, and wing tip have also been captured and analyzed.

Launch Pad in a Box

Science.gov (United States)

Mantovani, James; Tamasy, Gabor; Mueller, Rob; Townsend, Van; Sampson, Jeff; Lane, Mike

2016-01-01

NASA Kennedy Space Center (KSC) is developing a new deployable launch system capability to support a small class of launch vehicles for NASA and commercial space companies to test and launch their vehicles. The deployable launch pad concept was first demonstrated on a smaller scale at KSC in 2012 in support of NASA Johnson Space Center's Morpheus Lander Project. The main objective of the Morpheus Project was to test a prototype planetary lander as a vertical takeoff and landing test-bed for advanced spacecraft technologies using a hazard field that KSC had constructed at the Shuttle Landing Facility (SLF). A steel pad for launch or landing was constructed using a modular design that allowed it to be reconfigurable and expandable. A steel flame trench was designed as an optional module that could be easily inserted in place of any modular steel plate component. The concept of a transportable modular launch and landing pad may also be applicable to planetary surfaces where the effects of rocket exhaust plume on surface regolith is problematic for hardware on the surface that may either be damaged by direct impact of high speed dust particles, or impaired by the accumulation of dust (e.g., solar array panels and thermal radiators). During the Morpheus free flight campaign in 2013-14, KSC performed two studies related to rocket plume effects. One study compared four different thermal ablatives that were applied to the interior of a steel flame trench that KSC had designed and built. The second study monitored the erosion of a concrete landing pad following each landing of the Morpheus vehicle on the same pad located in the hazard field. All surfaces of a portable flame trench that could be directly exposed to hot gas during launch of the Morpheus vehicle were coated with four types of ablatives. All ablative products had been tested by NASA KSC and/or the manufacturer. The ablative thicknesses were measured periodically following the twelve Morpheus free flight tests

Kinetic modeling of auroral ion outflows observed by the VISIONS sounding rocket

Science.gov (United States)

Albarran, R. M.; Zettergren, M. D.

2017-12-01

The VISIONS (VISualizing Ion Outflow via Neutral atom imaging during a Substorm) sounding rocket was launched on Feb. 7, 2013 at 8:21 UTC from Poker Flat, Alaska, into an auroral substorm with the objective of identifying the drivers and dynamics of the ion outflow below 1000km. Energetic ion data from the VISIONS polar cap boundary crossing show evidence of an ion "pressure cooker" effect whereby ions energized via transverse heating in the topside ionosphere travel upward and are impeded by a parallel potential structure at higher altitudes. VISIONS was also instrumented with an energetic neutral atom (ENA) detector which measured neutral particles ( 50-100 eV energy) presumably produced by charge-exchange with the energized outflowing ions. Hence, inferences about ion outflow may be made via remotely-sensing measurements of ENAs. This investigation focuses on modeling energetic outflowing ion distributions observed by VISIONS using a kinetic model. This kinetic model traces large numbers of individual particles, using a guiding-center approximation, in order to allow calculation of ion distribution functions and moments. For the present study we include mirror and parallel electric field forces, and a source of ion cyclotron resonance (ICR) wave heating, thought to be central to the transverse energization of ions. The model is initiated with a steady-state ion density altitude profile and Maxwellian velocity distribution characterizing the initial phase-space conditions for multiple particle trajectories. This project serves to advance our understanding of the drivers and particle dynamics in the auroral ionosphere and to improve data analysis methods for future sounding rocket and satellite missions.

Relationship of Worldwide Rocket Launch Crashes with Geophysical Parameters

Directory of Open Access Journals (Sweden)

N. Romanova

2013-01-01

Full Text Available A statistical comparison of launch crashes at different worldwide space ports with geophysical factors has been performed. A comprehensive database has been compiled, which includes 50 years of information from the beginning of the space age in 1957 about launch crashes occurring world-wide. Special attention has been paid to statistics concerning launches at the largest space ports: Plesetsk, Baikonur, Cape Canaveral, and Vandenberg. In search of a possible influence of geophysical factors on launch failures, such parameters as the vehicle type, local time, season, sunspot number, high-energy electron fluxes, and solar proton events have been examined. Also, we have analyzed correlations with the geomagnetic indices as indirect indicators of the space weather condition. Regularities found in this study suggest that further detailed studies of space weather effects on launcher systems, especially in the high-latitude regions, should be performed.

The reusable launch vehicle technology program

Science.gov (United States)

Cook, S.

1995-01-01

Today's launch systems have major shortcomings that will increase in significance in the future, and thus are principal drivers for seeking major improvements in space transportation. They are too costly; insufficiently reliable, safe, and operable; and increasingly losing market share to international competition. For the United States to continue its leadership in the human exploration and wide ranging utilization of space, the first order of business must be to achieve low cost, reliable transportatin to Earth orbit. NASA's Access to Space Study, in 1993, recommended the development of a fully reusable single-stage-to-orbit (SSTO) rocket vehicle as an Agency goal. The goal of the Reusable Launch Vehicle (RLV) technology program is to mature the technologies essential for a next-generation reusable launch system capable of reliably serving National space transportation needs at substantially reduced costs. The primary objectives of the RLV technology program are to (1) mature the technologies required for the next-generation system, (2) demonstrate the capability to achieve low development and operational cost, and rapid launch turnaround times and (3) reduce business and technical risks to encourage significant private investment in the commercial development and operation of the next-generation system. Developing and demonstrating the technologies required for a Single Stage to Orbit (SSTO) rocket is a focus of the program becuase past studies indicate that it has the best potential for achieving the lowest space access cost while acting as an RLV technology driver (since it also encompasses the technology requirements of reusable rocket vehicles in general).

The reusable launch vehicle technology program

Science.gov (United States)

Cook, S.

Today's launch systems have major shortcomings that will increase in significance in the future, and thus are principal drivers for seeking major improvements in space transportation. They are too costly; insufficiently reliable, safe, and operable; and increasingly losing market share to international competition. For the United States to continue its leadership in the human exploration and wide ranging utilization of space, the first order of business must be to achieve low cost, reliable transportatin to Earth orbit. NASA's Access to Space Study, in 1993, recommended the development of a fully reusable single-stage-to-orbit (SSTO) rocket vehicle as an Agency goal. The goal of the Reusable Launch Vehicle (RLV) technology program is to mature the technologies essential for a next-generation reusable launch system capable of reliably serving National space transportation needs at substantially reduced costs. The primary objectives of the RLV technology program are to (1) mature the technologies required for the next-generation system, (2) demonstrate the capability to achieve low development and operational cost, and rapid launch turnaround times and (3) reduce business and technical risks to encourage significant private investment in the commercial development and operation of the next-generation system. Developing and demonstrating the technologies required for a Single Stage to Orbit (SSTO) rocket is a focus of the program becuase past studies indicate that it has the best potential for achieving the lowest space access cost while acting as an RLV technology driver (since it also encompasses the technology requirements of reusable rocket vehicles in general).

First results of the Auroral Turbulance II rocket experiment

DEFF Research Database (Denmark)

Danielides, M.A.; Ranta, A.; Ivchenco, N.

1999-01-01

The Auroral Turbulance II sounding rocket was launched on February 11, 1997 into moderately active nightside aurora from the Poker Flat Research Range, Alaska, US. The experiment consisted of three independent, completely instrumented payloads launched by a single vehicle. The aim of the experiment...

78 FR 40196 - National Environmental Policy Act; Sounding Rockets Program; Poker Flat Research Range

Science.gov (United States)

2013-07-03

...; Sounding Rockets Program; Poker Flat Research Range AGENCY: National Aeronautics and Space Administration... Sounding Rockets Program (SRP) at Poker Flat Research Range (PFRR), Alaska. SUMMARY: Pursuant to the... government agencies, and educational institutions have conducted suborbital rocket launches from the PFRR...

Lymphocytes on sounding rocket flights.

Science.gov (United States)

Cogoli-Greuter, M; Pippia, P; Sciola, L; Cogoli, A

1994-05-01

Cell-cell interactions and the formation of cell aggregates are important events in the mitogen-induced lymphocyte activation. The fact that the formation of cell aggregates is only slightly reduced in microgravity suggests that cells are moving and interacting also in space, but direct evidence was still lacking. Here we report on two experiments carried out on a flight of the sounding rocket MAXUS 1B, launched in November 1992 from the base of Esrange in Sweden. The rocket reached the altitude of 716 km and provided 12.5 min of microgravity conditions.

Aerodynamic Problems of Launch Vehicles

Directory of Open Access Journals (Sweden)

Kyong Chol Chou

1984-09-01

Full Text Available The airflow along the surface of a launch vehicle together with vase flow of clustered nozzles cause problems which may affect the stability or efficiency of the entire vehicle. The problem may occur when the vehicle is on the launching pad or even during flight. As for such problems, local steady-state loads, overall steady-state loads, buffet, ground wind loads, base heating and rocket-nozzle hinge moments are examined here specifically.

Wave-particle interaction phenomena observed by antarctic rockets

International Nuclear Information System (INIS)

Kimura, I.; Hirasawa, T.

1979-01-01

Rocket measurements of wave and particles activities made at Syowa Station in Antarctica during IMS period are reviewed. Nine rockets were used for such observations, out of which 6 rockets were launched in the auroral sky. In the VLF frequency range, 0 - 10 KHz, wideband spectra of wave electric and magnetic fields, Poynting flux and the direction of propagation vector were measured for chorus, ELF and VLF hiss, and for electrostatic noises. In the MF and HF range, the dynamic frequency spectra of 0.1 - 10 MHz were measured. The relationship of these wave phenomena with energetic particle activities measured by the same rockets are discussed. (author)

Introduction to the Special Issue on Sounding Rockets and Instrumentation

OpenAIRE

Christe, Steven; Zeiger, Ben; Pfaff, Rob; Garcia, Michael

2016-01-01

Rocket technology, originally developed for military applications, has provided a low-cost observing platform to carry critical and rapid-response scientific investigations for over 70 years. Even with the development of launch vehicles that could put satellites into orbit, high altitude sounding rockets have remained relevant. In addition to science observations, sounding rockets provide a unique technology test platform and a valuable training ground for scientists and engineers. Most impor...

Project Stratos; reaching space with a student-built rocket

NARCIS (Netherlands)

Haneveer, M.

2013-01-01

In the spring of 2009 a team of 15 TU Delft students travelled to Kiruna, Sweden with only one goal: to launch the rocket Stratos I they had been working on for 2 years to an altitude of over 12km, thereby claiming the European Amateur Rocket Altitude record. These students were part of Delft

The Isinglass Auroral Sounding Rocket Campaign: data synthesis incorporating remote sensing, in situ observations, and modelling

Science.gov (United States)

Lynch, K. A.; Clayton, R.; Roberts, T. M.; Hampton, D. L.; Conde, M.; Zettergren, M. D.; Burleigh, M.; Samara, M.; Michell, R.; Grubbs, G. A., II; Lessard, M.; Hysell, D. L.; Varney, R. H.; Reimer, A.

2017-12-01

The NASA auroral sounding rocket mission Isinglass was launched from Poker Flat Alaska in winter 2017. This mission consists of two separate multi-payload sounding rockets, over an array of groundbased observations, including radars and filtered cameras. The science goal is to collect two case studies, in two different auroral events, of the gradient scale sizes of auroral disturbances in the ionosphere. Data from the in situ payloads and the groundbased observations will be synthesized and fed into an ionospheric model, and the results will be studied to learn about which scale sizes of ionospheric structuring have significance for magnetosphere-ionosphere auroral coupling. The in situ instrumentation includes thermal ion sensors (at 5 points on the second flight), thermal electron sensors (at 2 points), DC magnetic fields (2 point), DC electric fields (one point, plus the 4 low-resource thermal ion RPA observations of drift on the second flight), and an auroral precipitation sensor (one point). The groundbased array includes filtered auroral imagers, the PFISR and SuperDarn radars, a coherent scatter radar, and a Fabry-Perot interferometer array. The ionospheric model to be used is a 3d electrostatic model including the effects of ionospheric chemistry. One observational and modelling goal for the mission is to move both observations and models of auroral arc systems into the third (along-arc) dimension. Modern assimilative tools combined with multipoint but low-resource observations allow a new view of the auroral ionosphere, that should allow us to learn more about the auroral zone as a coupled system. Conjugate case studies such as the Isinglass rocket flights allow for a test of the models' intepretation by comparing to in situ data. We aim to develop and improve ionospheric models to the point where they can be used to interpret remote sensing data with confidence without the checkpoint of in situ comparison.

Air Launch from a Towed Glider

Data.gov (United States)

National Aeronautics and Space Administration — This research effort is exploring the concept of launching a rocket from a glider that is towed by an aircraft. The idea is to build a relatively inexpensive...

An X-ray Experiment with Two-Stage Korean Sounding Rocket

Directory of Open Access Journals (Sweden)

Uk-Won Nam

1998-12-01

Full Text Available The test result of the X-ray observation system is presented which have been developed at Korea Astronomy Observatory for 3 years (1995-1997. The instrument, which is composed of detector and signal processing parts, is designed for the future observations of compact X-ray sources. The performance of the instrument was tested by mounting on the two-stage Korean Sounding Rocket, which was launched from Taean rocket flight center on June 11 at 10:00 KST 1998. Telemetry data were received from individual parts of the instrument for 32 and 55.7 sec, respectively, since the launch of the rocket. In this paper, the result of the data analysis based on the telemetry data and discussion about the performance of the instrument is reported.

Vertical Wind Tunnel for Prediction of Rocket Flight Dynamics

Directory of Open Access Journals (Sweden)

Hoani Bryson

2016-03-01

Full Text Available A customized vertical wind tunnel has been built by the University of Canterbury Rocketry group (UC Rocketry. This wind tunnel has been critical for the success of UC Rocketry as it allows the optimization of avionics and control systems before flight. This paper outlines the construction of the wind tunnel and includes an analysis of flow quality including swirl. A minimal modelling methodology for roll dynamics is developed that can extrapolate wind tunnel behavior at low wind speeds to much higher velocities encountered during flight. The models were shown to capture the roll flight dynamics in two rocket launches with mean roll angle errors varying from 0.26° to 1.5° across the flight data. The identified model parameters showed consistent and predictable variations over both wind tunnel tests and flight, including canard–fin interaction behavior. These results demonstrate that the vertical wind tunnel is an important tool for the modelling and control of sounding rockets.

Design options for advanced manned launch systems

Science.gov (United States)

Freeman, Delma C.; Talay, Theodore A.; Stanley, Douglas O.; Lepsch, Roger A.; Wilhite, Alan W.

1995-03-01

Various concepts for advanced manned launch systems are examined for delivery missions to space station and polar orbit. Included are single-and two-stage winged systems with rocket and/or air-breathing propulsion systems. For near-term technologies, two-stage reusable rocket systems are favored over single-stage rocket or two-stage air-breathing/rocket systems. Advanced technologies enable viable single-stage-to-orbit (SSTO) concepts. Although two-stage rocket systems continue to be lighter in dry weight than SSTO vehicles, advantages in simpler operations may make SSTO vehicles more cost-effective over the life cycle. Generally, rocket systems maintain a dry-weight advantage over air-breathing systems at the advanced technology levels, but to a lesser degree than when near-term technologies are used. More detailed understanding of vehicle systems and associated ground and flight operations requirements and procedures is essential in determining quantitative discrimination between these latter concepts.

Structural strengthening of rocket nozzle extension by means of laser metal deposition

Science.gov (United States)

Honoré, M.; Brox, L.; Hallberg, M.

2012-03-01

Commercial space operations strive to maximize the payload per launch in order to minimize the costs of each kg launched into orbit; this yields demand for ever larger launchers with larger, more powerful rocket engines. Volvo Aero Corporation in collaboration with Snecma and Astrium has designed and tested a new, upgraded Nozzle extension for the Vulcain 2 engine configuration, denoted Vulcain 2+ NE Demonstrator The manufacturing process for the welding of the sandwich wall and the stiffening structure is developed in close cooperation with FORCE Technology. The upgrade is intended to be available for future development programs for the European Space Agency's (ESA) highly successful commercial launch vehicle, the ARIANE 5. The Vulcain 2+ Nozzle Extension Demonstrator [1] features a novel, thin-sheet laser-welded configuration, with laser metal deposition built-up 3D-features for the mounting of stiffening structure, flanges and for structural strengthening, in order to cope with the extreme load- and thermal conditions, to which the rocket nozzle extension is exposed during launch of the 750 ton ARIANE 5 launcher. Several millimeters of material thickness has been deposited by laser metal deposition without disturbing the intricate flow geometry of the nozzle cooling channels. The laser metal deposition process has been applied on a full-scale rocket nozzle demonstrator, and in excess of 15 kilometers of filler wire has been successfully applied to the rocket nozzle. The laser metal deposition has proven successful in two full-throttle, full-scale tests, firing the rocket engine and nozzle in the ESA test facility P5 by DLR in Lampoldshausen, Germany.

X-ray scanning of overhead aurorae from rockets

International Nuclear Information System (INIS)

Barcus, J.R.; Goldberg, R.A.

1981-01-01

Two Nike Tomahawk rocket payloads were launched into energetic auroral events to investigate their structure and effects on the atmosphere. The instrument complement included X-ray scintillation detectors with energy discrimination in four ranges to measure the deposition of bremsstrahlung produced X-rays within the stratosphere and mesosphere. For this purpose, each instrument was designed for wide angle viewing; however, properties of the rocket motion have permitted coarse observation of distinct spatial X-ray structure. The detectors were mounted at a 45 0 angle with respect to the payload axis to permit scanning of the upper hemisphere, with rocket spin rates near 5 c/s during the upleg portion of each flight. Here, atmospheric shielding reduced energetic particle contamination effects to insignificant values below 65 to 75 km. Iterative computer techniques were used to reconstruct X-ray source maps at 100 km, taking atmospheric absorption effects into account. Payload 18.178 was launched on 21 September (0302 LMT) into an aurora observed to have two distinct azimuthal regions of optical brightness. Payload 18.179 (23 September, 0147 LMT) was launched into an aurora of more diffuse character. The presence of a two component spectrum is indicated for each event with the hard component originating in the more diffuse, optically faint regions. (author)

Water Rockets. Get Funny With Newton's Laws

Directory of Open Access Journals (Sweden)

Manuel Roca Vicent

2017-01-01

Full Text Available The study of the movement of the rocket has been used for decades to encourage students in the study of physics. This system has an undeniable interest to introduce concepts such as properties of gases, laws of Newton,  exchange  between  different  types  of  energy  and  its  conservation  or fluid  mechanics.  Our  works has  been  to  build  and  launch  these  rockets  in  different  educational  levels  and  in  each  of  these  ones  have introduced  the  part  of  Physics  more  suited  to  the  knowledge  of  our  students.  The  aim  of  the  learning experience  is  to  launch  the  rocket  as  far  as  possible  and  learn  to  predict  the  travelled  distance,  using Newton's  laws  and fluid  mechanics.  After  experimentation  we  demonstrated  to  be  able  to  control  the parameters that improve the performance of our rocket, such as the  fill factor, the volume and mass of the empty  bottle,  liquid  density,  launch  angle,  pressure  prior  air  release.  In addition, it is a fun experience can be attached to all levels of education in primary and high school.

The 4D-var Estimation of North Korean Rocket Exhaust Emissions Into the Ionosphere

Science.gov (United States)

Ssessanga, Nicholas; Kim, Yong Ha; Choi, Byungyu; Chung, Jong-Kyun

2018-03-01

We have developed a four-dimensional variation data assimilation technique (4D-var) and utilized it to reconstruct three-dimensional images of the ionospheric hole created during Kwangmyongsong-4 rocket launch. Kwangmyongsong-4 was launched southward from North Korea Sohae space center (124.7°E, 39.6°N) at 00:30 UT on 7 February 2016. The data assimilated were Global Positioning System total electron content from the South Korean Global Positioning System-receiver network. Due to lack of publicized information about Kwangmyongsong-4, the rocket was assumed to inherit its technology from previous launches (Taepodong-2). The created ionospheric hole was assumed to be made by neutral molecules, water (H2O) and hydrogen (H2), deposited in exhaust plumes. The dispersion model was developed based on advection and diffusion equation, and a simple asymmetric diffusion model assumed. From the analysis, using the adjoint technique, we estimated an ionospheric hole with the largest depletion existing around 6-7 min after launch and gradually recovering within 30 min. These results are in agreement with temporal total electron content analyses of the same event from previous studies. Furthermore, Kwangmyongsong-4 second stage exhaust emissions were estimated as 1.9 × 1026 s-1 of which 40% was H2 and the rest H2O.

The Chameleon Solid Rocket Propulsion Model

International Nuclear Information System (INIS)

Robertson, Glen A.

2010-01-01

The Khoury and Weltman (2004a and 2004b) Chameleon Model presents an addition to the gravitation force and was shown by the author (Robertson, 2009a and 2009b) to present a new means by which one can view other forces in the Universe. The Chameleon Model is basically a density-dependent model and while the idea is not new, this model is novel in that densities in the Universe to include the vacuum of space are viewed as scalar fields. Such an analogy gives the Chameleon scalar field, dark energy/dark matter like characteristics; fitting well within cosmological expansion theories. In respect to this forum, in this paper, it is shown how the Chameleon Model can be used to derive the thrust of a solid rocket motor. This presents a first step toward the development of new propulsion models using density variations verse mass ejection as the mechanism for thrust. Further, through the Chameleon Model connection, these new propulsion models can be tied to dark energy/dark matter toward new space propulsion systems utilizing the vacuum scalar field in a way understandable by engineers, the key toward the development of such systems. This paper provides corrections to the Chameleon rocket model in Robertson (2009b).

Rocket measurements of positive ions during polar mesosphere winter echo conditions

Directory of Open Access Journals (Sweden)

A. Brattli

2006-01-01

Full Text Available On 18 January 2005, two small, instrumented rockets were launched from Andøya Rocket Range (69.3° N, 16° E during conditions with Polar Mesosphere Winter Echoes (PMWE. Each of the rockets was equipped with a Positive Ion Probe (PIP and a Faraday rotation/differential absorption experiment, and was launched as part of a salvo of meteorological rockets measuring temperature and wind using falling spheres and chaff. Layers of PMWE were detected between 55 and 77 km by the 53.5 MHz ALWIN radar. The rockets were launched during a solar proton event, and measured extremely high ion densities, of order 1010 m−3, in the region where PMWE were observed. The density measurements were analyzed with the wavelet transform technique. At large length scales, ~103 m, the power spectral density can be fitted with a k−3 wave number dependence, consistent with saturated gravity waves. Outside the PMWE layers the k−3 spectrum extends down to approximately 102 m where the fluctuations are quickly damped and disappear into the instrumental noise. Inside the PMWE layers the spectrum at smaller length scales is well fitted with a k−5/3 dependence over two decades of scales. The PMWE are therefore clearly indicative of turbulence, and the data are consistent with the turbulent dissipation of breaking gravity waves. We estimate a lower limit for the turbulent energy dissipation rate of about 10−2 W/kg in the upper (72 km layer.

Using NASA's Space Launch System to Enable Game Changing Science Mission Designs

Science.gov (United States)

Creech, Stephen D.

2013-01-01

NASA's Marshall Space Flight Center is directing efforts to build the Space Launch System (SLS), a heavy-lift rocket that will help restore U.S. leadership in space by carrying the Orion Multi-Purpose Crew Vehicle and other important payloads far beyond Earth orbit. Its evolvable architecture will allow NASA to begin with Moon fly-bys and then go on to transport humans or robots to distant places such as asteroids, Mars, and the outer solar system. Designed to simplify spacecraft complexity, the SLS rocket will provide improved mass margins and radiation mitigation, and reduced mission durations. These capabilities offer attractive advantages for ambitious missions such as a Mars sample return, by reducing infrastructure requirements, cost, and schedule. For example, if an evolved expendable launch vehicle (EELV) were used for a proposed mission to investigate the Saturn system, a complicated trajectory would be required with several gravity-assist planetary fly-bys to achieve the necessary outbound velocity. The SLS rocket, using significantly higher C3 energies, can more quickly and effectively take the mission directly to its destination, reducing trip times and cost. As this paper will report, the SLS rocket will launch payloads of unprecedented mass and volume, such as monolithic telescopes and in-space infrastructure. Thanks to its ability to co-manifest large payloads, it also can accomplish complex missions in fewer launches. Future analyses will include reviews of alternate mission concepts and detailed evaluations of SLS figures of merit, helping the new rocket revolutionize science mission planning and design for years to come.

Design optimization of space launch vehicles using a genetic algorithm

Science.gov (United States)

Bayley, Douglas James

The United States Air Force (USAF) continues to have a need for assured access to space. In addition to flexible and responsive spacelift, a reduction in the cost per launch of space launch vehicles is also desirable. For this purpose, an investigation of the design optimization of space launch vehicles has been conducted. Using a suite of custom codes, the performance aspects of an entire space launch vehicle were analyzed. A genetic algorithm (GA) was employed to optimize the design of the space launch vehicle. A cost model was incorporated into the optimization process with the goal of minimizing the overall vehicle cost. The other goals of the design optimization included obtaining the proper altitude and velocity to achieve a low-Earth orbit. Specific mission parameters that are particular to USAF space endeavors were specified at the start of the design optimization process. Solid propellant motors, liquid fueled rockets, and air-launched systems in various configurations provided the propulsion systems for two, three and four-stage launch vehicles. Mass properties models, an aerodynamics model, and a six-degree-of-freedom (6DOF) flight dynamics simulator were all used to model the system. The results show the feasibility of this method in designing launch vehicles that meet mission requirements. Comparisons to existing real world systems provide the validation for the physical system models. However, the ability to obtain a truly minimized cost was elusive. The cost model uses an industry standard approach, however, validation of this portion of the model was challenging due to the proprietary nature of cost figures and due to the dependence of many existing systems on surplus hardware.

The UK sounding rocket and balloon programme

International Nuclear Information System (INIS)

Delury, J.T.

1980-01-01

The UK civil science balloon and rocket programmes for 1979/80/81 are summarised and the areas of scientific interest for the period 1981/85 mentioned. In the main the facilities available are 10 in number balloons up to 40 m cu ft launched from USA or Australia and up to 10 in number 7 1/2'' diameter Petrel rockets. This paper outlines the 1979 and 1980 programmes and explains the longer term plans covering the next 5 years. (Auth.)

Coupled Fluid-Structure Interaction Analysis of Solid Rocket Motor with Flexible Inhibitors

Science.gov (United States)

Yang, H. Q.; West, Jeff; Harris, Robert E.

2014-01-01

Flexible inhibitors are generally used in solid rocket motors (SRMs) as a means to control the burning of propellant. Vortices generated by the flow of propellant around the flexible inhibitors have been identified as a driving source of instabilities that can lead to thrust oscillations in launch vehicles. Potential coupling between the SRM thrust oscillations and structural vibration modes is an important risk factor in launch vehicle design. As a means to predict and better understand these phenomena, a multidisciplinary simulation capability that couples the NASA production CFD code, Loci/CHEM, with CFDRC's structural finite element code, CoBi, has been developed. This capability is crucial to the development of NASA's new space launch system (SLS). This paper summarizes the efforts in applying the coupled software to demonstrate and investigate fluid-structure interaction (FSI) phenomena between pressure waves and flexible inhibitors inside reusable solid rocket motors (RSRMs). The features of the fluid and structural solvers are described in detail, and the coupling methodology and interfacial continuity requirements are then presented in a general Eulerian-Lagrangian framework. The simulations presented herein utilize production level CFD with hybrid RANS/LES turbulence modeling and grid resolution in excess of 80 million cells. The fluid domain in the SRM is discretized using a general mixed polyhedral unstructured mesh, while full 3D shell elements are utilized in the structural domain for the flexible inhibitors. Verifications against analytical solutions for a structural model under a steady uniform pressure condition and under dynamic modal analysis show excellent agreement in terms of displacement distribution and eigenmode frequencies. The preliminary coupled results indicate that due to acoustic coupling, the dynamics of one of the more flexible inhibitors shift from its first modal frequency to the first acoustic frequency of the solid rocket motor

Russian Meteorological and Geophysical Rockets of New Generation

Science.gov (United States)

Yushkov, V.; Gvozdev, Yu.; Lykov, A.; Shershakov, V.; Ivanov, V.; Pozin, A.; Afanasenkov, A.; Savenkov, Yu.; Kuznetsov, V.

2015-09-01

To study the process in the middle and upper atmosphere, ionosphere and near-Earth space, as well as to monitor the geophysical environment in Russian Federal Service for Hydrology and Environmental Monitoring (ROSHYDROMET) the development of new generation of meteorological and geophysical rockets has been completed. The modern geophysical research rocket system MR-30 was created in Research and Production Association RPA "Typhoon". The basis of the complex MR-30 is a new geophysical sounding rocket MN-300 with solid propellant, Rocket launch takes place at an angle of 70º to 90º from the launcher, which is a farm with a guide rail type required for imparting initial rotation rocket. The Rocket is spin stabilized with a spin rate between 5 and 7 Hz. Launch weight is 1564 kg, and the mass of the payload of 50 to 150 kg. MR-300 is capable of lifting up to 300 km, while the area of dispersion points for booster falling is an ellipse with parameters 37x 60 km. The payload of the rocket MN-300 consists of two sections: a sealed, located below the instrument compartment, and not sealed, under the fairing. Block of scientific equipment is formed on the platform in a modular layout. This makes it possible to solve a wide range of tasks and conduct research and testing technologies using a unique environment of space, as well as to conduct technological experiments testing and research systems and spacecraft equipment. New Russian rocket system MERA (MEteorological Rocket for Atmospheric Research) belongs to so called "dart" technique that provide lifting of small scientific payload up to altitude 100 km and descending with parachute. It was developed at Central Aerological Observatory jointly with State Unitary Enterprise Instrument Design Bureau. The booster provides a very rapid acceleration to about Mach 5. After the burning phase of the buster the dart is separated and continues ballistic flight for about 2 minutes. The dart carries the instrument payload+ parachute

Space Launch System Development Status

Science.gov (United States)

Lyles, Garry

2014-01-01

Development of NASA's Space Launch System (SLS) heavy lift rocket is shifting from the formulation phase into the implementation phase in 2014, a little more than three years after formal program approval. Current development is focused on delivering a vehicle capable of launching 70 metric tons (t) into low Earth orbit. This "Block 1" configuration will launch the Orion Multi-Purpose Crew Vehicle (MPCV) on its first autonomous flight beyond the Moon and back in December 2017, followed by its first crewed flight in 2021. SLS can evolve to a130-t lift capability and serve as a baseline for numerous robotic and human missions ranging from a Mars sample return to delivering the first astronauts to explore another planet. Benefits associated with its unprecedented mass and volume include reduced trip times and simplified payload design. Every SLS element achieved significant, tangible progress over the past year. Among the Program's many accomplishments are: manufacture of Core Stage test panels; testing of Solid Rocket Booster development hardware including thrust vector controls and avionics; planning for testing the RS-25 Core Stage engine; and more than 4,000 wind tunnel runs to refine vehicle configuration, trajectory, and guidance. The Program shipped its first flight hardware - the Multi-Purpose Crew Vehicle Stage Adapter (MSA) - to the United Launch Alliance for integration with the Delta IV heavy rocket that will launch an Orion test article in 2014 from NASA's Kennedy Space Center. Objectives of this Earth-orbit flight include validating the performance of Orion's heat shield and the MSA design, which will be manufactured again for SLS missions to deep space. The Program successfully completed Preliminary Design Review in 2013 and Key Decision Point C in early 2014. NASA has authorized the Program to move forward to Critical Design Review, scheduled for 2015 and a December 2017 first launch. The Program's success to date is due to prudent use of proven

Microcomputers, Model Rockets, and Race Cars.

Science.gov (United States)

Mirus, Edward A., Jr.

1985-01-01

The industrial education orientation program at Wisconsin School for the Deaf (WSD) presents problem-solving situations to all seventh- and eighth-grade hearing-impaired students. WSD developed user-friendly microcomputer software to guide students individually through complex computations involving model race cars and rockets while freeing…

Ionospheric effects of rocket exhaust products: Skylab and HEAO-C

International Nuclear Information System (INIS)

Zinn, J.; Sutherland, C.D.; Duncan, L.M.; Stone, S.N.

1981-01-01

This paper is about ionospheric F-layer depletions produced by chemical reactions with exhaust gases from large rockets. It describes a 2-dimensional computer model of the ionosphere, and it compares model results with experimental data on the structure and variability of the natural ionosphere, as well as data on ionospheric holes produced by the launches of Skylab (May, 1973) and HEAO-C (September, 1979). It also describes measurements made in conjunction with the HEAO-C launch. The computer model includes an approximate representation of thermospheric tidal winds and E fields in addition to vertical motions associated with diurnal changes in temperature. The computed ionospheric structure is sensitive to all the above. For a small number of cases, results are compared of computations of the normal diurnal variations of ionospheric structure with incoherent scatter and total electron content data. Computations of ionospheric depletions from the Skylab and HEAO-C launches are in satisfactory agreement with the observations. The winds appear to be essential for interpretation of the Skylab results

The Water Recovery X-ray Rocket (WRX-R)

Science.gov (United States)

Miles, Drew

2017-08-01

The Water Recovery X-ray Rocket (WRX-R) is a diffuse soft X-ray spectrometer that will launch on a sounding rocket from the Kwajalein Atoll. WRX-R has a field of view of >10 deg2 and will observe the Vela supernova remnant. A mechanical collimator, state-of-the-art off-plane reflection grating array and hybrid CMOS detector will allow WRX to achieve the most highly-resolved spectrum of the Vela SNR ever recorded. In addition, this payload will fly a hard X-ray telescope that is offset from the soft X-ray spectrometer in order to observe the pulsar at the center of the remnant. We present here an introduction to the instrument, the expected science return, and an update on the state of the payload as we work towards launch.

Hybrids - Best of both worlds. [liquid and solid propellants mated for safe reliable and low cost launch vehicles

Science.gov (United States)

Goldberg, Ben E.; Wiley, Dan R.

1991-01-01

An overview is presented of hybrid rocket propulsion systems whereby combining solids and liquids for launch vehicles could produce a safe, reliable, and low-cost product. The primary subsystems of a hybrid system consist of the oxidizer tank and feed system, an injector system, a solid fuel grain enclosed in a pressure vessel case, a mixing chamber, and a nozzle. The hybrid rocket has an inert grain, which reduces costs of development, transportation, manufacturing, and launch by avoiding many safety measures that must be taken when operating with solids. Other than their use in launch vehicles, hybrids are excellent for simulating the exhaust of solid rocket motors for material development.

Nitrous Oxide/Paraffin Hybrid Rocket Engines

Science.gov (United States)

Zubrin, Robert; Snyder, Gary

2010-01-01

Nitrous oxide/paraffin (N2OP) hybrid rocket engines have been invented as alternatives to other rocket engines especially those that burn granular, rubbery solid fuels consisting largely of hydroxyl- terminated polybutadiene (HTPB). Originally intended for use in launching spacecraft, these engines would also be suitable for terrestrial use in rocket-assisted takeoff of small airplanes. The main novel features of these engines are (1) the use of reinforced paraffin as the fuel and (2) the use of nitrous oxide as the oxidizer. Hybrid (solid-fuel/fluid-oxidizer) rocket engines offer advantages of safety and simplicity over fluid-bipropellant (fluid-fuel/fluid-oxidizer) rocket en - gines, but the thrusts of HTPB-based hybrid rocket engines are limited by the low regression rates of the fuel grains. Paraffin used as a solid fuel has a regression rate about 4 times that of HTPB, but pure paraffin fuel grains soften when heated; hence, paraffin fuel grains can, potentially, slump during firing. In a hybrid engine of the present type, the paraffin is molded into a 3-volume-percent graphite sponge or similar carbon matrix, which supports the paraffin against slumping during firing. In addition, because the carbon matrix material burns along with the paraffin, engine performance is not appreciably degraded by use of the matrix.

Space Launch System Scale Model Acoustic Test Ignition Overpressure Testing

Science.gov (United States)

Nance, Donald; Liever, Peter; Nielsen, Tanner

2015-01-01

The overpressure phenomenon is a transient fluid dynamic event occurring during rocket propulsion system ignition. This phenomenon results from fluid compression of the accelerating plume gas, subsequent rarefaction, and subsequent propagation from the exhaust trench and duct holes. The high-amplitude unsteady fluid-dynamic perturbations can adversely affect the vehicle and surrounding structure. Commonly known as ignition overpressure (IOP), this is an important design-to environment for the Space Launch System (SLS) that NASA is currently developing. Subscale testing is useful in validating and verifying the IOP environment. This was one of the objectives of the Scale Model Acoustic Test, conducted at Marshall Space Flight Center. The test data quantifies the effectiveness of the SLS IOP suppression system and improves the analytical models used to predict the SLS IOP environments. The reduction and analysis of the data gathered during the SMAT IOP test series requires identification and characterization of multiple dynamic events and scaling of the event waveforms to provide the most accurate comparisons to determine the effectiveness of the IOP suppression systems. The identification and characterization of the overpressure events, the waveform scaling, the computation of the IOP suppression system knockdown factors, and preliminary comparisons to the analytical models are discussed.

Space Launch System Scale Model Acoustic Test Ignition Overpressure Testing

Science.gov (United States)

Nance, Donald K.; Liever, Peter A.

2015-01-01

The overpressure phenomenon is a transient fluid dynamic event occurring during rocket propulsion system ignition. This phenomenon results from fluid compression of the accelerating plume gas, subsequent rarefaction, and subsequent propagation from the exhaust trench and duct holes. The high-amplitude unsteady fluid-dynamic perturbations can adversely affect the vehicle and surrounding structure. Commonly known as ignition overpressure (IOP), this is an important design-to environment for the Space Launch System (SLS) that NASA is currently developing. Subscale testing is useful in validating and verifying the IOP environment. This was one of the objectives of the Scale Model Acoustic Test (SMAT), conducted at Marshall Space Flight Center (MSFC). The test data quantifies the effectiveness of the SLS IOP suppression system and improves the analytical models used to predict the SLS IOP environments. The reduction and analysis of the data gathered during the SMAT IOP test series requires identification and characterization of multiple dynamic events and scaling of the event waveforms to provide the most accurate comparisons to determine the effectiveness of the IOP suppression systems. The identification and characterization of the overpressure events, the waveform scaling, the computation of the IOP suppression system knockdown factors, and preliminary comparisons to the analytical models are discussed.

A Reference Model for Virtual Machine Launching Overhead

Energy Technology Data Exchange (ETDEWEB)

Wu, Hao; Ren, Shangping; Garzoglio, Gabriele; Timm, Steven; Bernabeu, Gerard; Chadwick, Keith; Noh, Seo-Young

2016-07-01

Cloud bursting is one of the key research topics in the cloud computing communities. A well designed cloud bursting module enables private clouds to automatically launch virtual machines (VMs) to public clouds when more resources are needed. One of the main challenges in developing a cloud bursting module is to decide when and where to launch a VM so that all resources are most effectively and efficiently utilized and the system performance is optimized. However, based on system operational data obtained from FermiCloud, a private cloud developed by the Fermi National Accelerator Laboratory for scientific workflows, the VM launching overhead is not a constant. It varies with physical resource utilization, such as CPU and I/O device utilizations, at the time when a VM is launched. Hence, to make judicious decisions as to when and where a VM should be launched, a VM launching overhead reference model is needed. In this paper, we first develop a VM launching overhead reference model based on operational data we have obtained on FermiCloud. Second, we apply the developed reference model on FermiCloud and compare calculated VM launching overhead values based on the model with measured overhead values on FermiCloud. Our empirical results on FermiCloud indicate that the developed reference model is accurate. We believe, with the guidance of the developed reference model, efficient resource allocation algorithms can be developed for cloud bursting process to minimize the operational cost and resource waste.

From A-4 to Explorer 1. [U.S. rocket and missile technology historical review

Science.gov (United States)

Debus, K. H.

1973-01-01

Historical review of the development of rocket and missile technology in the United States over the period from 1945 to 1958. Attention is given to the organization of activities, the launch facilities, and the scope of test rocket firings at the White Sands Proving Ground area during the initial phase of research with captured German V2 rockets. The development of the Redstone missiles is outlined by discussing aspects of military involvement, cooperation with industrial suppliers, details of ground support equipment, and results of initial test firings. Subsequent development of the Jupiter missiles is examined in a similar manner, and attention is given to activities involved in the launching of the Explorer 1 satellite.

A Real Time Differential GPS Tracking System for NASA Sounding Rockets

Science.gov (United States)

Bull, Barton; Bauer, Frank (Technical Monitor)

2000-01-01

Sounding rockets are suborbital launch vehicles capable of carrying scientific payloads to several hundred miles in altitude. These missions return a variety of scientific data including: chemical makeup and physical processes taking place in the atmosphere, natural radiation surrounding the Earth, data on the Sun, stars, galaxies and many other phenomena. In addition, sounding rockets provide a reasonably economical means of conducting engineering tests for instruments and devices to be used on satellites and other spacecraft prior to their use in these more expensive missions. Typically around thirty of these rockets are launched each year, from established ranges at Wallops Island, Virginia; Poker Flat Research Range, Alaska; White Sands Missile Range, New Mexico and from a number of ranges outside the United States. Many times launches are conducted from temporary launch ranges in remote parts of the world requiring considerable expense to transport and operate tracking radars. In order to support these missions, an inverse differential GPS system has been developed. The flight system consists of a small, inexpensive receiver, a preamplifier and a wrap-around antenna. A rugged, compact, portable ground station extracts GPS data from the raw payload telemetry stream, performs a real time differential solution and graphically displays the rocket's path relative to a predicted trajectory plot. In addition to generating a real time navigation solution, the system has been used for payload recovery, timing, data timetagging, precise tracking of multiple payloads and slaving of optical tracking systems for over the horizon acquisition. This paper discusses, in detail, the flight and ground hardware, as well as data processing and operational aspects of the system, and provides evidence of the system accuracy.

Plasma waves observed by sounding rockets

International Nuclear Information System (INIS)

Kimura, I.

1977-01-01

Observations of plasma wave phenomena have been conducted with several rockets launched at Kagoshima Space Center, Kyushu, Japan, and at Showa Base, Antarctica. This report presents some results of the observations in anticipation of having valuable comments from other plasma physicists, especially from those who are concerned with laboratory plasma. In the K-9M-41 rocket experiment, VLF plasma waves were observed. In this experiment, the electron beam of several tens of uA was emitted from a hot cathode when a positive dc bias changing from 0 to 10V at 1V interval each second was applied to a receiving dipole antenna. The discrete emissions with 'U' shaped frequency spectrum were observed for the dc bias over 3 volts. The U emissions appeared twice per spin period of the rocket. Similar rocket experiment was performed at Showa Base using a loop and dipole antenna and without hot cathode. Emissions were observed with varying conditions. At present, the authors postulate that such emissions may be produced just in the vicinity of a rocket due to a kind of wake effect. (Aoki, K.)

Method for Producing Launch/Landing Pads and Structures Project

Science.gov (United States)

Mueller, Robert P. (Compiler)

2015-01-01

Current plans for deep space exploration include building landing-launch pads capable of withstanding the rocket blast of much larger spacecraft that that of the Apollo days. The proposed concept will develop lightweight launch and landing pad materials from in-situ materials, utilizing regolith to produce controllable porous cast metallic foam brickstiles shapes. These shapes can be utilized to lay a landing launch platform, as a construction material or as more complex parts of mechanical assemblies.

Self-absorption theory applied to rocket measurements of the nitric oxide (1, 0) gamma band in the daytime thermosphere

Science.gov (United States)

Eparvier, F. G.; Barth, C. A.

1992-01-01

Observations of the UV fluorescent emissions of the NO (1, 0) and (0, 1) gamma bands in the lower-thermospheric dayglow, made with a sounding rocket launched on March 7, 1989 from Poker Flat, Alaska, were analyzed. The resonant (1, 0) gamma band was found to be attenuated below an altitude of about 120 km. A self-absorption model based on Holstein transmission functions was developed for the resonant (1, 0) gamma band under varying conditions of slant column density and temperature and was applied for the conditions of the rocket flight. The results of the model agreed with the measured attenuation of the band, indicating the necessity of including self-absorption theory in the analysis of satellite and rocket limb data of NO.

The Role of CFD Simulation in Rocket Propulsion Support Activities

Science.gov (United States)

West, Jeff

2011-01-01

Outline of the presentation: CFD at NASA/MSFC (1) Flight Projects are the Customer -- No Science Experiments (2) Customer Support (3) Guiding Philosophy and Resource Allocation (4) Where is CFD at NASA/MSFC? Examples of the expanding Role of CFD at NASA/MSFC (1) Liquid Rocket Engine Applications : Evolution from Symmetric and Steady to 3D Unsteady (2)Launch Pad Debris Transport-> Launch Pad Induced Environments (a) STS and Launch Pad Geometry-steady (b) Moving Body Shuttle Launch Simulations (c) IOP and Acoustics Simulations (3)General Purpose CFD Applications (4) Turbomachinery Applications

Modeling the Virtual Machine Launching Overhead under Fermicloud

Energy Technology Data Exchange (ETDEWEB)

Garzoglio, Gabriele [Fermilab; Wu, Hao [Fermilab; Ren, Shangping [IIT, Chicago; Timm, Steven [Fermilab; Bernabeu, Gerard [Fermilab; Noh, Seo-Young [KISTI, Daejeon

2014-11-12

FermiCloud is a private cloud developed by the Fermi National Accelerator Laboratory for scientific workflows. The Cloud Bursting module of the FermiCloud enables the FermiCloud, when more computational resources are needed, to automatically launch virtual machines to available resources such as public clouds. One of the main challenges in developing the cloud bursting module is to decide when and where to launch a VM so that all resources are most effectively and efficiently utilized and the system performance is optimized. However, based on FermiCloud’s system operational data, the VM launching overhead is not a constant. It varies with physical resource (CPU, memory, I/O device) utilization at the time when a VM is launched. Hence, to make judicious decisions as to when and where a VM should be launched, a VM launch overhead reference model is needed. The paper is to develop a VM launch overhead reference model based on operational data we have obtained on FermiCloud and uses the reference model to guide the cloud bursting process.

Numerical study for flame deflector design of a space launch vehicle

Science.gov (United States)

Oh, Hwayoung; Lee, Jungil; Um, Hyungsik; Huh, Hwanil

2017-04-01

A flame deflector is a structure that prevents damage to a launch vehicle and a launch pad due to exhaust plumes of a lifting-off launch vehicle. The shape of a flame deflector should be designed to restrain the discharged gas from backdraft inside the deflector and to reflect the impact to the surrounding environment and the engine characteristics of the vehicle. This study presents the five preliminary flame deflector configurations which are designed for the first-stage rocket engine of the Korea Space Launch Vehicle-II and surroundings of the Naro space center. The gas discharge patterns of the designed flame deflectors are investigated using the 3D flow field analysis by assuming that the air, in place of the exhaust gas, forms the plume. In addition, a multi-species unreacted flow model is investigated through 2D analysis of the first-stage engine of the KSLV-II. The results indicate that the closest Mach number and temperature distributions to the reacted flow model can be achieved from the 4-species unreacted flow model which employs H2O, CO2, and CO and specific heat-corrected plume.

NASA's Space Launch System Development Status

Science.gov (United States)

Lyles, Garry

2014-01-01

Development of the National Aeronautics and Space Administration's (NASA's) Space Launch System (SLS) heavy lift rocket is shifting from the formulation phase into the implementation phase in 2014, a little more than 3 years after formal program establishment. Current development is focused on delivering a vehicle capable of launching 70 metric tons (t) into low Earth orbit. This "Block 1" configuration will launch the Orion Multi-Purpose Crew Vehicle (MPCV) on its first autonomous flight beyond the Moon and back in December 2017, followed by its first crewed flight in 2021. SLS can evolve to a130t lift capability and serve as a baseline for numerous robotic and human missions ranging from a Mars sample return to delivering the first astronauts to explore another planet. Benefits associated with its unprecedented mass and volume include reduced trip times and simplified payload design. Every SLS element achieved significant, tangible progress over the past year. Among the Program's many accomplishments are: manufacture of core stage test barrels and domes; testing of Solid Rocket Booster development hardware including thrust vector controls and avionics; planning for RS- 25 core stage engine testing; and more than 4,000 wind tunnel runs to refine vehicle configuration, trajectory, and guidance. The Program shipped its first flight hardware - the Multi-Purpose Crew Vehicle Stage Adapter (MSA) - to the United Launch Alliance for integration with the Delta IV heavy rocket that will launch an Orion test article in 2014 from NASA's Kennedy Space Center. The Program successfully completed Preliminary Design Review in 2013 and will complete Key Decision Point C in 2014. NASA has authorized the Program to move forward to Critical Design Review, scheduled for 2015 and a December 2017 first launch. The Program's success to date is due to prudent use of proven technology, infrastructure, and workforce from the Saturn and Space Shuttle programs, a streamlined management

The National Aeronautics and Space Administration (NASA)/Goddard Space Flight Center (GSFC) sounding-rocket program

Science.gov (United States)

Guidotti, J. G.

1976-01-01

An overall introduction to the NASA sounding rocket program as managed by the Goddard Space Flight Center is presented. The various sounding rockets, auxiliary systems (telemetry, guidance, etc.), launch sites, and services which NASA can provide are briefly described.

Space Launch System Base Heating Test: Experimental Operations & Results

Science.gov (United States)

Dufrene, Aaron; Mehta, Manish; MacLean, Matthew; Seaford, Mark; Holden, Michael

2016-01-01

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 based on the Space Shuttle base heating tests with several improvements including doubling of the maximum chamber pressures and duplication of freestream conditions. Test methodology and conditions are presented, and base heating results from 76 runs are reported in non-dimensional form. Regions of high heating are identified and comparisons of various configuration and conditions are highlighted. Base pressure and radiometer results are also reported.

Preliminary results of rocket attitude and auroral green line emission rate in the DELTA campaign

Science.gov (United States)

Iwagami, Naomoto; Komada, Sayaka; Takahashi, Takao

2006-09-01

The attitude of a sounding rocket launched in the DELTA (Dynamics and Energetics of the Lower Thermosphere in Aurora) campaign was determined with IR horizon sensors and geomagnetic sensors. Since the payload was separated into two portions, two sets of attitude sensors were needed. A new IR sensor was developed for the present experiment, and found the zenith-angle of the spin-axis of the rocket with an accuracy of 2°. By combining information obtained by both type of sensors, the absolute attitudes were determined. The auroral green line emission rate was measured by a photometer on board the same rocket launched under active auroral conditions, and the energy flux of the auroral particle precipitation was estimated.

High Altitude Launch for a Practical SSTO

Science.gov (United States)

Landis, Geoffrey A.; Denis, Vincent

2003-01-01

Existing engineering materials allow the constuction of towers to heights of many kilometers. Orbital launch from a high altitude has significant advantages over sea-level launch due to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. High-altitude launch sites are particularly advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2% of the initial launch mass. An earlier paper enumerated some of the advantages of high altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25 kilometer altitudes above sea level. The performance increase can be directly translated into increased payload capability to orbit, ranging from 5 to 20% increase in the mass to orbit. For a candidate vehicle with an initial payload fraction of 2% of gross lift-off weight, this corresponds to 31% increase in payload (for 5-km launch altitude) to 122% additional payload (for 25-km launch altitude).

Ablative Material Testing at Lewis Rocket Lab

Science.gov (United States)

1997-01-01

The increasing demand for a low-cost, reliable way to launch commercial payloads to low- Earth orbit has led to the need for inexpensive, expendable propulsion systems for new launch vehicles. This, in turn, has renewed interest in less complex, uncooled rocket engines that have combustion chambers and exhaust nozzles fabricated from ablative materials. A number of aerospace propulsion system manufacturers have utilized NASA Lewis Research Center's test facilities with a high degree of success to evaluate candidate materials for application to new propulsion devices.

Sounding rocket experiments during the IMS period at Syowa Station, Antarctica

International Nuclear Information System (INIS)

Hirasawa, T.; Nagata, T.

1979-01-01

During IMS Period, 19 sounding rockets were launched into auroras at various stages of polar substorms from Syowa Station (Geomag. lat. = -69.6 0 , Geomag. log. = 77.1 0 ), Antarctica. Through the successful rocket flights, the significant physical quantities in auroras were obtained: 19 profiles of electron density and temperature, 11 energy spectra of precipitating electrons, 15 frequency spectra of VLF and HF plasma waves and 4 vertical profiles of electric and magnetic fields. These rocket data have been analyzed and compared with the coordinated ground-based observation data for studies of polar substorms. (author)

Feasibility and Performance of the Microwave Thermal Rocket Launcher

OpenAIRE

Parkin, Kevin L. G.; Culick, Fred E. C.

2004-01-01

Beamed-energy launch concepts employing a microwave thermal thruster are feasible in principle, and microwave sources of sufficient power to launch tons into LEO already exist. Microwave thermal thrusters operate on an analogous principle to nuclear thermal thrusters, which have experimentally demonstrated specific impulses exceeding 850 seconds. Assuming such performance, simple application of the rocket equation suggests that payload fractions of 10% are possible for a single stage to orbit...

Water-cooled spacecraft : DART to be launched by Russian Volna (Stingray) rocket

NARCIS (Netherlands)

Van Baten, T.; Buursink, J.; Hartmann, L.

2002-01-01

A25 September 2005, Barents Sea, near Murmansk.Ten metres under the surface of the sea, the launch tube of the Mstislav, a Rostropovich class nuclear submarine, grinds open. The countdown for the launch of a Volna R-29R slbm (Submarine-Launched Ballistic Missile) starts: For many years, satellites

On the coordination of EISCAT measurements with rocket and satellite observations

International Nuclear Information System (INIS)

Hultqvist, B.

1977-01-01

The scientific interest of combining EISCAT measurements of the thermal ionospheric plasma with sounding rocket and/or satellite measurements of the hot plasma distribution function and other variables is discussed briefly. Some examples are presented where such coordinated measurements are of great interest. The importance of being able to launch rockets through, or at least quite close to, the radar beam is emphasized. (Auth.)

High Resolution Modeling of the Thermospheric Response to Energy Inputs During the RENU-2 Rocket Flight

Science.gov (United States)

Walterscheid, R. L.; Brinkman, D. G.; Clemmons, J. H.; Hecht, J. H.; Lessard, M.; Fritz, B.; Hysell, D. L.; Clausen, L. B. N.; Moen, J.; Oksavik, K.; Yeoman, T. K.

2017-12-01

The Earth's magnetospheric cusp provides direct access of energetic particles to the thermosphere. These particles produce ionization and kinetic (particle) heating of the atmosphere. The increased ionization coupled with enhanced electric fields in the cusp produces increased Joule heating and ion drag forcing. These energy inputs cause large wind and temperature changes in the cusp region. The Rocket Experiment for Neutral Upwelling -2 (RENU-2) launched from Andoya, Norway at 0745UT on 13 December 2015 into the ionosphere-thermosphere beneath the magnetic cusp. It made measurements of the energy inputs (e.g., precipitating particles, electric fields) and the thermospheric response to these energy inputs (e.g., neutral density and temperature, neutral winds). Complementary ground based measurements were made. In this study, we use a high resolution two-dimensional time-dependent non hydrostatic nonlinear dynamical model driven by rocket and ground based measurements of the energy inputs to simulate the thermospheric response during the RENU-2 flight. Model simulations will be compared to the corresponding measurements of the thermosphere to see what they reveal about thermospheric structure and the nature of magnetosphere-ionosphere-thermosphere coupling in the cusp. Acknowledgements: This material is based upon work supported by the National Aeronautics and Space Administration under Grants: NNX16AH46G and NNX13AJ93G. This research was also supported by The Aerospace Corporation's Technical Investment program

Subscale Winged Rocket Development and Application to Future Reusable Space Transportation

Directory of Open Access Journals (Sweden)

Koichi YONEMOTO

2018-03-01

Full Text Available Kyushu Institute of Technology has been studying unmanned suborbital winged rocket called WIRES (WInged REusable Sounding rocket and its research subjects concerning aerodynamics, NGC (Navigation, Guidance and Control, cryogenic composite tanks etc., and conducting flight demonstration of small winged rocket since 2005. WIRES employs the original aerodynamic shape of HIMES (HIghly Maneuverable Experimental Sounding rocket studied by ISAS (Institute of Space and Astronautical Science of JAXA (Japan Aerospace Exploration Agency in 1980s. This paper presents the preliminary design of subscale non-winged and winged rockets called WIRES#013 and WIRES#015, respectively, that are developed in collaboration with JAXA, USC (University of Southern California, UTEP (University of Texas at El Paso and Japanese industries. WIRES#013 is a conventional pre-test rocket propelled by two IPA-LOX (Isopropyl Alcohol and Liquid Oxygen engines under development by USC. It has the total length of 4.6m, and the weight of 1000kg to reach the altitude of about 6km. The flight objective is validation of the telemetry and ground communication system, recovery parachute system, and launch operation of liquid engine. WIRES#015, which has the same length of WIRES#013 and the weight of 1000kg, is a NGC technology demonstrator propelled by a fully expander-cycle LOX-Methane engine designed and developed by JAXA to reach the altitude more than 6km. The flight tests of both WIRES#013 and WIRES#015 will be conducted at the launch facility of FAR (Friends of Amateur Rocketry, Inc., which is located at Mojave Desert of California in United States of America, in May 2018 and March 2019 respectively. After completion of WIRES#015 flight tests, the suborbital demonstrator called WIRES-X will be developed and its first flight test well be performed in 2020. Its application to future fully reusable space transportation systems, such as suborbital space tour vehicles and two

THE STERN PROJECT–HANDS ON ROCKETS SCIENCE FOR UNIVERSITY STUDENT

OpenAIRE

Schüttauf, Katharina; Stamminger, Andreas; Lappöhn, Karsten

2017-01-01

In April 2012, the German Aerospace Center DLR initiated a sponsorship program for university students to develop, build and launch their own rockets over a period of three years. The program designation STERN was abbreviated from the German “STudentische Experimental-RaketeN”, which translates to Student- Experimental-Rockets. The primary goal of the STERN program is to inspire students in the subject of space transportation through hands-on activities within a pro...

Multivariable optimization of liquid rocket engines using particle swarm algorithms

Science.gov (United States)

Jones, Daniel Ray

Liquid rocket engines are highly reliable, controllable, and efficient compared to other conventional forms of rocket propulsion. As such, they have seen wide use in the space industry and have become the standard propulsion system for launch vehicles, orbit insertion, and orbital maneuvering. Though these systems are well understood, historical optimization techniques are often inadequate due to the highly non-linear nature of the engine performance problem. In this thesis, a Particle Swarm Optimization (PSO) variant was applied to maximize the specific impulse of a finite-area combustion chamber (FAC) equilibrium flow rocket performance model by controlling the engine's oxidizer-to-fuel ratio and de Laval nozzle expansion and contraction ratios. In addition to the PSO-controlled parameters, engine performance was calculated based on propellant chemistry, combustion chamber pressure, and ambient pressure, which are provided as inputs to the program. The performance code was validated by comparison with NASA's Chemical Equilibrium with Applications (CEA) and the commercially available Rocket Propulsion Analysis (RPA) tool. Similarly, the PSO algorithm was validated by comparison with brute-force optimization, which calculates all possible solutions and subsequently determines which is the optimum. Particle Swarm Optimization was shown to be an effective optimizer capable of quick and reliable convergence for complex functions of multiple non-linear variables.

The Launch Systems Operations Cost Model

Science.gov (United States)

Prince, Frank A.; Hamaker, Joseph W. (Technical Monitor)

2001-01-01

One of NASA's primary missions is to reduce the cost of access to space while simultaneously increasing safety. A key component, and one of the least understood, is the recurring operations and support cost for reusable launch systems. In order to predict these costs, NASA, under the leadership of the Independent Program Assessment Office (IPAO), has commissioned the development of a Launch Systems Operations Cost Model (LSOCM). LSOCM is a tool to predict the operations & support (O&S) cost of new and modified reusable (and partially reusable) launch systems. The requirements are to predict the non-recurring cost for the ground infrastructure and the recurring cost of maintaining that infrastructure, performing vehicle logistics, and performing the O&S actions to return the vehicle to flight. In addition, the model must estimate the time required to cycle the vehicle through all of the ground processing activities. The current version of LSOCM is an amalgamation of existing tools, leveraging our understanding of shuttle operations cost with a means of predicting how the maintenance burden will change as the vehicle becomes more aircraft like. The use of the Conceptual Operations Manpower Estimating Tool/Operations Cost Model (COMET/OCM) provides a solid point of departure based on shuttle and expendable launch vehicle (ELV) experience. The incorporation of the Reliability and Maintainability Analysis Tool (RMAT) as expressed by a set of response surface model equations gives a method for estimating how changing launch system characteristics affects cost and cycle time as compared to today's shuttle system. Plans are being made to improve the model. The development team will be spending the next few months devising a structured methodology that will enable verified and validated algorithms to give accurate cost estimates. To assist in this endeavor the LSOCM team is part of an Agency wide effort to combine resources with other cost and operations professionals to

Rocket in situ observation of equatorial plasma irregularities in the region between E and F layers over Brazil

Directory of Open Access Journals (Sweden)

S. Savio Odriozola

2017-03-01

Full Text Available A two-stage VS-30 Orion rocket was launched from the equatorial rocket launching station in Alcântara, Brazil, on 8 December 2012 soon after sunset (19:00 LT, carrying a Langmuir probe operating alternately in swept and constant bias modes. At the time of launch, ground equipment operated at equatorial stations showed rapid rise in the base of the F layer, indicating the pre-reversal enhancement of the F region vertical drift and creating ionospheric conditions favorable for the generation of plasma bubbles. Vertical profiles of electron density estimated from Langmuir probe data showed wave patterns and small- and medium-scale plasma irregularities in the valley region (100–300 km during the rocket upleg and downleg. These irregularities resemble those detected by the very high frequency (VHF radar installed at Jicamarca and so-called equatorial quasi-periodic echoes. We present evidence suggesting that these observations could be the first detection of this type of irregularity made by instruments onboard a rocket.

Soyuz Spacecraft Transported to Launch Pad

Science.gov (United States)

2003-01-01

The Soyuz TMA-3 spacecraft and its booster rocket (rear view) is shown on a rail car for transport to the launch pad where it was raised to a vertical launch position at the Baikonur Cosmodrome, Kazakhstan on October 16, 2003. Liftoff occurred on October 18th, transporting a three man crew to the International Space Station (ISS). Aboard were Michael Foale, Expedition-8 Commander and NASA science officer; Alexander Kaleri, Soyuz Commander and flight engineer, both members of the Expedition-8 crew; and European Space agency (ESA) Astronaut Pedro Duque of Spain. Photo Credit: 'NASA/Bill Ingalls'

Space Launch System Spacecraft and Payload Elements: Progress Toward Crewed Launch and Beyond

Science.gov (United States)

Schorr, Andrew A.; Smith, David Alan; Holcomb, Shawn; Hitt, David

2017-01-01

While significant and substantial progress continues to be accomplished toward readying the Space Launch System (SLS) rocket for its first test flight, work is already underway on preparations for the second flight - using an upgraded version of the vehicle - and beyond. Designed to support human missions into deep space, SLS is the most powerful human-rated launch vehicle the United States has ever undertaken, and is one of three programs being managed by the National Aeronautics and Space Administration's (NASA's) Exploration Systems Development division. The Orion spacecraft program is developing a new crew vehicle that will support human missions beyond low Earth orbit (LEO), and the Ground Systems Development and Operations (GSDO) program is transforming Kennedy Space Center (KSC) into a next-generation spaceport capable of supporting not only SLS but also multiple commercial users. Together, these systems will support human exploration missions into the proving ground of cislunar space and ultimately to Mars. For its first flight, SLS will deliver a near-term heavy-lift capability for the nation with its 70-metric-ton (t) Block 1 configuration. Each element of the vehicle now has flight hardware in production in support of the initial flight of the SLS, which will propel Orion around the moon and back. Encompassing hardware qualification, structural testing to validate hardware compliance and analytical modeling, progress is on track to meet the initial targeted launch date. In Utah and Mississippi, booster and engine testing are verifying upgrades made to proven shuttle hardware. At Michoud Assembly Facility (MAF) in Louisiana, the world's largest spacecraft welding tool is producing tanks for the SLS core stage. Providing the Orion crew capsule/launch vehicle interface and in-space propulsion via a cryogenic upper stage, the Spacecraft/Payload Integration and Evolution (SPIE) element serves a key role in achieving SLS goals and objectives. The SPIE element

A Suborbital Spaceship for Short Duration Space and Microsat Launch

OpenAIRE

Bahn, Pat

2005-01-01

The TGV Rockets corporation is working on a small Vertical Takeoff Vertical Landing Suborbital Rocketship capable of carrying 1000 kg to 100 km for low cost. This provides unique and interesting capabilities for payload test and qualification, development and short duration experimentation. Theoretical possibilities include micro-sat launch. TGV Rockets was founded in 1997 on a desire to commercialize the Delta Clipper-Experimental (DC-X)1,5,8. Subsequently TGV has been working towards th...

Computational study of variable area ejector rocket flowfields

Science.gov (United States)

Etele, Jason

Access to space has always been a scientific priority for countries which can afford the prohibitive costs associated with launch. However, the large scale exploitation of space by the business community will require the cost of placing payloads into orbit be dramatically reduced for space to become a truly profitable commodity. To this end, this work focuses on a next generation propulsive technology called the Rocket Based Combined Cycle (RBCC) engine in which rocket, ejector, ramjet, and scramjet cycles operate within the same engine environment. Using an in house numerical code solving the axisymmetric version of the Favre averaged Navier Stokes equations (including the Wilcox ko turbulence model with dilatational dissipation) a systematic study of various ejector designs within an RBCC engine is undertaken. It is shown that by using a central rocket placed along the axisymmetric axis in combination with an annular rocket placed along the outer wall of the ejector, one can obtain compression ratios of approximately 2.5 for the case where both the entrained air and rocket exhaust mass flows are equal. Further, it is shown that constricting the exit area, and the manner in which this constriction is performed, has a significant positive impact on the compression ratio. For a decrease in area of 25% a purely conical ejector can increase the compression ratio by an additional 23% compared to an equal length unconstricted ejector. The use of a more sharply angled conical section followed by a cylindrical section to maintain equivalent ejector lengths can further increase the compression ratio by 5--7% for a total increase of approximately 30%.

Rockets for Extended Source Soft X-ray Spectroscopy

Science.gov (United States)

McEntaffer, Randall

The soft X-ray background surrounds our local galactic environment yet very little is known about the physical characteristics of this plasma. A high-resolution spectrum could unlock the properties of this million degree gas but the diffuse, low intensity nature of the background have made it difficult to observe, especially with a dispersive spectrograph. Previous observations have relied on X-ray detector energy resolution which produces poorly defined spectra that are poorly fit by complex plasma models. Here we propose a series of suborbital rocket flights that will begin the characterization of this elusive source through high-resolution X-ray grating spectroscopy. The rocket-based spectrograph can resolve individual emission lines over the soft X-ray band and place tight constraints on the temperature, density, abundance, ionization state and age of the plasma. These payloads will draw heavily from the heritage gained from previous rocket missions, while also benefiting from related NASA technology development programs. The Pennsylvania State University (PSU) team has a history of designing and flying spectrometer components onboard rockets while also being scientific leaders in the field of diffuse soft X-ray astronomy. The PSU program will provide hands-on training of young scientists in the techniques of instrumental and observational X-ray astronomy. The proposed rocket program will also expose these researchers to a full experiment cycle: design, fabrication, tolerance analysis, assembly, flight-qualification, calibration, integration, launch, and data analysis; using a combination of technologies suitable for adaptation to NASA's major missions. The PSU program in suborbital X-ray astronomy represents an exciting mix of compelling science, heritage, cutting-edge technology development, and training of future scientists.

Launch vehicle selection model

Science.gov (United States)

Montoya, Alex J.

1990-01-01

Over the next 50 years, humans will be heading for the Moon and Mars to build scientific bases to gain further knowledge about the universe and to develop rewarding space activities. These large scale projects will last many years and will require large amounts of mass to be delivered to Low Earth Orbit (LEO). It will take a great deal of planning to complete these missions in an efficient manner. The planning of a future Heavy Lift Launch Vehicle (HLLV) will significantly impact the overall multi-year launching cost for the vehicle fleet depending upon when the HLLV will be ready for use. It is desirable to develop a model in which many trade studies can be performed. In one sample multi-year space program analysis, the total launch vehicle cost of implementing the program reduced from 50 percent to 25 percent. This indicates how critical it is to reduce space logistics costs. A linear programming model has been developed to answer such questions. The model is now in its second phase of development, and this paper will address the capabilities of the model and its intended uses. The main emphasis over the past year was to make the model user friendly and to incorporate additional realistic constraints that are difficult to represent mathematically. We have developed a methodology in which the user has to be knowledgeable about the mission model and the requirements of the payloads. We have found a representation that will cut down the solution space of the problem by inserting some preliminary tests to eliminate some infeasible vehicle solutions. The paper will address the handling of these additional constraints and the methodology for incorporating new costing information utilizing learning curve theory. The paper will review several test cases that will explore the preferred vehicle characteristics and the preferred period of construction, i.e., within the next decade, or in the first decade of the next century. Finally, the paper will explore the interaction

A Concept of Two-Stage-To-Orbit Reusable Launch Vehicle

Science.gov (United States)

Yang, Yong; Wang, Xiaojun; Tang, Yihua

2002-01-01

Reusable Launch Vehicle (RLV) has a capability of delivering a wide rang of payload to earth orbit with greater reliability, lower cost, more flexibility and operability than any of today's launch vehicles. It is the goal of future space transportation systems. Past experience on single stage to orbit (SSTO) RLVs, such as NASA's NASP project, which aims at developing an rocket-based combined-cycle (RBCC) airplane and X-33, which aims at developing a rocket RLV, indicates that SSTO RLV can not be realized in the next few years based on the state-of-the-art technologies. This paper presents a concept of all rocket two-stage-to-orbit (TSTO) reusable launch vehicle. The TSTO RLV comprises an orbiter and a booster stage. The orbiter is mounted on the top of the booster stage. The TSTO RLV takes off vertically. At the altitude about 50km the booster stage is separated from the orbiter, returns and lands by parachutes and airbags, or lands horizontally by means of its own propulsion system. The orbiter continues its ascent flight and delivers the payload into LEO orbit. After completing orbit mission, the orbiter will reenter into the atmosphere, automatically fly to the ground base and finally horizontally land on the runway. TSTO RLV has less technology difficulties and risk than SSTO, and maybe the practical approach to the RLV in the near future.

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

Closure Letter Report for Corrective Action Unit 496: Buried Rocket Site - Antelope Lake (TTR)

International Nuclear Information System (INIS)

NSTec Environmental Restoration

2007-01-01

A Streamlined Approach for Environmental Restoration (SAFER) Plan for investigation and closure of CAU 496, Corrective Action Site (CAS) TA-55-008-TAAL (Buried Rocket), at the Tonopah Test Range (TTR), was approved by the Nevada Department of Environmental Protection (NDEP) on July 21,2004. Approval to transfer CAS TA-55-008-TAAL from CAU 496 to CAU 4000 (No Further Action Sites) was approved by NDEP on December 21, 2005, based on the assumption that the rocket did not present any environmental concern. The approval letter included the following condition: ''NDEP understands, from the NNSA/NSO letter dated November 30,2005, that a search will be conducted for the rocket during the planned characterization of other sites at the Tonopah Test Range and, if found, the rocket will be removed as a housekeeping measure''. NDEP and U.S. Department of Energy, National Nuclear Security Administration Nevada Site Office personnel located the rocket on Mid Lake during a site visit to TTR, and a request to transfer CAS TA-55-008-TAAL from CAU 4000 back to CAU 496 was approved by NDEP on September 11,2006. CAS TA-55-008-TAAL was added to the ''Federal Facility Agreement and Consent Order'' of 1996, based on an interview with a retired TTR worker in 1993. The original interview documented that a rocket was launched from Area 9 to Antelope Lake and was never recovered due to the high frequency of rocket tests being conducted during this timeframe. The interviewee recalled the rocket being an M-55 or N-55 (the M-50 ''Honest John'' rocket was used extensively at TTR from the 1960s to early 1980s). A review of previously conducted interviews with former TTR personnel indicated that the interviewees confused information from several sites. The location of the CAU 496 rocket on Mid Lake is directly south of the TTR rocket launch facility in Area 9 and is consistent with information gathered on the lost rocket during recent interviews. Most pertinently, an interview in 2005 with a

Hydrocarbon Rocket Technology Impact Forecasting

Science.gov (United States)

Stuber, Eric; Prasadh, Nishant; Edwards, Stephen; Mavris, Dimitri N.

2012-01-01

Forecasting method is a normative forecasting technique that allows the designer to quantify the effects of adding new technologies on a given design. This method can be used to assess and identify the necessary technological improvements needed to close the gap that exists between the current design and one that satisfies all constraints imposed on the design. The TIF methodology allows for more design knowledge to be brought to the earlier phases of the design process, making use of tools such as Quality Function Deployments, Morphological Matrices, Response Surface Methodology, and Monte Carlo Simulations.2 This increased knowledge allows for more informed decisions to be made earlier in the design process, resulting in shortened design cycle time. This paper will investigate applying the TIF method, which has been widely used in aircraft applications, to the conceptual design of a hydrocarbon rocket engine. In order to reinstate a manned presence in space, the U.S. must develop an affordable and sustainable launch capability. Hydrocarbon-fueled rockets have drawn interest from numerous major government and commercial entities because they offer a low-cost heavy-lift option that would allow for frequent launches1. However, the development of effective new hydrocarbon rockets would likely require new technologies in order to overcome certain design constraints. The use of advanced design methods, such as the TIF method, enables the designer to identify key areas in need of improvement, allowing one to dial in a proposed technology and assess its impact on the system. Through analyses such as this one, a conceptual design for a hydrocarbon-fueled vehicle that meets all imposed requirements can be achieved.

Dynamical Model of Rocket Propellant Loading with Liquid Hydrogen

Data.gov (United States)

National Aeronautics and Space Administration — A dynamical model describing the multi-stage process of rocket propellant loading has been developed. It accounts for both the nominal and faulty regimes of...

Using Discrete Event Simulation to Model Integrated Commodities Consumption for a Launch Campaign of the Space Launch System

Science.gov (United States)

Leonard, Daniel; Parsons, Jeremy W.; Cates, Grant

2014-01-01

In May 2013, NASA's GSDO Program requested a study to develop a discrete event simulation (DES) model that analyzes the launch campaign process of the Space Launch System (SLS) from an integrated commodities perspective. The scope of the study includes launch countdown and scrub turnaround and focuses on four core launch commodities: hydrogen, oxygen, nitrogen, and helium. Previously, the commodities were only analyzed individually and deterministically for their launch support capability, but this study was the first to integrate them to examine the impact of their interactions on a launch campaign as well as the effects of process variability on commodity availability. The study produced a validated DES model with Rockwell Arena that showed that Kennedy Space Center's ground systems were capable of supporting a 48-hour scrub turnaround for the SLS. The model will be maintained and updated to provide commodity consumption analysis of future ground system and SLS configurations.

The issue of ensuring the safe explosion of the spent orbital stages of a launch vehicle with propulsion rocket engine

Directory of Open Access Journals (Sweden)

Trushlyakov Valeriy I.

2017-01-01

Full Text Available A method for increasing the safe explosion of the spent orbital stages of a space launch vehicle (SLV with a propulsion rocket engine (PRE based on the gasification of unusable residues propellant and venting fuel tanks. For gasification and ventilation the hot gases used produced by combustion of the specially selected gas generating composition (GGC with a set of physical and chemical properties. Excluding the freezing of the drainage system on reset gasified products (residues propellant+pressurization gas+hot gases in the near-Earth space is achieved by selecting the physical-chemical characteristics of the GGC. Proposed steps to ensure rotation of gasified products due to dumping through the drainage system to ensure the most favorable conditions for propellant gasification residues. For example, a tank with liquid oxygen stays with the orbital spent second stage of the SLV “Zenit”, which shows the effectiveness of the proposed method.

Examination of Cross-Scale Coupling During Auroral Events using RENU2 and ISINGLASS Sounding Rocket Data.

Science.gov (United States)

Kenward, D. R.; Lessard, M.; Lynch, K. A.; Hysell, D. L.; Hampton, D. L.; Michell, R.; Samara, M.; Varney, R. H.; Oksavik, K.; Clausen, L. B. N.; Hecht, J. H.; Clemmons, J. H.; Fritz, B.

2017-12-01

The RENU2 sounding rocket (launched from Andoya rocket range on December 13th, 2015) observed Poleward Moving Auroral Forms within the dayside cusp. The ISINGLASS rockets (launched from Poker Flat rocket range on February 22, 2017 and March 2, 2017) both observed aurora during a substorm event. Despite observing very different events, both campaigns witnessed a high degree of small scale structuring within the larger auroral boundary, including Alfvenic signatures. These observations suggest a method of coupling large-scale energy input to fine scale structures within aurorae. During RENU2, small (sub-km) scale drivers persist for long (10s of minutes) time scales and result in large scale ionospheric (thermal electron) and thermospheric response (neutral upwelling). ISINGLASS observations show small scale drivers, but with short (minute) time scales, with ionospheric response characterized by the flight's thermal electron instrument (ERPA). The comparison of the two flights provides an excellent opportunity to examine ionospheric and thermospheric response to small scale drivers over different integration times.

Rocket center Peenemünde — Personal memories

Science.gov (United States)

Dannenberg, Konrad; Stuhlinger, Ernst

Von Braun built his first rockets as a young teenager. At 14, he started making plans for rockets for human travel to the Moon and Mars. The German Army began a rocket program in 1929. Two years later, Colonel (later General) Becker contacted von Braun who experimented with rockets in Berlin, gave him a contract in 1932, and, jointly with the Air Force, in 1936 built the rocket center Peenemünde where von Braun and his team developed the A-4 (V-2) rocket under Army auspices, while the Air Force developed the V-1 (buzz bomb), wire-guided bombs, and rocket planes. Albert Speer, impressed by the work of the rocketeers, allowed a modest growth of the Peenemünde project; this brought Dannenberg to the von Braun team in 1940. Hitler did not believe in rockets; he ignored the A-4 project until 1942 when he began to support it, expecting that it could turn the fortunes of war for him. He drastically increased the Peenemünde work force and allowed the transfer of soldiers from the front to Peenemünde; that was when Stuhlinger, in 1943, came to Peenemünde as a Pfc.-Ph.D. Later that year, Himmler wrenched the authority over A-4 production out of the Army's hands, put it under his command, and forced production of the immature rocket at Mittelwerk, and its military deployment against targets in France, Belgium, and England. Throughout the development of the A-4 rocket, von Braun was the undisputed leader of the project. Although still immature by the end of the war, the A-4 had proceeded to a status which made it the first successful long-range precision rocket, the prototype for a large number of military rockets built by numerous nations after the war, and for space rockets that launched satellites and traveled to the Moon and the planets.

Global atmospheric response to emissions from a proposed reusable space launch system

Science.gov (United States)

Larson, Erik J. L.; Portmann, Robert W.; Rosenlof, Karen H.; Fahey, David W.; Daniel, John S.; Ross, Martin N.

2017-01-01

Modern reusable launch vehicle technology may allow high flight rate space transportation at low cost. Emissions associated with a hydrogen fueled reusable rocket system are modeled based on the launch requirements of developing a space-based solar power system that generates present-day global electric energy demand. Flight rates from 104 to 106 per year are simulated and sustained to a quasisteady state. For the assumed rocket engine, H2O and NOX are the primary emission products; this also includes NOX produced during reentry heating. For a base case of 105 flights per year, global stratospheric and mesospheric water vapor increase by approximately 10 and 100%, respectively. As a result, high-latitude cloudiness increases in the lower stratosphere and near the mesopause by as much as 20%. Increased water vapor also results in global effective radiative forcing of about 0.03 W/m2. NOX produced during reentry exceeds meteoritic production by more than an order of magnitude, and along with in situ stratospheric emissions, results in a 0.5% loss of the globally averaged ozone column, with column losses in the polar regions exceeding 2%.

PLUMEX II: A second set of coincident radar and rocket observations of equatorial spread-F

International Nuclear Information System (INIS)

Szuszczewicz, E.P.; Tsunoda, R.T.; Narcisi, R.; Holmes, J.C.

1981-01-01

PLUMEX II, the second rocket in a two-rocket operation that successfully executed coincident rocket and radar measurements of backscatter plumes and plasma depletions, was launched into the mid-phase of well-developed equatorial spread-F. In contrast with the first operation, the PLUMEX II results show large scale F-region irregularities only on the bottomside gradient with smaller scale irregularities (i.e., small scale structure imbedded in larger scale features) less intense than corresponding observations in PLUMEX I. The latter result could support current interpretations of east-west plume asymmetry which suggests that during initial upwelling the western wall of a plume (the PLUMEX I case) is more unstable than its eastern counterpart (the PLUMEX II case). In addition, ion mass spectrometer results are found to provide further support for an ion transport model which ''captures'' bottomside ions in an upwelling bubble and transports them to high altitudes

Multicamera High Dynamic Range High-Speed Video of Rocket Engine Tests and Launches

Data.gov (United States)

National Aeronautics and Space Administration — High-speed video recording of rocket engine tests has several challenges. The scenes that are imaged have both bright and dark regions associated with plume emission...

LV-IMLI: Integrated MLI/Aeroshell for Cryogenic Launch Vehicles, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Cryogenic propellants have the highest energy density of any rocket fuel, and are used in most NASA and commercial launch vehicles to power their ascent. Cryogenic...

Feasibility and Performance of the Microwave Thermal Rocket Launcher

Science.gov (United States)

Parkin, Kevin L. G.; Culick, Fred E. C.

2004-03-01

Beamed-energy launch concepts employing a microwave thermal thruster are feasible in principle, and microwave sources of sufficient power to launch tons into LEO already exist. Microwave thermal thrusters operate on an analogous principle to nuclear thermal thrusters, which have experimentally demonstrated specific impulses exceeding 850 seconds. Assuming such performance, simple application of the rocket equation suggests that payload fractions of 10% are possible for a single stage to orbit (SSTO) microwave thermal rocket. We present an SSTO concept employing a scaled X-33 aeroshell. The flat aeroshell underside is covered by a thin-layer microwave absorbent heat-exchanger that forms part of the thruster. During ascent, the heat-exchanger faces the microwave beam. A simple ascent trajectory analysis incorporating X-33 aerodynamic data predicts a 10% payload fraction for a 1 ton craft of this type. In contrast, the Saturn V had 3 non-reusable stages and achieved a payload fraction of 4%.

Lessons from half a century experience of Japanese solid rocketry since Pencil rocket

Science.gov (United States)

Matogawa, Yasunori

2007-12-01

50 years have passed since a tiny rocket "Pencil" was launched horizontally at Kokubunji near Tokyo in 1955. Though there existed high level of rocket technology in Japan before the end of the second World War, it was not succeeded by the country after the War. Pencil therefore was the substantial start of Japanese rocketry that opened the way to the present stage. In the meantime, a rocket group of the University of Tokyo contributed to the International Geophysical Year in 1957-1958 by developing bigger rockets, and in 1970, the group succeeded in injecting first Japanese satellite OHSUMI into earth orbit. It was just before the launch of OHSUMI that Japan had built up the double feature system of science and applications in space efforts. The former has been pursued by ISAS (the Institute of Space and Astronautical Science) of the University of Tokyo, and the latter by NASDA (National Space Development Agency). This unique system worked quite efficiently because space activities in scientific and applicational areas could develop rather independently without affecting each other. Thus Japan's space science ran up rapidly to the international stage under the support of solid propellant rocket technology, and, after a 20 year technological introduction period from the US, a big liquid propellant launch vehicle, H-II, at last was developed on the basis of Japan's own technology in the early 1990's. On October 1, 2003, as a part of Governmental Reform, three Japanese space agencies were consolidated into a single agency, JAXA (Japan Aerospace Exploration Agency), and Japan's space efforts began to walk toward the future in a globally coordinated fashion, including aeronautics, astronautics, space science, satellite technology, etc., at the same time. This paper surveys the history of Japanese rocketry briefly, and draws out the lessons from it to make a new history of Japan's space efforts more meaningful.

Intense auroral field-aligned currents and electrojets detected by rocket-borne fluxgate magnetometer

International Nuclear Information System (INIS)

Tohyama, Fumio; Fukunishi, Hiroshi; Takahashi, Takao; Kokubun, Susumu; Fujii, Ryoichi; Yamagishi, Hisao.

1988-01-01

The S-310JA-11 and S-310JA-12 rockets, having a vector magnetometer with high sensitivity (1.8 nT) and high sampling frequency (100 Hz), were launched into the aurora on May 29 and July 12, 1985, from Syowa Station, Antarctica. The S-310JA-11 rocket penetrated twice quiet arcs, while the S-310JA-12 rocket traversed across intense and active auroral arcs during a large magnetic substorm. In the S-310JA-12 rocket experiment, intense field-aligned currents of 400 - 600 nT were observed when the rocket penetrated an active arc during the descending flight. The magnetometer on board the S-310JA-12 rocket also detected intense electrojet currents with a center at 110 km on the upward leg and at 108 km on the downward leg. The magnetometer data of the S-310JA-11 rocket showed no distinguished magnetic field variation due to field-aligned current and electrojet. (author)

STS-37 Breakfast / Ingress / Launch & ISO Camera Views

Science.gov (United States)

1991-01-01

The primary objective of the STS-37 mission was to deploy the Gamma Ray Observatory. The mission was launched at 9:22:44 am on April 5, 1991, onboard the space shuttle Atlantis. The mission was led by Commander Steven Nagel. The crew was Pilot Kenneth Cameron and Mission Specialists Jerry Ross, Jay Apt, and Linda Godwing. This videotape shows the crew having breakfast on the launch day, with the narrator introducing them. It then shows the crew's final preparations and the entry into the shuttle, while the narrator gives information about each of the crew members. The countdown and launch is shown including the shuttle separation from the solid rocket boosters. The launch is reshown from 17 different camera views. Some of the other camera views were in black and white.

Design and analysis of a single stage to orbit nuclear thermal rocket reactor engine

Energy Technology Data Exchange (ETDEWEB)

Labib, Satira, E-mail: Satira.Labib@duke-energy.com; King, Jeffrey, E-mail: kingjc@mines.edu

2015-06-15

Graphical abstract: - Highlights: • Three NTR reactors are optimized for the single stage launch of 1–15 MT payloads. • The proposed rocket engines have specific impulses in excess of 700 s. • Reactivity and submersion criticality requirements are satisfied for each reactor. - Abstract: Recent advances in the development of high power density fuel materials have renewed interest in nuclear thermal rockets (NTRs) as a viable propulsion technology for future space exploration. This paper describes the design of three NTR reactor engines designed for the single stage to orbit launch of payloads from 1 to 15 metric tons. Thermal hydraulic and rocket engine analyses indicate that the proposed rocket engines are able to reach specific impulses in excess of 800 s. Neutronics analyses performed using MCNP5 demonstrate that the hot excess reactivity, shutdown margin, and submersion criticality requirements are satisfied for each NTR reactor. The reactors each consist of a 40 cm diameter core packed with hexagonal tungsten cermet fuel elements. The core is surrounded by radial and axial beryllium reflectors and eight boron carbide control drums. The 40 cm long reactor meets the submersion criticality requirements (a shutdown margin of at least $1 subcritical in all submersion scenarios) with no further modifications. The 80 and 120 cm long reactors include small amounts of gadolinium nitride as a spectral shift absorber to keep them subcritical upon submersion in seawater or wet sand following a launch abort.

Design and analysis of a single stage to orbit nuclear thermal rocket reactor engine

International Nuclear Information System (INIS)

Labib, Satira; King, Jeffrey

2015-01-01

Graphical abstract: - Highlights: • Three NTR reactors are optimized for the single stage launch of 1–15 MT payloads. • The proposed rocket engines have specific impulses in excess of 700 s. • Reactivity and submersion criticality requirements are satisfied for each reactor. - Abstract: Recent advances in the development of high power density fuel materials have renewed interest in nuclear thermal rockets (NTRs) as a viable propulsion technology for future space exploration. This paper describes the design of three NTR reactor engines designed for the single stage to orbit launch of payloads from 1 to 15 metric tons. Thermal hydraulic and rocket engine analyses indicate that the proposed rocket engines are able to reach specific impulses in excess of 800 s. Neutronics analyses performed using MCNP5 demonstrate that the hot excess reactivity, shutdown margin, and submersion criticality requirements are satisfied for each NTR reactor. The reactors each consist of a 40 cm diameter core packed with hexagonal tungsten cermet fuel elements. The core is surrounded by radial and axial beryllium reflectors and eight boron carbide control drums. The 40 cm long reactor meets the submersion criticality requirements (a shutdown margin of at least $1 subcritical in all submersion scenarios) with no further modifications. The 80 and 120 cm long reactors include small amounts of gadolinium nitride as a spectral shift absorber to keep them subcritical upon submersion in seawater or wet sand following a launch abort

The Potential for Ozone Depletion in Solid Rocket Motor Plumes by Heterogeneous Chemistry

National Research Council Canada - National Science Library

Hanning-Lee, M

1996-01-01

... (hydroxylated alumina), respectively, over the temperature range -60 to 200 degrees C. This work addresses the potential for stratospheric ozone depletion by launch vehicle solid rocket motor exhaust...

Launch Pad Escape System Design (Human Spaceflight)

Science.gov (United States)

Maloney, Kelli

2011-01-01

A launch pad escape system for human spaceflight is one of those things that everyone hopes they will never need but is critical for every manned space program. Since men were first put into space in the early 1960s, the need for such an Emergency Escape System (EES) has become apparent. The National Aeronautics and Space Administration (NASA) has made use of various types of these EESs over the past 50 years. Early programs, like Mercury and Gemini, did not have an official launch pad escape system. Rather, they relied on a Launch Escape System (LES) of a separate solid rocket motor attached to the manned capsule that could pull the astronauts to safety in the event of an emergency. This could only occur after hatch closure at the launch pad or during the first stage of flight. A version of a LES, now called a Launch Abort System (LAS) is still used today for all manned capsule type launch vehicles. However, this system is very limited in that it can only be used after hatch closure and it is for flight crew only. In addition, the forces necessary for the LES/LAS to get the capsule away from a rocket during the first stage of flight are quite high and can cause injury to the crew. These shortcomings led to the development of a ground based EES for the flight crew and ground support personnel as well. This way, a much less dangerous mode of egress is available for any flight or ground personnel up to a few seconds before launch. The early EESs were fairly simple, gravity-powered systems to use when thing's go bad. And things can go bad very quickly and catastrophically when dealing with a flight vehicle fueled with millions of pounds of hazardous propellant. With this in mind, early EES designers saw such a passive/unpowered system as a must for last minute escapes. This and other design requirements had to be derived for an EES, and this section will take a look at the safety design requirements had to be derived for an EES, and this section will take a look at

Design of a rocket-borne radiometer for stratospheric ozone measurements

International Nuclear Information System (INIS)

Barnes, R.A.; Simeth, P.G.

1989-01-01

A four-filter ultraviolet radiometer for measuring stratospheric ozone is described. The payload is launched aboard a Super-Loki rocket to an apogee of 70 km. The instrument measures the solar ultraviolet irradiance over its filter wavelengths as it descends on a parachute. The amount of ozone in the path between the radiometer and the sun is calculated from the attenuation of solar flux using the Beer-Lambert law. Radar at the launch site measures the height of the instrument throughout its flight. The fundamental ozone value measured by the ROCOZ-A radiometer is the vertical ozone overburden as a function of geometric altitude. Ozone measurements are obtained for altitudes from 55 to 20 km, extending well above the altitude range of balloon-borne ozone-measuring instruments. The optics and electronics in the radiometer have been designed within relatively severe size and weight limitations imposed by the launch vehicle. The electronics in the improved rocket ozonesonde (ROCOZ-A) provide essentially drift-free outputs throughout 40-min ozone soundings at stratospheric temperatures. The modest cost of the payload precludes recovery and makes the instrument a versatile tool compared to larger ozonesondes

Five-Segment Solid Rocket Motor Development Status

Science.gov (United States)

Priskos, Alex S.

2012-01-01

In support of the National Aeronautics and Space Administration (NASA), Marshall Space Flight Center (MSFC) is developing a new, more powerful solid rocket motor for space launch applications. To minimize technical risks and development costs, NASA chose to use the Space Shuttle s solid rocket boosters as a starting point in the design and development. The new, five segment motor provides a greater total impulse with improved, more environmentally friendly materials. To meet the mass and trajectory requirements, the motor incorporates substantial design and system upgrades, including new propellant grain geometry with an additional segment, new internal insulation system, and a state-of-the art avionics system. Significant progress has been made in the design, development and testing of the propulsion, and avionics systems. To date, three development motors (one each in 2009, 2010, and 2011) have been successfully static tested by NASA and ATK s Launch Systems Group in Promontory, UT. These development motor tests have validated much of the engineering with substantial data collected, analyzed, and utilized to improve the design. This paper provides an overview of the development progress on the first stage propulsion system.

Rocket borne solar eclipse experiment to measure the temperature structure of the solar corona via lyman-α line profile observations

International Nuclear Information System (INIS)

Argo, H.V.

1981-01-01

A rocket borne experiment to measure the temperature structure of the inner solar corona via the doppler broadening of the resonance hydrogen Lyman-α (lambda1216A) radiation scattered by ambient neutral hydrogen atoms was attempted during the 16 Feb 1980 solar eclipse. Two Nike-Black Brant V sounding rockets carrying instrumented payloads were launched into the path of the advancing eclipse umbra from the San Marco satellite launch platform 3 miles off the east coast of Kenya

Comet mission is blown off course by faulty rocket

CERN Multimedia

Henderson, M

2003-01-01

ESA has announced that the launch of the Rosetta probe, has been delayed indefinitely because of problems with the unreliable Ariane-5 rocket on which it was to fly. The project will now have to be redesigned completely to target a different comet at a later date (1/2 page).

Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

Science.gov (United States)

Morris, Christopher I.

2005-01-01

Pulse detonation rocket engines (PDREs) have generated research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional rocket engines. The detonative mode of combustion employed by these devices offers a thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional rocket engines and gas turbines. However, while this theoretical advantage has spurred considerable interest in building PDRE devices, the unsteady blowdown process intrinsic to the PDRE has made realistic estimates of the actual propulsive performance problematic. The recent review article by Kailasanath highlights some of the progress that has been made in comparing the available experimental measurements with analytical and numerical models. In recent work by the author, a quasi-one-dimensional, finite rate chemistry CFD model was utilized to study the gasdynamics and performance characteristics of PDREs over a range of blowdown pressure ratios from 1-1000. Models of this type are computationally inexpensive, and enable first-order parametric studies of the effect of several nozzle and extension geometries on PDRE performance over a wide range of conditions. However, the quasi-one-dimensional approach is limited in that it cannot properly capture the multidimensional blast wave and flow expansion downstream of the PDRE, nor can it resolve nozzle flow separation if present. Moreover, the previous work was limited to single-pulse calculations. In this paper, an axisymmetric finite rate chemistry model is described and utilized to study these issues in greater detail. Example Mach number contour plots showing the multidimensional blast wave and nozzle exhaust plume are shown. The performance results are compared with the quasi-one-dimensional results from the previous paper. Both Euler and Navier-Stokes solutions are calculated in order to determine the effect of viscous

Demand-type gas supply system for rocket borne thin-window proportional counters

Science.gov (United States)

Acton, L. W.; Caravalho, R.; Catura, R. C.; Joki, E. G.

1977-01-01

A simple closed loop control system has been developed to maintain the gas pressure in thin-window proportional counters during rocket flights. This system permits convenient external control of detector pressure and system flushing rate. The control system is activated at launch with the sealing of a reference volume at the existing system pressure. Inflight control to plus or minus 2 torr at a working pressure of 760 torr has been achieved on six rocket flights.

Sensitivity Analysis of Launch Vehicle Debris Risk Model

Science.gov (United States)

Gee, Ken; Lawrence, Scott L.

2010-01-01

As part of an analysis of the loss of crew risk associated with an ascent abort system for a manned launch vehicle, a model was developed to predict the impact risk of the debris resulting from an explosion of the launch vehicle on the crew module. The model consisted of a debris catalog describing the number, size and imparted velocity of each piece of debris, a method to compute the trajectories of the debris and a method to calculate the impact risk given the abort trajectory of the crew module. The model provided a point estimate of the strike probability as a function of the debris catalog, the time of abort and the delay time between the abort and destruction of the launch vehicle. A study was conducted to determine the sensitivity of the strike probability to the various model input parameters and to develop a response surface model for use in the sensitivity analysis of the overall ascent abort risk model. The results of the sensitivity analysis and the response surface model are presented in this paper.

Studies of an extensively axisymmetric rocket based combined cycle (RBCC) engine powered single-stage-to-orbit (SSTO) vehicle

Science.gov (United States)

Foster, Richard W.; Escher, William J. D.; Robinson, John W.

1989-01-01

The present comparative performance study has established that rocket-based combined cycle (RBCC) propulsion systems, when incorporated by essentially axisymmetric SSTO launch vehicle configurations whose conical forebody maximizes both capture-area ratio and total capture area, are capable of furnishing payload-delivery capabilities superior to those of most multistage, all-rocket launchers. Airbreathing thrust augmentation in the rocket-ejector mode of an RBCC powerplant is noted to make a major contribution to final payload capability, by comparison to nonair-augmented rocket engine propulsion systems.

Nonlinear acoustic propagation of launch vehicle and military jet aircraft noise

Science.gov (United States)

Gee, Kent L.

2010-10-01

The noise from launch vehicles and high-performance military jet aircraft has been shown to travel nonlinearly as a result of an amplitude-dependent speed of sound. Because acoustic pressure compressions travel faster than rarefactions, the waveform steepens and shocks form. This process results in a very different (and readily audible) noise signature and spectrum than predicted by linear models. On-going efforts to characterize the nonlinearity using statistical and spectral measures are described with examples from recent static tests of solid rocket boosters and the F-22 Raptor.

Rationales for the Lightning Launch Commit Criteria

Science.gov (United States)

Willett, John C. (Editor); Merceret, Francis J. (Editor); Krider, E. Philip; O'Brien, T. Paul; Dye, James E.; Walterscheid, Richard L.; Stolzenburg, Maribeth; Cummins, Kenneth; Christian, Hugh J.; Madura, John T.

2016-01-01

Since natural and triggered lightning are demonstrated hazards to launch vehicles, payloads, and spacecraft, NASA and the Department of Defense (DoD) follow the Lightning Launch Commit Criteria (LLCC) for launches from Federal Ranges. The LLCC were developed to prevent future instances of a rocket intercepting natural lightning or triggering a lightning flash during launch from a Federal Range. NASA and DoD utilize the Lightning Advisory Panel (LAP) to establish and develop robust rationale from which the criteria originate. The rationale document also contains appendices that provide additional scientific background, including detailed descriptions of the theory and observations behind the rationales. The LLCC in whole or part are used across the globe due to the rigor of the documented criteria and associated rationale. The Federal Aviation Administration (FAA) adopted the LLCC in 2006 for commercial space transportation and the criteria were codified in the FAA's Code of Federal Regulations (CFR) for Safety of an Expendable Launch Vehicle (Appendix G to 14 CFR Part 417, (G417)) and renamed Lightning Flight Commit Criteria in G417.

Launch and Landing Effects Ground Operations (LLEGO) Model

Science.gov (United States)

2008-01-01

LLEGO is a model for understanding recurring launch and landing operations costs at Kennedy Space Center for human space flight. Launch and landing operations are often referred to as ground processing, or ground operations. Currently, this function is specific to the ground operations for the Space Shuttle Space Transportation System within the Space Shuttle Program. The Constellation system to follow the Space Shuttle consists of the crewed Orion spacecraft atop an Ares I launch vehicle and the uncrewed Ares V cargo launch vehicle. The Constellation flight and ground systems build upon many elements of the existing Shuttle flight and ground hardware, as well as upon existing organizations and processes. In turn, the LLEGO model builds upon past ground operations research, modeling, data, and experience in estimating for future programs. Rather than to simply provide estimates, the LLEGO model s main purpose is to improve expenses by relating complex relationships among functions (ground operations contractor, subcontractors, civil service technical, center management, operations, etc.) to tangible drivers. Drivers include flight system complexity and reliability, as well as operations and supply chain management processes and technology. Together these factors define the operability and potential improvements for any future system, from the most direct to the least direct expenses.

A Monte Carlo Approach to Modeling the Breakup of the Space Launch System EM-1 Core Stage with an Integrated Blast and Fragment Catalogue

Science.gov (United States)

Richardson, Erin; Hays, M. J.; Blackwood, J. M.; Skinner, T.

2014-01-01

The Liquid Propellant Fragment Overpressure Acceleration Model (L-FOAM) is a tool developed by Bangham Engineering Incorporated (BEi) that produces a representative debris cloud from an exploding liquid-propellant launch vehicle. Here it is applied to the Core Stage (CS) of the National Aeronautics and Space Administration (NASA) Space Launch System (SLS launch vehicle). A combination of Probability Density Functions (PDF) based on empirical data from rocket accidents and applicable tests, as well as SLS specific geometry are combined in a MATLAB script to create unique fragment catalogues each time L-FOAM is run-tailored for a Monte Carlo approach for risk analysis. By accelerating the debris catalogue with the BEi blast model for liquid hydrogen / liquid oxygen explosions, the result is a fully integrated code that models the destruction of the CS at a given point in its trajectory and generates hundreds of individual fragment catalogues with initial imparted velocities. The BEi blast model provides the blast size (radius) and strength (overpressure) as probabilities based on empirical data and anchored with analytical work. The coupling of the L-FOAM catalogue with the BEi blast model is validated with a simulation of the Project PYRO S-IV destruct test. When running a Monte Carlo simulation, L-FOAM can accelerate all catalogues with the same blast (mean blast, 2 s blast, etc.), or vary the blast size and strength based on their respective probabilities. L-FOAM then propagates these fragments until impact with the earth. Results from L-FOAM include a description of each fragment (dimensions, weight, ballistic coefficient, type and initial location on the rocket), imparted velocity from the blast, and impact data depending on user desired application. LFOAM application is for both near-field (fragment impact to escaping crew capsule) and far-field (fragment ground impact footprint) safety considerations. The user is thus able to use statistics from a Monte Carlo

Reusable Rocket Engine Advanced Health Management System. Architecture and Technology Evaluation: Summary

Science.gov (United States)

Pettit, C. D.; Barkhoudarian, S.; Daumann, A. G., Jr.; Provan, G. M.; ElFattah, Y. M.; Glover, D. E.

1999-01-01

In this study, we proposed an Advanced Health Management System (AHMS) functional architecture and conducted a technology assessment for liquid propellant rocket engine lifecycle health management. The purpose of the AHMS is to improve reusable rocket engine safety and to reduce between-flight maintenance. During the study, past and current reusable rocket engine health management-related projects were reviewed, data structures and health management processes of current rocket engine programs were assessed, and in-depth interviews with rocket engine lifecycle and system experts were conducted. A generic AHMS functional architecture, with primary focus on real-time health monitoring, was developed. Fourteen categories of technology tasks and development needs for implementation of the AHMS were identified, based on the functional architecture and our assessment of current rocket engine programs. Five key technology areas were recommended for immediate development, which (1) would provide immediate benefits to current engine programs, and (2) could be implemented with minimal impact on the current Space Shuttle Main Engine (SSME) and Reusable Launch Vehicle (RLV) engine controllers.

INTELSAT III LIFTS OFF FROM LC 17A ABOARD A DELTA LAUNCH VEHICLE

Science.gov (United States)

1968-01-01

A Delta launch vehicle carrying the Intelsat III spacecraft was launched from Complex 17 at 8:09 p.m. EDT. A malfunction in flight resulted in the rocket breaking up some 102 seconds into the mission. Destruct action was initiated by the Air Force East Test Range some six seconds later when it was apparent that the mission could not succeed.

VSB-30 sounding rocket: history of flight performance

Directory of Open Access Journals (Sweden)

Wolfgang Jung

2011-09-01

Full Text Available The VSB-30 vehicle is a two-stage, unguided, rail launched sounding rocket, consisting of two solid propellant motors, payload, with recovery and service system. By the end of 2010, ten vehicles had already been launched, three from Brazil (Alcântara and seven from Sweden (Esrange. The objective of this paper is to give an overview of the main characteristics of the first ten flights of the VSB-30, with emphasis on performance and trajectory data. The circular 3σ dispersion area for payload impact point has around 50 km of radius. In most launchings of such vehicle, the impact of the payload fell within 2 sigma. This provides the possibility for further studies to decrease the area of dispersion from the impact point.

Simultaneous measurements of auroral particles and electric currents by a rocket-borne instrument system - Introductory remarks

Science.gov (United States)

Anderson, H. R.; Cloutier, P. A.

1975-01-01

A rocket-borne experiment package has been designed to obtain simultaneous in situ measurements of the pitch angle distributions and energy spectra of primary auroral particles, the flux of neutral hydrogen at auroral energies, the electric currents flowing in the vicinity of the auroral arc as determined from vector magnetic data, and the modulation of precipitating electrons in the frequency range 0.5-10 MHz. The experiment package was launched by a Nike-Tomahawk rocket from Poker Flat, Alaska, at 0722 UT on Feb. 25, 1972, over a bright auroral band. This paper is intended to serve as an introduction to the detailed discussion of results given in the companion papers. As such it includes a brief review of the general problem, a discussion of the rocket instrumentation, a delineation of the auroral and geomagnetic conditions at the time of launch, and comments on the overall payload performance.

Simultaneous measurements of auroral particles and electric currents by a rocket-borne instrument system: introductory remarks

International Nuclear Information System (INIS)

Anderson, H.R.; Cloutier, P.A.

1975-01-01

A rocket-borne experiment package has been designed to obtain simultaneous in situ measurements of the pitch angle distribution and energy spectra of primary auroral particles, the flux of neutral hydrogen at auroral energies, the electric currents flowing in the vicinity of the auroral arc as determined from vector magnetic data, and the modulation of precipitating electrons in the frequency range 0.5-10 MHz. The experiment package was launched by a Nike-Tomahawk rocket from Poker Flat, Alaska, at 0722 UT on February 25, 1972, over a bright auroral band. This paper is intended to serve as an introduction to the detailed discussion of results given in the companion papers. As such it includes a brief review of the general problem, a discussion of the rocket instrumentation, a delineation of the auroral and geomagnetic conditions at the time of launch, and comments on the overall payload performance

Rocket Based Combined Cycle (RBCC) Engine

Science.gov (United States)

2004-01-01

Pictured is an artist's concept of the Rocket Based Combined Cycle (RBCC) launch. The RBCC's overall objective is to provide a technology test bed to investigate critical technologies associated with opperational usage of these engines. The program will focus on near term technologies that can be leveraged to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsions systems and ultimately a Single Stage To Orbit (SSTO) air breathing propulsion system.

Improved Background Removal in Sounding Rocket Neutral Atom Imaging Data

Science.gov (United States)

Smith, M. R.; Rowland, D. E.

2017-12-01

The VISIONS sounding rocket, launched into a substorm on Feb 7, 2013 from Poker Flat, Alaska had a novel miniaturized energetic neutral atom (ENA) imager onboard. We present further analysis of the ENA data from this rocket flight, including improved removal of ultraviolet and electron contamination. In particular, the relative error source contributions due to geocoronal, auroral, and airglow UV, as well as energetic electrons from 10 eV to 3 keV were assessed. The resulting data provide a more clear understanding of the spatial and temporal variations of the ion populations that are energized to tens or hundreds of eV.

SSTO rockets. A practical possibility

Science.gov (United States)

Bekey, Ivan

1994-07-01

Most experts agree that single-stage-to-orbit (SSTO) rockets would become feasible if more advanced technologies were available to reduce the vehicle dry weight, increase propulsion system performance, or both. However, these technologies are usually judged to be very ambitious and very far off. This notion persists despite major advances in technology and vehicle design in the past decade. There appears to be four major misperceptions about SSTOs, regarding their mass fraction, their presumed inadequate performance margin, their supposedly small payloads, and their extreme sensitivity to unanticipated vehicle weight growth. These misperceptions can be dispelled for SSTO rockets using advanced technologies that could be matured and demonstrated in the near term. These include a graphite-composite primary structure, graphite-composite and Al-Li propellant tanks with integral reusable thermal protection, long-life tripropellant or LOX-hydrogen engines, and several technologies related to operational effectiveness, including vehicle health monitoring, autonomous avionics/flight control, and operable launch and ground handling systems.

Designing on-Board Data Handling for EDF (Electric Ducted Fan) Rocket

Science.gov (United States)

Mulyana, A.; Faiz, L. A. A.

2018-02-01

The EDF (Electric Ducted Fan) rocket to launch requires a system of monitoring, tracking and controlling to allow the rocket to glide properly. One of the important components in the rocket is OBDH (On-Board Data Handling) which serves as a medium to perform commands and data processing. However, TTC (Telemetry, Tracking, and Command) are required to communicate between GCS (Ground Control Station) and OBDH on EDF rockets. So the design control system of EDF rockets and GCS for telemetry and telecommand needs to be made. In the design of integrated OBDH controller uses a lot of electronics modules, to know the behavior of rocket used IMU sensor (Inertial Measurement Unit) in which consist of 3-axis gyroscope sensor and Accelerometer 3-axis. To do tracking using GPS, compass sensor as a determinant of the direction of the rocket as well as a reference point on the z-axis of gyroscope sensor processing and used barometer sensors to measure the height of the rocket at the time of glide. The data can be known in real-time by sending data through radio modules at 2.4 GHz frequency using XBee-Pro S2B to GCS. By using windows filter, noises can be reduced, and it used to guarantee monitoring and controlling system can work properly.

Low-Cost Propellant Launch From a Tethered Balloon

Science.gov (United States)

Wilcox, Brian

2006-01-01

A document presents a concept for relatively inexpensive delivery of propellant to a large fuel depot in low orbit around the Earth, for use in rockets destined for higher orbits, the Moon, and for remote planets. The propellant is expected to be at least 85 percent of the mass needed in low Earth orbit to support the NASA Exploration Vision. The concept calls for the use of many small ( 10 ton) spin-stabilized, multistage, solid-fuel rockets to each deliver 250 kg of propellant. Each rocket would be winched up to a balloon tethered above most of the atmospheric mass (optimal altitude 26 2 km). There, the rocket would be aimed slightly above the horizon, spun, dropped, and fired at a time chosen so that the rocket would arrive in orbit near the depot. Small thrusters on the payload (powered, for example, by boil-off gases from cryogenic propellants that make up the payload) would precess the spinning rocket, using data from a low-cost inertial sensor to correct for small aerodynamic and solid rocket nozzle misalignment torques on the spinning rocket; would manage the angle of attack and the final orbit insertion burn; and would be fired on command from the depot in response to observations of the trajectory of the payload so as to make small corrections to bring the payload into a rendezvous orbit and despin it for capture by the depot. The system is low-cost because the small rockets can be mass-produced using the same techniques as those to produce automobiles and low-cost munitions, and one or more can be launched from a U.S. territory on the equator (Baker or Jarvis Islands in the mid-Pacific) to the fuel depot on each orbit (every 90 minutes, e.g., any multiple of 6,000 per year).

Injection of a microsatellite in circular orbits using a three-stage launch vehicle

Science.gov (United States)

Marchi, L. O.; Murcia, J. O.; Prado, A. F. B. A.; Solórzano, C. R. H.

2017-10-01

The injection of a satellite into orbit is usually done by a multi-stage launch vehicle. Nowadays, the space market demonstrates a strong tendency towards the use of smaller satellites, because the miniaturization of the systems improve the cost/benefit of a mission. A study to evaluate the capacity of the Brazilian Microsatellite Launch Vehicle (VLM) to inject payloads into Low Earth Orbits is presented in this paper. All launches are selected to be made to the east side of the Alcântara Launch Center (CLA). The dynamical model to calculate the trajectory consists of the three degrees of freedom (3DOF) associated with the translational movement of the rocket. Several simulations are performed according to a set of restrictions imposed to the flight. The altitude reached in the separation of the second stage, the altitude and velocity of injection, the flight path angle at the moment of the activation of the third stage and the duration of the ballistic flight are presented as a function of the payload carried.

Launch Environment Water Flow Simulations Using Smoothed Particle Hydrodynamics

Science.gov (United States)

Vu, Bruce T.; Berg, Jared J.; Harris, Michael F.; Crespo, Alejandro C.

2015-01-01

This paper describes the use of Smoothed Particle Hydrodynamics (SPH) to simulate the water flow from the rainbird nozzle system used in the sound suppression system during pad abort and nominal launch. The simulations help determine if water from rainbird nozzles will impinge on the rocket nozzles and other sensitive ground support elements.

Polar 5, a Norwegian US electron accelerator sounding rocket

International Nuclear Information System (INIS)

Jacobsen, T.A.; Maehlum, B.N.; Troeim, J.

1976-01-01

A technical description of a mother daughter experiment including an electron gun is given. The payload was launched by a Nike/Tomahawk rocket from Andenes, North-Norway near 2030 local time on February 1, 1976. A few preliminary observations obtained by the HF-wave propagation experiment, the retarding potential analyzer and the energetic electron counters are be presented

Reusable Launch Vehicle Technology Program

Science.gov (United States)

Freeman, Delma C., Jr.; Talay, Theodore A.; Austin, R. Eugene

1997-01-01

Industry/NASA reusable launch vehicle (RLV) technology program efforts are underway to design, test, and develop technologies and concepts for viable commercial launch systems that also satisfy national needs at acceptable recurring costs. Significant progress has been made in understanding the technical challenges of fully reusable launch systems and the accompanying management and operational approaches for achieving a low cost program. This paper reviews the current status of the RLV technology program including the DC-XA, X-33 and X-34 flight systems and associated technology programs. It addresses the specific technologies being tested that address the technical and operability challenges of reusable launch systems including reusable cryogenic propellant tanks, composite structures, thermal protection systems, improved propulsion and subsystem operability enhancements. The recently concluded DC-XA test program demonstrated some of these technologies in ground and flight test. Contracts were awarded recently for both the X-33 and X-34 flight demonstrator systems. The Orbital Sciences Corporation X-34 flight test vehicle will demonstrate an air-launched reusable vehicle capable of flight to speeds of Mach 8. The Lockheed-Martin X-33 flight test vehicle will expand the test envelope for critical technologies to flight speeds of Mach 15. A propulsion program to test the X-33 linear aerospike rocket engine using a NASA SR-71 high speed aircraft as a test bed is also discussed. The paper also describes the management and operational approaches that address the challenge of new cost effective, reusable launch vehicle systems.

Validation and Simulation of Ares I Scale Model Acoustic Test - 2 - Simulations at 5 Foot Elevation for Evaluation of Launch Mount Effects

Science.gov (United States)

Strutzenberg, Louise L.; Putman, Gabriel C.

2011-01-01

The Ares I Scale Model Acoustics Test (ASMAT) is a series of live-fire tests of scaled rocket motors meant to simulate the conditions of the Ares I launch configuration. These tests have provided a well documented set of high fidelity measurements useful for validation including data taken over a range of test conditions and containing phenomena like Ignition Over-Pressure and water suppression of acoustics. Expanding from initial simulations of the ASMAT setup in a held down configuration, simulations have been performed using the Loci/CHEM computational fluid dynamics software for ASMAT tests of the vehicle at 5 ft. elevation (100 ft. real vehicle elevation) with worst case drift in the direction of the launch tower. These tests have been performed without water suppression and have compared the acoustic emissions for launch structures with and without launch mounts. In addition, simulation results have also been compared to acoustic and imagery data collected from similar live-fire tests to assess the accuracy of the simulations. Simulations have shown a marked change in the pattern of emissions after removal of the launch mount with a reduction in the overall acoustic environment experienced by the vehicle and the formation of highly directed acoustic waves moving across the platform deck. Comparisons of simulation results to live-fire test data showed good amplitude and temporal correlation and imagery comparisons over the visible and infrared wavelengths showed qualitative capture of all plume and pressure wave evolution features.

The Advanced Solid Rocket Motor

Science.gov (United States)

Mitchell, Royce E.

1992-01-01

The Advanced Solid Rocket Motor will utilize improved design features and automated manufacturing methods to produce an inherently safer propulsive system for the Space Shuttle and future launch systems. This second-generation motor will also provide an additional 12,000 pounds of payload to orbit, enhancing the utility and efficiency of the Shuttle system. The new plant will feature strip-wound, asbestos-free insulation; propellant continuous mixing and casting; and extensive robotic systems. Following a series of static tests at the Stennis Space Center, MS flights are targeted to begin in early 1997.

Spectroscopic observation of the middle ultraviolet earth albedo by S-520-4 rocket and mesospheric ozone density profile

International Nuclear Information System (INIS)

Suzuki, Katsuhisa; Ogawa, Toshihiro.

1982-01-01

The ozone Hartey absorption band in the middle ultraviolet range is commonly adopted for the ozone measurement by rocket and satellite observations. In Japan, since 1965 the ozone absorption in the solar ultraviolet radiation has been observed by rocket-borne uv photometers. On the other hand the spectroscopic measurements of the scattered solar ultraviolet radiation from the terrestrial atmosphere will be performed by the EXOS-C satellite which will be launched in 1984. We tested the spectrometer for this satellite experiment by S-520-4 rocket launched on 5 September 1981. This instrument observed the scattered radiation of 2500 A -- 3300 A and the visible earth albedo of 4030 A. The spectrometer is consisted of a concave grating and has about 10 A wavelength resolution. A photomultiplier having a Cs-Te photocathode is used as a uv detector. The visible albedo is measured by a photometer consisting of an interference filter and a phototube. We estimated the atmospheric ozone profile, comparing the uv spectrum obtained by this experiment with the model calculations. The estimated ozone density profile higher than 30 km altitude has good agreement with the profile obtained by the previous uv photometer experiments at Uchinoura. There are differences between the observed spectrum and the calculated one in = 3100 A. We can explain them by the effect of Mie scattering and the uv stray light. In the present experiment we could successfully test the functions of the instrument in the space. rocket, spectrometer, solar ultraviolet radiation, earth albedo, ozone (author)

The international trade in launch services : the effects of U.S. laws, policies and practices on its development

NARCIS (Netherlands)

Fenema, van H.P.

1999-01-01

Rockets or launch vehicles, though sharing the same technology, have both military and civil applications: they can be used as missiles or as 'ordinary' transportation vehicles. As a consequence, national security and foreign policy considerations stand in the way of the international launch

The techniques of quality operations computational and experimental researches of the launch vehicles in the drawing-board stage

Science.gov (United States)

Rozhaeva, K.

2018-01-01

The aim of the researchis the quality operations of the design process at the stage of research works on the development of active on-Board system of the launch vehicles spent stages descent with liquid propellant rocket engines by simulating the gasification process of undeveloped residues of fuel in the tanks. The design techniques of the gasification process of liquid rocket propellant components residues in the tank to the expense of finding and fixing errors in the algorithm calculation to increase the accuracy of calculation results is proposed. Experimental modelling of the model liquid evaporation in a limited reservoir of the experimental stand, allowing due to the false measurements rejection based on given criteria and detected faults to enhance the results reliability of the experimental studies; to reduce the experiments cost.

Lockheed Martin approach to a Reusable Launch Vehicle (RLV)

Science.gov (United States)

Elvin, John D.

1996-03-01

This paper discusses Lockheed Martin's perspective on the development of a cost effective Reusable Launch Vehicle (RLV). Critical to a successful Single Stage To Orbit (SSTO) program are; an economic development plan sensitive to fiscal constraints; a vehicle concept satisfying present and future US launch needs; and an operations concept commensurate with a market driven program. Participation in the economic plan by government, industry, and the commercial sector is a key element of integrating our development plan and funding profile. The RLV baseline concept design, development evolution and several critical trade studies illustrate the superior performance achieved by our innovative approach to the problem of SSTO. Findings from initial aerodynamic and aerothermodynamic wind tunnel tests and trajectory analyses on this concept confirm the superior characteristics of the lifting body shape combined with the Linear Aerospike rocket engine. This Aero Ballistic Rocket (ABR) concept captures the essence of The Skunk Works approach to SSTO RLV technology integration and system engineering. These programmatic and concept development topics chronicle the key elements to implementing an innovative market driven next generation RLV.

A new sounding rocket payload for solar plasma studies

Science.gov (United States)

Bruner, Marilyn E.; Brown, William A.; Appert, Kevin L.

1989-01-01

A sounding rocket payload developed for studies of high-temperature plasmas associated with solar active regions and flares is described. The payload instruments will record both spectra and images in the UV, EUV, and soft X-ray regions of the spectrum. The instruments, including the Dual Range Spectrograph, the Flat Field Soft X-ray Spectrograph, the Normal Incidence Soft X-ray Imager, the UV Filtergraph, and the H-alpha Imaging system, are described. Attention is also given to the new structural system of the payload, based on a large optical table suspended within the payload cavity, which will support the optical elements in their correct positions and orientations and will maintain these alignments throughout the rocket launch environment.

The electromagnetic rocket gun - a means to reach ultrahigh velocities

International Nuclear Information System (INIS)

Winterberg, F.

1983-01-01

A novel kind of electromagnetic launcher for the acceleration of multigram-size macroparticles, up to velocities required for impact fusion, is proposed. The novel launcher concept combines the efficiency of a gun with the much higher velocities attainable by a rocket. In the proposed concept a rocket-like projectile is launched inside a gun barrel, drawing its energy from a travelling magnetic wave. The travelling magnetic wave heats and ionizes the exhaust jet of the rocket. As a result, the projectile i propelled both by the recoil from the jet and the magnetic pressure of the travelling magnetic wave. In comparison to magnetic linear accelerators, accelerating either superconducting or ferromagnetic projectiles, the proposed concept has several important advantages. First, the exhaust jet is much longer than the rocket-like projectile and which permits a much longer switching time to turn on the travelling magnetic wave. Second, the proposed concept does not require superconducting projectiles, or projectiles made from expensive ferromagnetic material. Third, unlike in railgun accelerators, the projectile can be kept away from the wall, and thereby can reach much larger velocities. (orig.)

MODELING A ROCKET ELASTIC STRUCTURE AS A BECK’S COLUMN UNDER FOLLOWER FORCE

OpenAIRE

Brejão, Leandro Forne; Brasil, Reyolando M. L. R. F.

2017-01-01

It is intended, in this paper, to develop a mathematical model of an elastic space rocket structure as a Beck’s column excited by a follower (or circulatory) force. This force represents the rocket motor thrust that should be always in the direction of the tangent to the structure deformed axis at the base of the vehicle. We present a simplified two degree of freedom rigid bars discrete model. Its system of two second order nonlinear ordinary differential equations of motion are derived via L...

Dynamic Beam Solutions for Real-Time Simulation and Control Development of Flexible Rockets

Science.gov (United States)

Su, Weihua; King, Cecilia K.; Clark, Scott R.; Griffin, Edwin D.; Suhey, Jeffrey D.; Wolf, Michael G.

2016-01-01

In this study, flexible rockets are structurally represented by linear beams. Both direct and indirect solutions of beam dynamic equations are sought to facilitate real-time simulation and control development for flexible rockets. The direct solution is completed by numerically integrate the beam structural dynamic equation using an explicit Newmark-based scheme, which allows for stable and fast transient solutions to the dynamics of flexile rockets. Furthermore, in the real-time operation, the bending strain of the beam is measured by fiber optical sensors (FOS) at intermittent locations along the span, while both angular velocity and translational acceleration are measured at a single point by the inertial measurement unit (IMU). Another study in this paper is to find the analytical and numerical solutions of the beam dynamics based on the limited measurement data to facilitate the real-time control development. Numerical studies demonstrate the accuracy of these real-time solutions to the beam dynamics. Such analytical and numerical solutions, when integrated with data processing and control algorithms and mechanisms, have the potential to increase launch availability by processing flight data into the flexible launch vehicle's control system.

Rocket observation of electron density irregularities in the lower E region

International Nuclear Information System (INIS)

Watanabe, Yuzo; Nakamura, Yoshiharu; Amemiya, Hiroshi.

1990-01-01

Local ionospheric electron density irregularities in the scale size of 3 m to 300 m have been measured on the ascending path from 74 km to 93 km by a fix biased Langmuir probe on board the S-310-16 sounding rocket. The rocket was launched at 22:40:00 on February 1, 1986 from Kagoshima Space Center in Japan. It is found from frequency analysis of the data that the spectral index of the irregularities is 0.9 to 1.8 and the irregularity amplitude is 1 to 15 %. The altitude where the amplitude reaches its maximum is 88 km. The generation mechanism of these irregularities is explained by the neutral turbulence theory, which indicates that the spectral index is 5/3 and has been confirmed by a chemical release experiment using rockets over India to be valid up to about 110 km. From frequency analysis of the data observed during the descent in the lower E region, we have found that the rocket-wake effect becomes larger when the probe is situated near the edge of the rocket-wake, and that this is also the case even when the rocket-wake effect does not clearly appear in the DC current signal which approximately changes in proportion to the electron density, where the probe is completely situated inside the rocket-wake region. (author)

An introduction to the water recovery x-ray rocket

Science.gov (United States)

Miles, Drew M.; McEntaffer, Randall L.; Schultz, Ted B.; Donovan, Benjamin D.; Tutt, James H.; Yastishock, Daniel; Steiner, Tyler; Hillman, Christopher R.; McCoy, Jake A.; Wages, Mitchell; Hull, Sam; Falcone, Abe; Burrows, David N.; Chattopadhyay, Tanmoy; Anderson, Tyler; McQuaide, Maria

2017-08-01

The Water Recovery X-ray Rocket (WRXR) is a sounding rocket payload that will launch from the Kwajalein Atoll in April 2018 and seeks to be the first astrophysics sounding rocket payload to be water recovered by NASA. WRXR's primary instrument is a grating spectrometer that consists of a mechanical collimator, X-ray reflection gratings, grazing-incidence mirrors, and a hybrid CMOS detector. The instrument will obtain a spectrum of the diffuse soft X-ray emission from the northern part of the Vela supernova remnant and is optimized for 3rd and 4th order OVII emission. Utilizing a field of view of 3.25° × 3.25° and resolving power of λ/δλ ≍40-50 in the lines of interest, the WRXR spectrometer aims to achieve the most highly-resolved spectrum of Vela's diffuse soft X-ray emission. This paper presents introductions to the payload and the science target.

Human Health Risk Assessment Simulations in a Distributed Environment for Shuttle Launch

Science.gov (United States)

Thirumalainambi, Rajkumar; Bardina, Jorge

2004-01-01

During the launch of a rocket under prevailing weather conditions, commanders at Cape Canaveral Air Force station evaluate the possibility of whether wind blown toxic emissions might reach civilian and military personnel in the near by area. In our model, we focused mainly on Hydrogen chloride (HCL), Nitrogen oxides (NOx) and Nitric acid (HNO3), which are non-carcinogenic chemicals as per United States Environmental Protection Agency (USEPA) classification. We have used the hazard quotient model to estimate the number of people at risk. It is based on the number of people with exposure above a reference exposure level that is unlikely to cause adverse health effects. The risk to the exposed population is calculated by multiplying the individual risk and the number in exposed population. The risk values are compared against the acceptable risk values and GO or NO-go situation is decided based on risk values for the Shuttle launch. The entire model is simulated over the web and different scenaria can be generated which allows management to choose an optimum decision.

Current status of rocket developments in universities -development of a small hybrid rocket with a swirling oxidizer flow type engine

OpenAIRE

Yuasa, Saburo; Kitagawa, Koki

2005-01-01

To develop an experimental small hybrid rocket with a swirling gaseous oxygen flow type engine, we made a flight model engine. Burning tests of the engine showed that a maximum thrust of 692 N and a specific impulse of 263 s (at sea level) were achieved. We designed a small hybrid rocket with this engine. The rocket measured 1.8 m in length and 15.4 kg in mass. To confirm the flight stability of the rocket, wind tunnel tests using a 112-scale model of the rocket and simulations of the flight ...

Solid Rocket Testing at AFRL (Briefing Charts)

Science.gov (United States)

2016-10-21

Distribution Unlimited. PA#16492 2 Agenda • Solid Rocket Motors • History of Sea Level Testing • Small Component Testing • Full-scale Testing • Altitude...Facility • History of Testing • Questions -Distribution A: Approved for Public Release; Distribution Unlimited. PA#16492 3 RQ-West • AFRL/RQ...INTEGRATION FACILITY NATIONAL HOVER TEST FACILITY TITAN SRM TEST FACILITY TS-1C1-125 LARGE ENGINE/COMPONENT TEST FACILITY TS-1A 1-120 1-115 X-33 LAUNCH

Investigation of Electron Density Profile in the ionospheric D and E region by Kagoshima rocket experiment

Science.gov (United States)

Ashihara, Y.; Ishisaka, K.; Miyake, T.; Okada, T.; Nagano, I.; Abe, T.; Ono, T.

2007-12-01

The radio wave propagation characteristic in the lower ionosphere is important because of its effect on commercial radio communication, navigation, and broadcast services. The electron density is of primary interest in this region because the high ion-neutral collision frequencies result in radio wave absorption. In order to investigate the ionization structure in the ionospheric D and E region by using the propagation characteristics of MF-band and LF-band radio waves, S-310-37 and S-520-23 sounding rocket experiments have been carried out at Uchinoura Space Center (USC). S-310-37 sounding rocket was launched at 11:20 LT on January 16, 2007. The apex of rocket trajectory was about 138 km. Then S-520-23 sounding rocket was launched at 19:20 LT on September 2, 2007. The apex was about 279 km. As a common measurement, these sounding rockets measure the fields intensities and the waveform of radio waves from NHK Kumamoto broadcasting station (873kHz, 500kW) and JJY signals from Haganeyama LF radio station (60kHz, 50kW). The approximate electron density profile can be determined from the comparison between these experimental results and propagation characteristics calculated by the full wave method. We will get the most probable electron density profile in the ionosphere. In presentation, we will show the propagation characteristic of LF/MF radio waves measured by two sounding rocket experiments. Then we will discuss the analysis method and the estimated electron density profile in the ionosphere.

Ionospheric effects of rocket exhaust products (HEAO-C, Skylab and SPS-HLLV)

International Nuclear Information System (INIS)

Zinn, J.; Sutherland, D.; Stone, S.N.; Duncan, L.M.; Behnke, R.

1980-10-01

This paper reviews the current state of our understanding of the problem of ionospheric F-layer depletions produced by chemical effects of the exhaust gases from large rockets, with particular emphasis on the Heavy Lift Launch Vehicles (HLLV) proposed for use in the construction of solar power satellites. The currently planned HLLV flight profile calls for main second-stage propulsion confined to altitudes below 124 km, and a brief orbit-circularization maneuver at apogee. The second-stage engines deposit 9 x 10 31 H 2 O and H 2 molecules between 56 and 124 km. Model computations show that they diffuse gradually into the ionospheric F region, where they lead to weak but widespread and persistent depletions of ionization and continuous production of H atoms. The orbit-circularization burn deposits 9 x 10 29 exhaust molecules at about 480-km altitude. These react rapidly with the F2 region 0 + ions, leading to a substantial (factor-of-three) reduction in plasma density, which extends over a 1000- by 2000-km region and persists for four to five hours. Also described are experimental airglow and incoherent-scatter radar measurements performed in conjunction with the 1979 launch of satellite HEAO-C, together with prelaunch and post-launch computations of the ionospheric effects. Several improvements in the model have been driven by the experimental observations. The computer model is described in some detail

Pre-launch simulation experiment of microwave-ionosphere nonlinear interaction rocket experiment in the space plasma chamber

Energy Technology Data Exchange (ETDEWEB)

Kaya, N. (Kobe University, Kobe, Japan); Tsutsui, M. (Kyoto University, Uji, Japan); Matsumoto, H. (Kyoto University, Kyoto, Japan)

1980-09-01

A pre-flight test experiment of a microwave-ionosphere nonlinear interaction rocket experiment (MINIX) has been carried out in a space plasma simulation chamber. Though the first rocket experiment ended up in failure because of a high voltage trouble, interesting results are observed in the pre-flight experiment. A significant microwave heating of plasma up to 300% temperature increase is observed. Strong excitations of plasma waves by the transmitted microwaves in the VLF and HF range are observed as well. These microwave effects may have to be taken into account in solar power satellite projects in the future.

Bantam: A Systematic Approach to Reusable Launch Vehicle Technology Development

Science.gov (United States)

Griner, Carolyn; Lyles, Garry

1999-01-01

capability to expand the capability of a reusable first stage, including ground launch, powered return to the launch site, and a fully reusable rocket propulsion system. A Bantam technology goal is to demonstrate twenty-five flights with no unplanned rocket engine maintenance and only minor planned maintenance or inspections. The design goal of the propulsion system is a mission life of one hundred.

Additive Manufacturing for Affordable Rocket Engines

Science.gov (United States)

West, Brian; Robertson, Elizabeth; Osborne, Robin; Calvert, Marty

2016-01-01

Additive manufacturing (also known as 3D printing) technology has the potential to drastically reduce costs and lead times associated with the development of complex liquid rocket engine systems. NASA is using 3D printing to manufacture rocket engine components including augmented spark igniters, injectors, turbopumps, and valves. NASA is advancing the process to certify these components for flight. Success Story: MSFC has been developing rocket 3D-printing technology using the Selective Laser Melting (SLM) process. Over the last several years, NASA has built and tested several injectors and combustion chambers. Recently, MSFC has 3D printed an augmented spark igniter for potential use the RS-25 engines that will be used on the Space Launch System. The new design is expected to reduce the cost of the igniter by a factor of four. MSFC has also 3D printed and tested a liquid hydrogen turbopump for potential use on an Upper Stage Engine. Additive manufacturing of the turbopump resulted in a 45% part count reduction. To understanding how the 3D printed parts perform and to certify them for flight, MSFC built a breadboard liquid rocket engine using additive manufactured components including injectors, turbomachinery, and valves. The liquid rocket engine was tested seven times in 2016 using liquid oxygen and liquid hydrogen. In addition to exposing the hardware to harsh environments, engineers learned to design for the new manufacturing technique, taking advantage of its capabilities and gaining awareness of its limitations. Benefit: The 3D-printing technology promises reduced cost and schedule for rocket engines. Cost is a function of complexity, and the most complicated features provide the largest opportunities for cost reductions. This is especially true where brazes or welds can be eliminated. The drastic reduction in part count achievable with 3D printing creates a waterfall effect that reduces the number of processes and drawings, decreases the amount of touch

Ricardo Dyrgalla (1910-1970), pioneer of rocket development in Argentina

Science.gov (United States)

de León, Pablo

2009-12-01

One of the most important developers of liquid propellant rocket engines in Argentina was Polish-born Ricardo Dyrgalla. Dyrgalla immigrated to Argentina from the United Kingdom in 1946, where he had been studying German weapons development at the end of the Second World War. A trained pilot and aeronautical engineer, he understood the intricacies of rocket propulsion and was eager to find practical applications to his recently gained knowledge. Dyrgalla arrived in Argentina during Juan Perón's first presidency, a time when technicians from all over Europe were being recruited to work in various projects for the recently created Argentine Air Force. Shortly after immigrating, Dyrgalla proposed to develop an advanced air-launched weapon, the Tábano, based on a rocket engine of his design, the AN-1. After a successful development program, the Tábano was tested between 1949 and 1951; however, the project was canceled by the government shortly after. Today, the AN-1 rocket engine is recognized as the first liquid propellant rocket to be developed in South America. Besides the AN-1, Dyrgalla also developed several other rockets systems in Argentina, including the PROSON, a solid-propellant rocket launcher developed by the Argentine Institute of Science and Technology for the Armed Forces (CITEFA). In the late 1960s, Dyrgalla and his family relocated to Brazil due mostly to the lack of continuation of rocket development in Argentina. There, he worked for the Institute of Aerospace Technology (ITA) until his untimely death in 1970. Ricardo Dyrgalla deserves to be recognized among the world's rocket pioneers and his contribution to the science and engineering of rocketry deserves a special place in the history of South America's rocketry and space flight advocacy programs.

Transient Acoustic Environment Prediction Tool for Launch Vehicles in Motion during Early Lift-Off, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Launch vehicles experience extreme acoustic loads dominated by rocket exhaust plume interactions with ground structures during lift-off, which can produce damaging...

NASA's Space Launch System: An Enabling Capability for International Exploration

Science.gov (United States)

Creech, Stephen D.; May, Todd A.; Robinson, Kimberly F.

2014-01-01

As the program moves out of the formulation phase and into implementation, work is well underway on NASA's new Space Launch System, the world's most powerful launch vehicle, which will enable a new era of human exploration of deep space. As assembly and testing of the rocket is taking place at numerous sites around the United States, mission planners within NASA and at the agency's international partners continue to evaluate utilization opportunities for this ground-breaking capability. Developed with the goals of safety, affordability, and sustainability in mind, the SLS rocket will launch the Orion Multi-Purpose Crew Vehicle (MPCV), equipment, supplies, and major science missions for exploration and discovery. NASA is developing this new capability in an austere economic climate, a fact which has inspired the SLS team to find innovative solutions to the challenges of designing, developing, fielding, and operating the largest rocket in history, via a path that will deliver an initial 70 metric ton (t) capability in December 2017 and then continuing through an incremental evolutionary strategy to reach a full capability greater than 130 t. SLS will be enabling for the first missions of human exploration beyond low Earth in almost half a century, and from its first crewed flight will be able to carry humans farther into space than they have ever voyaged before. In planning for the future of exploration, the International Space Exploration Coordination Group, representing 12 of the world's space agencies, has created the Global Exploration Roadmap, which outlines paths toward a human landing on Mars, beginning with capability-demonstrating missions to the Moon or an asteroid. The Roadmap and corresponding NASA research outline the requirements for reference missions for these destinations. SLS will offer a robust way to transport international crews and the air, water, food, and equipment they would need for such missions.

European souding-rocket, balloon and related research, with emphasis on experiments at high latitudes. Proceedings

International Nuclear Information System (INIS)

Halvorsen, T.; Battrick, B.; Rowley, C.

1978-06-01

The document contains 75 papers presented at the fourth symposium held within the framework of the ESA programme advisory committee on the special project concerning the launching of sounding rockets (PAC), which took place at Ajaccio, Corsica, from 24-29 April 1978. The symposium, organised jointly by the PAC and the Centre national d'etudes spaciales (CNES), was divided into 8 sessions embracing, respectively, a review of the national programmes, magnetospheric physics, launch range presentations, middle-atmosphere physics, astrophysic, material sciences, subsatellites and technology, as relevant to souding-rocket and balloon research. Working groups formed during the symposium formulated proposals to the PAC for new or intensified studies based on coordination of other experiments with EISCAT, optical experiments in connection with range and other experiment, middle-atmosphere programme experiments, and subsatellite experiments

Response of Launch Pad Structures to Random Acoustic Excitation

Directory of Open Access Journals (Sweden)

Ravi N. Margasahayam

1994-01-01

Full Text Available The design of launch pad structures, particularly those having a large area-to-mass ratio, is governed by launch-induced acoustics, a relatively short transient with random pressure amplitudes having a non-Gaussian distribution. The factors influencing the acoustic excitation and resulting structural responses are numerous and cannot be predicted precisely. Two solutions (probabilistic and deterministic for the random vibration problem are presented in this article from the standpoint of their applicability to predict the response of ground structures exposed to rocket noise. Deficiencies of the probabilistic method, especially to predict response in the low-frequency range of launch transients (below 20 Hz, prompted the development of the deterministic analysis. The relationship between the two solutions is clarified for future implementation in a finite element method (FEM code.

Effects of rocket exhaust products in the thermosphere and ionsphere

International Nuclear Information System (INIS)

Zinn, J.; Sutherland, C.D.

1980-02-01

This paper reviews the current state of understanding of the problem of ionospheric F-layer depletions produced by chemical effects of the exhaust gases from large rockets, with particular emphasis on the Heavy Lift Launch Vehicles (HLLV) proposed for use in the construction of solar power satellites. The currently planned HLLV flight profile calls for main second-stage propulsion confined to altitudes below 124 km, and a brief orbit circularization maneuver at apogee. The second stage engines deposit 9 x 10 31 H 2 O and H 2 molecules between 74 and 124 km. Model computations show that they diffuse gradually into the ionospheric F region, where they lead to weak but widespread and persistent depletions of ionization and continuous production of H atoms. The orbit circularization burn deposits 9 x 10 29 exhaust molecules at about 480-km altitude. These react rapidly with the F2 region 0 + ions, leading to a substantial (factor-of-three) reduction in plasma density, which extends over a 1000- by 2000-km region and persists for four to five hours. For purposes of computer model verification, a computation is included representing the Skylab I launch, for which observational data exist. The computations and data are compared, and the computer model is described

A new sounding rocket payload for solar plasma studies

International Nuclear Information System (INIS)

Bruner, M.E.; Brown, W.A.; Appert, K.L.

1989-01-01

A sounding rocket payload developed for studies of high-temperature plasmas associated with solar active regions and flares is described. The payload instruments will record both spectra and images in the UV, EUV, and soft X-ray regions of the spectrum. The instruments, including the Dual Range Spectrograph, the Flat Field Soft X-ray Spectrograph, the Normal Incidence Soft X-ray Imager, the UV Filtergraph, and the H-alpha Imaging system, are described. Attention is also given to the new structural system of the payload, based on a large optical table suspended within the payload cavity, which will support the optical elements in their correct positions and orientations and will maintain these alignments throughout the rocket launch environment. 8 refs

Hybrid adaptive ascent flight control for a flexible launch vehicle

Science.gov (United States)

Lefevre, Brian D.

For the purpose of maintaining dynamic stability and improving guidance command tracking performance under off-nominal flight conditions, a hybrid adaptive control scheme is selected and modified for use as a launch vehicle flight controller. This architecture merges a model reference adaptive approach, which utilizes both direct and indirect adaptive elements, with a classical dynamic inversion controller. This structure is chosen for a number of reasons: the properties of the reference model can be easily adjusted to tune the desired handling qualities of the spacecraft, the indirect adaptive element (which consists of an online parameter identification algorithm) continually refines the estimates of the evolving characteristic parameters utilized in the dynamic inversion, and the direct adaptive element (which consists of a neural network) augments the linear feedback signal to compensate for any nonlinearities in the vehicle dynamics. The combination of these elements enables the control system to retain the nonlinear capabilities of an adaptive network while relying heavily on the linear portion of the feedback signal to dictate the dynamic response under most operating conditions. To begin the analysis, the ascent dynamics of a launch vehicle with a single 1st stage rocket motor (typical of the Ares 1 spacecraft) are characterized. The dynamics are then linearized with assumptions that are appropriate for a launch vehicle, so that the resulting equations may be inverted by the flight controller in order to compute the control signals necessary to generate the desired response from the vehicle. Next, the development of the hybrid adaptive launch vehicle ascent flight control architecture is discussed in detail. Alterations of the generic hybrid adaptive control architecture include the incorporation of a command conversion operation which transforms guidance input from quaternion form (as provided by NASA) to the body-fixed angular rate commands needed by the

Structure of plasma blobs injected into the ionosphere from a rocket

International Nuclear Information System (INIS)

Aleksandrov, V.A.; Loevskii, A.S.; Popov, G.A.; Romanovskii, Iu.A.; Sobol, A.G.

1981-01-01

Plasma structure and dynamics have been studied by injecting plasma blobs (PB) into the ionosphere using two MR-12 rockets launched from Volgograd in 1978-1979. The mother-daughter system was used, and the daughter payload, which contained the plasma gun, was stabilized by rotation and separated at an altitude of 110 km at a speed of 2-3 m/s. The pulsed plasma was ejected to an altitude of from 110 to 145 km along and transverse to the rocket, which corresponded to pitch-angle ranges of 0-40 degrees and 70-110 degrees. The diagnostic equipment included an ion probe, a photometer, and electric and magnetic field detectors. A model of the PB processes was constructed, and it was confirmed by the experimental results: e.g., (1) the longitudinal size of the PB's reached several hundreds of meters, and the PB plasma density was found to be about 10 to the 7th to 10 to the 8th per cu cm and (2) the magnetic field had completely diffused into the PB

Level II Documentation of Launch Complex 31/32, Cape Canaveral Air Force Station, Florida

Science.gov (United States)

2008-12-01

Navigational Satellite Time and Ranging (NAVSTAR) Global Positioning System (GPS) program. Most of the above satellites have been launched from Cape...rocket, put together from existing hardware: an Honest John first stage, topped with two Nike stages, then a Recruit fourth stage, and finally, a T-55...a ground power oscillograph, exhaust fans, a sump pump, the dehydrator and humidity recorder, launch/perimeter and TV cameras, an Ampex- brand

A two-channel wave analyser for sounding rockets and satellites

International Nuclear Information System (INIS)

Brondz, E.

1989-04-01

Studies of low frequency electromagnetic waves, produced originally by lightning discharges penetrating the ionosphere, provide an important source of valuable information about the earth's surrounding plasma. Use of rockets and satellites supported by ground-based observations implies, unique opportunity for measuring in situ a number of parameters simultaneously in order to correlate data from various measurements. However, every rocket experiment has to be designed bearing in mind telemetry limitations and/or short flight duration. Typical flight duration for Norwegian rockets launched from Andoeya Rocket Range is 500 to 600 s. Therefore, the most desired way to use a rocket or satellite is to carry out data analyses on board in real time. Recent achievements in Digital Signal Processing (DSP) technology have made it possible to undertake very complex on board data manipulation. As a part of rocket instrumentation, a DSP based unit able to carry out on board analyses of low frequency electromagnetic waves in the ionosphere has been designed. The unit can be seen as a general purpose computer built on the basis of a fixed-point 16 bit signal processor. The unit is supplied with a program code in order to perform wave analyses on two independent channels simultaneously. The analyser is able to perform 256 point complex fast fourier transformations, and it produce a spectral power desity estimate on both channels every 85 ms. The design and construction of the DSP based unit is described and results from the tests are presented

A Business Analysis of a SKYLON-based European Launch Service Operator

Science.gov (United States)

Hempsell, Mark; Aprea, Julio; Gallagher, Ben; Sadlier, Greg

2016-04-01

Between 2012 and 2014 an industrial consortium led by Reaction Engines conducted a feasibility study for the European Space Agency with the objective to explore the feasibility of SKYLON as the basis for a launcher that meets the requirements established for the Next Generation European Launcher. SKYLON is a fully reusable single stage to orbit launch system that is enabled by the unique performance characteristic of the Synergetic Air-Breathing Rocket Engine and is under active development. The purpose of the study which was called ;SKYLON-based European Launch Service Operator (S-ELSO); was to support ESA decision making on launch service strategy by exploring the potential implications of this new launch system on future European launch capability and the European industry that supports it. The study explored both a SKYLON operator (S-ELSO) and SKYLON manufacturer as separate business ventures. In keeping with previous studies, the only strategy that was found that kept the purchase price of the SKYLON low enough for a viable operator business was to follow an ;airline; business model where the manufacturer sells SKYLONs to other operators in addition to S-ELSO. With the assumptions made in the study it was found that the SKYLON manufacturer with a total production run of between 30 and 100 SKYLONs could expect an Internal Rate of Return of around 10%. This was judged too low for all the funding to come from commercial funding sources, but is sufficiently high for a Public Private Partnership. The S-ELSO business model showed that the Internal Rate of Return would be high enough to consider operating without public support (i.e. commercial in operation, irrespective of any public funding of development), even when the average launch price is lowered to match the lowest currently quoted price for expendable systems.

Small Payload Launch Integrated Testing Services (SPLITS) - SPSDL

Science.gov (United States)

Plotner, Benjamin

2013-01-01

My experience working on the Small Payload Launch Integrated Testing Services project has been both educational and rewarding. I have been given the opportunity to work on and experiment with a number of exciting projects and initiatives, each offering different challenges and opportunities for teamwork and collaboration. One of my assignments is to aid in the design and construction of a small-scale two stage rocket as part of a Rocket University initiative. My duties include programming a microcontroller to control the various sensors on the rocket as well as process and transmit data. Additionally, I am writing a graphical user interface application for the ground station that will receive the transmitted data from the rocket and display the information on screen along with a 3D rendering displaying the rocket orientation. Another project I am working on is to design and develop the avionics that will be used to control a high altitude balloon flight that will test a sensor called a Micro Dosimeter that will measure the total ionizing dose absorbed by electrical components during a flight. This includes assembling and soldering the various sensors and components, programming a microcontroller to input and process data from the Micro Dosimeter, and transmitting the data down to a ground station as well as save the data to an on-board SD card. Additionally, I am aiding in the setup and development of ITOS (Integrated Test and Operations System) capability in the SPSDL (Spaceport Processing System Development Lab).

Discovery is in the VAB as STS-95 launch preparations continue

Science.gov (United States)

1998-01-01

In the Vehicle Assembly Building, the orbiter Discovery is mated with the external tank and solid rocket booster stack (seen behind the orbiter, to the left). The orbiter was recently painted with the NASA logo, termed the 'meatball,' on the left, or port, wing and both sides of the aft fuselage. Discovery (OV- 103) is the first of the orbiters to be launched with the new artwork. It is scheduled for its 25th flight, from Launch Pad 39B, on Oct. 29, 1998, for the STS-95 mission.

Thermohydraulic modeling of nuclear thermal rockets: The KLAXON code

International Nuclear Information System (INIS)

Hall, M.L.; Rider, W.J.; Cappiello, M.W.

1992-01-01

The hydrogen flow from the storage tanks, through the reactor core, and out the nozzle of a Nuclear Thermal Rocket is an integral design consideration. To provide an analysis and design tool for this phenomenon, the KLAXON code is being developed. A shock-capturing numerical methodology is used to model the gas flow (the Harten, Lax, and van Leer method, as implemented by Einfeldt). Preliminary results of modeling the flow through the reactor core and nozzle are given in this paper

Effect of Six Missile-Bay Baffle Configurations and a Rocket End Plate on Ejection Releases of an MB-1 Rocket from a 0.05 Scale Model of the Convair F-106A Airplane

Science.gov (United States)

Hinson, William F.; Lee, John B.

1959-01-01

As a continuation of an investigation of the release characteristics of an MB-1 rocket carried internally by the Convair F-106A airplane, six missile-bay baffle configurations and a rocket end plate have been investigated in the 27- by 27-inch preflight jet of the NASA Wallops Station. The MB-1 rocket used had retractable fins and was ejected from a missile bay modified by the addition of six different baffle configurations. For some tests a rocket end plate was added to the model. Dynamically scaled models (0.04956 scale) were tested at a simulated altitude of 22,450 feet and Mach numbers of 0.86, 1.59, and 1.98, and at a simulated altitude of 29,450 feet and a Mach number of 1.98. The results of this investigation indicate that the missile-bay baffle configurations and the rocket end plate may be used to reduce the positive pitch amplitude of the MB-1 rocket after release. The initial negative pitching velocity applied to the MB-1 rocket might then be reduced in order to maintain a near-level-flight attitude after release. As the fuselage angle of attack is increased, the negative pitch amplitude of the rocket is decreased.

Aerodynamics and flow characterisation of multistage rockets

Science.gov (United States)

Srinivas, G.; Prakash, M. V. S.

2017-05-01

The main objective of this paper is to conduct a systematic flow analysis on single, double and multistage rockets using ANSYS software. Today non-air breathing propulsion is increasing dramatically for the enhancement of space exploration. The rocket propulsion is playing vital role in carrying the payload to the destination. Day to day rocket aerodynamic performance and flow characterization analysis has becoming challenging task to the researchers. Taking this task as motivation a systematic literature is conducted to achieve better aerodynamic and flow characterization on various rocket models. The analyses on rocket models are very little especially in numerical side and experimental area. Each rocket stage analysis conducted for different Mach numbers and having different flow varying angle of attacks for finding the critical efficiency performance parameters like pressure, density and velocity. After successful completion of the analysis the research reveals that flow around the rocket body for Mach number 4 and 5 best suitable for designed payload. Another major objective of this paper is to bring best aerodynamics flow characterizations in both aero and mechanical features. This paper also brings feature prospectus of rocket stage technology in the field of aerodynamic design.

Solid rocket motor cost model

Science.gov (United States)

Harney, A. G.; Raphael, L.; Warren, S.; Yakura, J. K.

1972-01-01

A systematic and standardized procedure for estimating life cycle costs of solid rocket motor booster configurations. The model consists of clearly defined cost categories and appropriate cost equations in which cost is related to program and hardware parameters. Cost estimating relationships are generally based on analogous experience. In this model the experience drawn on is from estimates prepared by the study contractors. Contractors' estimates are derived by means of engineering estimates for some predetermined level of detail of the SRM hardware and program functions of the system life cycle. This method is frequently referred to as bottom-up. A parametric cost analysis is a useful technique when rapid estimates are required. This is particularly true during the planning stages of a system when hardware designs and program definition are conceptual and constantly changing as the selection process, which includes cost comparisons or trade-offs, is performed. The use of cost estimating relationships also facilitates the performance of cost sensitivity studies in which relative and comparable cost comparisons are significant.

Business Intelligence Modeling in Launch Operations

Science.gov (United States)

Bardina, Jorge E.; Thirumalainambi, Rajkumar; Davis, Rodney D.

2005-01-01

This technology project is to advance an integrated Planning and Management Simulation Model for evaluation of risks, costs, and reliability of launch systems from Earth to Orbit for Space Exploration. The approach builds on research done in the NASA ARC/KSC developed Virtual Test Bed (VTB) to integrate architectural, operations process, and mission simulations for the purpose of evaluating enterprise level strategies to reduce cost, improve systems operability, and reduce mission risks. The objectives are to understand the interdependency of architecture and process on recurring launch cost of operations, provide management a tool for assessing systems safety and dependability versus cost, and leverage lessons learned and empirical models from Shuttle and International Space Station to validate models applied to Exploration. The systems-of-systems concept is built to balance the conflicting objectives of safety, reliability, and process strategy in order to achieve long term sustainability. A planning and analysis test bed is needed for evaluation of enterprise level options and strategies for transit and launch systems as well as surface and orbital systems. This environment can also support agency simulation .based acquisition process objectives. The technology development approach is based on the collaborative effort set forth in the VTB's integrating operations. process models, systems and environment models, and cost models as a comprehensive disciplined enterprise analysis environment. Significant emphasis is being placed on adapting root cause from existing Shuttle operations to exploration. Technical challenges include cost model validation, integration of parametric models with discrete event process and systems simulations. and large-scale simulation integration. The enterprise architecture is required for coherent integration of systems models. It will also require a plan for evolution over the life of the program. The proposed technology will produce

Business intelligence modeling in launch operations

Science.gov (United States)

Bardina, Jorge E.; Thirumalainambi, Rajkumar; Davis, Rodney D.

2005-05-01

The future of business intelligence in space exploration will focus on the intelligent system-of-systems real-time enterprise. In present business intelligence, a number of technologies that are most relevant to space exploration are experiencing the greatest change. Emerging patterns of set of processes rather than organizational units leading to end-to-end automation is becoming a major objective of enterprise information technology. The cost element is a leading factor of future exploration systems. This technology project is to advance an integrated Planning and Management Simulation Model for evaluation of risks, costs, and reliability of launch systems from Earth to Orbit for Space Exploration. The approach builds on research done in the NASA ARC/KSC developed Virtual Test Bed (VTB) to integrate architectural, operations process, and mission simulations for the purpose of evaluating enterprise level strategies to reduce cost, improve systems operability, and reduce mission risks. The objectives are to understand the interdependency of architecture and process on recurring launch cost of operations, provide management a tool for assessing systems safety and dependability versus cost, and leverage lessons learned and empirical models from Shuttle and International Space Station to validate models applied to Exploration. The systems-of-systems concept is built to balance the conflicting objectives of safety, reliability, and process strategy in order to achieve long term sustainability. A planning and analysis test bed is needed for evaluation of enterprise level options and strategies for transit and launch systems as well as surface and orbital systems. This environment can also support agency simulation based acquisition process objectives. The technology development approach is based on the collaborative effort set forth in the VTB's integrating operations, process models, systems and environment models, and cost models as a comprehensive disciplined

E region neutral winds in the postmidnight diffuse aurora during the atmospheric response in aurora 1 rocket campaign

International Nuclear Information System (INIS)

Brinkman, D.G.; Walterscheid, R.L.; Lyons, L.R.

1995-01-01

Measured E region neutral winds from the Atmospheric Response in Aurora (ARIA 1) rocket campaign are compared with winds predicted by a high-resolution nonhydrostatic dynamical thermosphere model. The ARIA 1 rockets were launched into the postmidnight diffuse aurora during the recovery phase of a substorm. Simulations have shown that electrodynamical coupling between the auroral ionosphere and the thermosphere was expected to be strong during active diffuse auroral conditions. This is the first time that simulations using the time history of detailed specifications of the magnitude and latitudinal variation of the auroral forcing based on measurements have been compared to simultaneous wind measurements. Model inputs included electron densities derived from ground-based airglow measurements, precipitating electron fluxes measured by the rocket, electron densities measured on the rocket, electric fields derived from magnetometer and satellite ion drift measurements, and large-scale background winds from a thermospheric general circulation model. Our model predicted a strong jet of eastward winds at E region heights. A comparison between model predicted and observed winds showed modest agreement. Above 135 km the model predicted zonal winds with the correct sense, the correct profile shape, and the correct altitude of the peak wind. However, it overpredicted the magnitude of the eastward winds by more than a factor or 2. For the meridional winds the model predicted the general sense of the winds but was unable to predict the structure or strength of the winds seen in the observations. Uncertainties in the magnitude and latitudinal structure of the electric field and in the magnitude of the background winds are the most likely sources of error contributing to the differences between model and observed winds. Between 110 and 135 km the agreement between the model and observations was poor because of a large unmodeled jetlike feature in the observed winds

The Rocket Electric Field Sounding (REFS) Program: Prototype Design and Successful First Launch

Science.gov (United States)

1992-01-15

insulators surrounding the stators, and stator edges themselves, are fully covered by the rotor , so that any effects of charge on the insulators are...Jumper performed a separate analysis of the aerodynamics (primarily the " Magnus effect ") induced by the relative rotation of rocket body and shell. The...significant advantages over an aircraft in simplicity and calibration. A single cylindrical rotor covering most of the payload acts as the shutter for all

The Development of Rocketry Capability in New Zealand—World Record Rocket and First of Its Kind Rocketry Course

Directory of Open Access Journals (Sweden)

George Buchanan

2015-02-01

Full Text Available The University of Canterbury has developed a rocket research group, UC Rocketry, which recently broke the world altitude record for an I-class motor (impulse of 320–640 Ns and has run a rocketry course for the first time in New Zealand. This paper discusses the development and results of the world record rocket “Milly” and details all the fundamental elements of the rocketry final year engineering course, including the manufacturing processes, wind tunnel testing, avionics, control and the final rocket launch of “Smokey”. The rockets Milly and Smokey are an example of the design, implementation and testing methodologies that have significantly contributed to research and graduates for New Zealand’s space program.

Ionospheric E–F valley observed by a sounding rocket at the low-latitude station Hainan

Directory of Open Access Journals (Sweden)

J. K. Shi

2013-08-01

Full Text Available According to the sounding rocket experiment conducted at Hainan ionospheric observatory (19.5° N, 109.1° E, a valley between the E layer and F layer in the ionospheric electron density profile is observed and presented. The sounding rocket was launched in the morning (06:15 LT on 7 May 2011, and the observed electron density profile outside the valley agrees with the simultaneous observation by the DPS-4 digisonde at the same station. The width of the observed valley was about 42 km, the depth almost 50%, and the altitude of the electron density minimum 123.5 km. This is the first observation of the E–F valley in the low-latitude region in the East Asian sector. The results are also compared with models, and the physical mechanism of the observed valley is discussed in this paper.

Gigantic Circular Shock Acoustic Waves in the Ionosphere Triggered by the Launch of FORMOSAT-5 Satellite

Science.gov (United States)

Chou, Min-Yang; Shen, Ming-Hsueh; Lin, Charles C. H.; Yue, Jia; Chen, Chia-Hung; Liu, Jann-Yenq; Lin, Jia-Ting

2018-02-01

The launch of SpaceX Falcon 9 rocket delivered Taiwan's FORMOSAT-5 satellite to orbit from Vandenberg Air Force Base in California at 18:51:00 UT on 24 August 2017. To facilitate the delivery of FORMOSAT-5 to its mission orbit altitude of 720 km, the Falcon 9 made a steep initial ascent. During the launch, the supersonic rocket induced gigantic circular shock acoustic waves (SAWs) in total electron content (TEC) over the western United States beginning approximately 5 min after the liftoff. The circular SAWs emanated outward with 20 min duration, horizontal phase velocities of 629-726 m/s, horizontal wavelengths of 390-450 km, and period of 10.28 ± 1 min. This is the largest rocket-induced circular SAWs on record, extending approximately 114-128°W in longitude and 26-39°N in latitude ( 1,500 km in diameter), and was due to the unique, nearly vertical attitude of the rocket during orbit insertion. The rocket-exhaust plume subsequently created a large-scale ionospheric plasma hole ( 900 km in diameter) with 10-70% TEC depletions in comparison with the reference days. While the circular SAWs, with a relatively small amplitude of TEC fluctuations, likely did not introduce range errors into the Global Navigation Satellite Systems navigation and positioning system, the subsequent ionospheric plasma hole, on the other hand, could have caused spatial gradients in the ionospheric plasma potentially leading to a range error of 1 m.

The Cost-Optimal Size of Future Reusable Launch Vehicles

Science.gov (United States)

Koelle, D. E.

2000-07-01

The paper answers the question, what is the optimum vehicle size — in terms of LEO payload capability — for a future reusable launch vehicle ? It is shown that there exists an optimum vehicle size that results in minimum specific transportation cost. The optimum vehicle size depends on the total annual cargo mass (LEO equivalent) enviseaged, which defines at the same time the optimum number of launches per year (LpA). Based on the TRANSCOST-Model algorithms a wide range of vehicle sizes — from 20 to 100 Mg payload in LEO, as well as launch rates — from 2 to 100 per year — have been investigated. It is shown in a design chart how much the vehicle size as well as the launch rate are influencing the specific transportation cost (in MYr/Mg and USS/kg). The comparison with actual ELVs (Expendable Launch Vehicles) and Semi-Reusable Vehicles (a combination of a reusable first stage with an expendable second stage) shows that there exists only one economic solution for an essential reduction of space transportation cost: the Fully Reusable Vehicle Concept, with rocket propulsion and vertical take-off. The Single-stage Configuration (SSTO) has the best economic potential; its feasibility is not only a matter of technology level but also of the vehicle size as such. Increasing the vehicle size (launch mass) reduces the technology requirements because the law of scale provides a better mass fraction and payload fraction — practically at no cost. The optimum vehicle design (after specification of the payload capability) requires a trade-off between lightweight (and more expensive) technology vs. more conventional (and cheaper) technology. It is shown that the the use of more conventional technology and accepting a somewhat larger vehicle is the more cost-effective and less risky approach.

Low-Cost Propellant Launch to LEO from a Tethered Balloon - 'Propulsion Depots' Not 'Propellant Depots'

Science.gov (United States)

Wilcox, Brian H.; Schneider, Evan G.; Vaughan, David A.; Hall, Jeffrey L.; Yu, Chi Yau

2011-01-01

As we have previously reported, it may be possible to launch payloads into low-Earth orbit (LEO) at a per-kilogram cost that is one to two orders of magnitude lower than current launch systems, using only a relatively small capital investment (comparable to a single large present-day launch). An attractive payload would be large quantities of high-performance chemical rocket propellant (e.g. Liquid Oxygen/Liquid Hydrogen (LO2/LH2)) that would greatly facilitate, if not enable, extensive exploration of the moon, Mars, and beyond.

Measurements of auroral particles by means of sounding rockets of mother-daughter type

International Nuclear Information System (INIS)

Falck, A.

1985-11-01

The scientific objective of the S17 payloads was to study the ionosphere during auroral situations and especially with regards to the local fine structure and a possible separation of spatial and temporal variations of auroral phenomena. The intensities of 8 keV and 2 keV electrons have been measured from one sounding rocket launched into a breakup aurora of moderate activity and from another rocket launched into a very active substorm situation. Both the rockets were of mother-daughter type i.e. had two separated payloads. The general features in the data of different particle energies were very similar over the whole flight time of the rockets. Special events and gradients and well identifiable shapes in the particle intensities were studied to see if the intensity fluctuations obtained from two detectors in one payload or from detectors into separate payloads were time delayed. Such time delays in the particle flux intensities were obvious in both of the rocket measurements and most of these time shifts could be understood as caused by spatial variations in the particle precipitation. In parts of the rocket flights the particle intensity variations were true temporal changes. The time lags between 8 keV and 2 keV electron intensities detected in the same payload, which could be observed and were obtained by crosscorrelation analyses, were in the range less than 0.3 s and most of them less than 0.1 s. If the time differences are assumed to be caused by the velocity dispersion of the particles, the particle data reported here placed the modulation source at a distance of less than 10 000 km from the rocket position. Measurements at the S17-1 mother payload of the electric field have been compared with data of precipitating electrons and low-light-level-TV-recording of the auroral situation. An inverted-V precipitation event was observed and was associated with auroral arcs and with reversals of the measured electric field components implicating the possibility of

Transient Mathematical Modeling for Liquid Rocket Engine Systems: Methods, Capabilities, and Experience

Science.gov (United States)

Seymour, David C.; Martin, Michael A.; Nguyen, Huy H.; Greene, William D.

2005-01-01

The subject of mathematical modeling of the transient operation of liquid rocket engines is presented in overview form from the perspective of engineers working at the NASA Marshall Space Flight Center. The necessity of creating and utilizing accurate mathematical models as part of liquid rocket engine development process has become well established and is likely to increase in importance in the future. The issues of design considerations for transient operation, development testing, and failure scenario simulation are discussed. An overview of the derivation of the basic governing equations is presented along with a discussion of computational and numerical issues associated with the implementation of these equations in computer codes. Also, work in the field of generating usable fluid property tables is presented along with an overview of efforts to be undertaken in the future to improve the tools use for the mathematical modeling process.

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

This Is Rocket Science!

Science.gov (United States)

Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

2013-09-01

Using model rockets to teach physics can be an effective way to engage students in learning. In this paper, we present a curriculum developed in response to an expressed need for helping high school students review physics equations in preparation for a state-mandated exam. This required a mode of teaching that was more advanced and analytical than that offered by Estes Industries, but more basic than the analysis of Nelson et al. In particular, drag is neglected until the very end of the exercise, which allows the concept of conservation of energy to be shown when predicting the rocket's flight. Also, the variable mass of the rocket motor is assumed to decrease linearly during the flight (while the propulsion charge and recovery delay charge are burning) and handled simplistically by using an average mass value. These changes greatly simplify the equations needed to predict the times and heights at various stages of flight, making it more useful as a review of basic physics. Details about model rocket motors, range safety, and other supplemental information may be found online at Apogee Components4 and the National Association of Rocketry.5

Study of solid rocket motors for a space shuttle booster. Volume 1: Executive summary

Science.gov (United States)

1972-01-01

The design, development, production, and launch support analysis for determining the solid propellant rocket engine to be used with the space shuttle are discussed. Specific program objectives considered were: (1) definition of engine designs to satisfy the performance and configuration requirements of the various vehicle/booster concepts, (2) definition of requirements to produce booster stages at rates of 60, 40, 20, and 10 launches per year in a man-rated system, and (3) estimation of costs for the defined SRM booster stages.

NTR-Enhanced Lunar-Base Supply using Existing Launch Fleet Capabilities

Energy Technology Data Exchange (ETDEWEB)

John D. Bess; Emily Colvin; Paul G. Cummings

2009-06-01

During the summer of 2006, students at the Center for Space Nuclear Research sought to augment the current NASA lunar exploration architecture with a nuclear thermal rocket (NTR). An additional study investigated the possible use of an NTR with existing launch vehicles to provide 21 metric tons of supplies to the lunar surface in support of a lunar outpost. Current cost estimates show that the complete mission cost for an NTR-enhanced assembly of Delta-IV and Atlas V vehicles may cost 47-86% more than the estimated Ares V launch cost of $1.5B; however, development costs for the current NASA architecture have not been assessed. The additional cost of coordinating the rendezvous of four to six launch vehicles with an in-orbit assembly facility also needs more thorough analysis and review. Future trends in launch vehicle use will also significantly impact the results from this comparison. The utility of multiple launch vehicles allows for the development of a more robust and lower risk exploration architecture.

NTR-Enhanced Lunar-Base Supply using Existing Launch Fleet Capabilities

International Nuclear Information System (INIS)

Bess, John D.; Colvin, Emily; Cummings, Paul G.

2009-01-01

During the summer of 2006, students at the Center for Space Nuclear Research sought to augment the current NASA lunar exploration architecture with a nuclear thermal rocket (NTR). An additional study investigated the possible use of an NTR with existing launch vehicles to provide 21 metric tons of supplies to the lunar surface in support of a lunar outpost. Current cost estimates show that the complete mission cost for an NTR-enhanced assembly of Delta-IV and Atlas V vehicles may cost 47-86% more than the estimated Ares V launch cost of $1.5B; however, development costs for the current NASA architecture have not been assessed. The additional cost of coordinating the rendezvous of four to six launch vehicles with an in-orbit assembly facility also needs more thorough analysis and review. Future trends in launch vehicle use will also significantly impact the results from this comparison. The utility of multiple launch vehicles allows for the development of a more robust and lower risk exploration architecture

Thermohydraulic Design Analysis Modeling for Korea Advanced NUclear Thermal Engine Rocket for Space Application

Energy Technology Data Exchange (ETDEWEB)

Nam, Seung Hyun; Choi, Jae Young; Venneria, Paolo F.; Jeong, Yong Hoon; Chang, Soon Heung [KAIST, Daejeon (Korea, Republic of)

2015-05-15

Space exploration is a realistic and profitable goal for long-term humanity survival, although the harsh space environment imposes lots of severe challenges to space pioneers. To date, almost all space programs have relied upon Chemical Rockets (CRs) rating superior thrust level to transit from the Earth's surface to its orbit. However, CRs inherently have insurmountable barrier to carry out deep space missions beyond Earth's orbit due to its low propellant efficiency, and ensuing enormous propellant requirement and launch costs. Meanwhile, nuclear rockets typically offer at least two times the propellant efficiency of a CR and thus notably reduce the propellant demand. Particularly, a Nuclear Thermal Rocket (NTR) is a leading candidate for near-term manned missions to Mars and beyond because it satisfies a relatively high thrust as well as a high efficiency. The superior efficiency of NTRs is due to both high energy density of nuclear fuel and the low molecular weight propellant of Hydrogen (H{sub 2}) over the chemical reaction by-products. A NTR uses thermal energy released from a nuclear fission reactor to heat the H{sub 2} propellant and then exhausted the highly heated propellant through a propelling nozzle to produce thrust. A propellant efficiency parameter of rocket engines is specific impulse (I{sub s}p) which represents the ratio of the thrust over the propellant consumption rate. If the average exhaust H{sub 2} temperature of a NTR is around 3,000 K, the I{sub s}p can be achieved as high as 1,000 s as compared with only 450 - 500 s of the best CRs. For this reason, NTRs are favored for various space applications such as orbital tugs, lunar transports, and manned missions to Mars and beyond. The best known NTR development effort was conducted from 1955 to1974 under the ROVER and NERVA programs in the USA. These programs had successfully designed and tested many different reactors and engines. After these projects, the researches on NERVA derived

The Off-plane Grating Rocket Experiment

Science.gov (United States)

Donovan, Benjamin

2018-01-01

The next generation of X-ray spectrometers necessitate significant increases in both resolution and effective area to achieve the science goals set forth in the 2010 Decadal Survey and the 2013 Astrophysics Roadmap. The Off-plane Grating Rocket Experiment (OGRE), an X-ray spectroscopy suborbital rocket payload currently scheduled for launch in Q3 2020, will serve as a testbed for several key technologies which can help achieve the desired performance increases of future spectrometers. OGRE will be the first instrument to fly mono-crystalline silicon X-ray mirrors developed at NASA Goddard Space Flight Center. The payload will also utilize an array of off-plane gratings manufactured at The Pennsylvania State University. Additionally, the focal plane will be populated with an array of four electron-multiplying CCDs developed by the Open University and XCAM Ltd. With these key technologies, OGRE hopes to achieve the highest resolution on-sky soft X-ray spectrum to date. We discuss the optical design, expected performance, and the current status of the payload.

Registration of ELF waves in rocket-satellite experiment with plasma injection

Science.gov (United States)

Korobeinikov, V. G.; Oraevskii, V. N.; Ruzhin, Iu. Ia.; Sobolev, Ia. P.; Skomarovskii, V. S.; Chmyrev, V. M.; Namazov, C. A.; Pokhunkov, A. A.; Nesmeianov, V. I.

1992-12-01

Two rocket KOMBI-SAMA experiments with plasma injection at height 100-240 km were performed in August 1987 in the region of Brazilian magnetic anomaly (L = 1.25). The launching time of the rocket was determined so that plasma injection was at the time when COSMOS 1809 satellite passed as close as possible to magnetic tube of injection. Caesium plasma jet was produced during not less than 300 s by an electric plasma generator separated from the payload. When the satellite passed the geomagnetic tube intersecting the injection region an enhancement of ELF emission at 140 Hz, 450 Hz by a factor of 2 was registered on board the satellite. An enhancement of energetic particle flux by a factor of 4-5 was registered on board the rocket. Observed ELF emission below 100 Hz is interpreted as the generation of oblique electromagnetic ion-cyclotron waves due to drift plasma instability at the front of the plasma jet.

NASA's Space Launch System: A New Capability for Science and Exploration

Science.gov (United States)

Crumbly, Christopher M.; May, Todd A.; Robinson, Kimberly F.

2014-01-01

The National Aeronautics and Space Administration's (NASA's) Marshall Space Flight Center (MSFC) is directing efforts to build the Space Launch System (SLS), a heavy-lift rocket that will launch the Orion Multi-Purpose Crew Vehicle (MPCV) and other high-priority payloads into deep space. Its evolvable architecture will allow NASA to begin with human missions beyond the Moon and then go on to transport astronauts or robots to distant places such as asteroids and Mars. Developed with the goals of safety, affordability, and sustainability in mind, SLS will start with 10 percent more thrust than the Saturn V rocket that launched astronauts to the Moon 40 years ago. From there it will evolve into the most powerful launch vehicle ever flown, via an upgrade approach that will provide building blocks for future space exploration. This paper will explain how NASA will execute this development within flat budgetary guidelines by using existing engines assets and heritage technology, from the initial 70 metric ton (t) lift capability through a block upgrade approach to an evolved 130-t capability, and will detail the progress that has already been made toward a first launch in 2017. This paper will also explore the requirements needed for human missions to deep-space destinations and for game-changing robotic science missions, and the capability of SLS to meet those requirements and enable those missions, along with the evolution strategy that will increase that capability. The International Space Exploration Coordination Group, representing 12 of the world's space agencies, has worked together to create the Global Exploration Roadmap, which outlines paths towards a human landing on Mars, beginning with capability-demonstrating missions to the Moon or an asteroid. The Roadmap and corresponding NASA research outline the requirements for reference missions for all three destinations. The SLS will offer a robust way to transport international crews and the air, water, food, and

Tabletop Experimental Track for Magnetic Launch Assist

Science.gov (United States)

2000-01-01

Marshall Space Flight Center's (MSFC's) Advanced Space Transportation Program has developed the Magnetic Launch Assist System, formerly known as the Magnetic Levitation (MagLev) technology that could give a space vehicle a running start to break free from Earth's gravity. A Magnetic Launch Assist system would use magnetic fields to levitate and accelerate a vehicle along a track at speeds up to 600 mph. The vehicle would shift to rocket engines for launch into orbit. Similar to high-speed trains and roller coasters that use high-strength magnets to lift and propel a vehicle a couple of inches above a guideway, a Magnetic Launch Assist system would electromagnetically propel a space vehicle along the track. The tabletop experimental track for the system shown in this photograph is 44-feet long, with 22-feet of powered acceleration and 22-feet of passive braking. A 10-pound carrier with permanent magnets on its sides swiftly glides by copper coils, producing a levitation force. The track uses a linear synchronous motor, which means the track is synchronized to turn the coils on just before the carrier comes in contact with them, and off once the carrier passes. Sensors are positioned on the side of the track to determine the carrier's position so the appropriate drive coils can be energized. MSFC engineers have conducted tests on the indoor track and a 50-foot outdoor track. The major advantages of launch assist for NASA launch vehicles is that it reduces the weight of the take-off, the landing gear, the wing size, and less propellant resulting in significant cost savings. The US Navy and the British MOD (Ministry of Defense) are planning to use magnetic launch assist for their next generation aircraft carriers as the aircraft launch system. The US Army is considering using this technology for launching target drones for anti-aircraft training.

Design considerations for a pressure-driven multi-stage rocket

Science.gov (United States)

Sauerwein, Steven Craig

2002-01-01

The purpose of this study was to examine the feasibility of using propellant tank pressurization to eliminate the use of high-pressure turbopumps in multi-stage liquid-fueled satellite launchers. Several new technologies were examined to reduce the mass of such a rocket. Composite materials have a greater strength-to-weight ratio than metals and can be used to reduce the weight of rocket propellant tanks and structure. Catalytically combined hydrogen and oxygen can be used to heat pressurization gas, greatly reducing the amount of gas required. Ablatively cooled rocket engines can reduce the complexity and cost of the rocket. Methods were derived to estimate the mass of the various rocket components. These included a method to calculate the amount of gas needed to pressurize a propellant tank by modeling the behavior of the pressurization gas as the liquid propellant flows out of the tank. A way to estimate the mass and size of a ablatively cooled composite cased rocket engine. And a method to model the flight of such a rocket through the atmosphere in conjunction with optimization of the rockets trajectory. The results show that while a liquid propellant rocket using tank pressurization are larger than solid propellant rockets and turbopump driven liquid propellant rockets, they are not impractically large.

The Rocket Balloon (Rocketball): Applications to Science, Technology, and Education

Science.gov (United States)

Esper, Jaime

2009-01-01

Originally envisioned to study upper atmospheric phenomena, the Rocket Balloon system (or Rocketball for short) has utility in a range of applications, including sprite detection and in-situ measurements, near-space measurements and calibration correlation with orbital assets, hurricane observation and characterization, technology testing and validation, ground observation, and education. A salient feature includes the need to reach space and near-space within a critical time-frame and in adverse local meteorological conditions. It can also provide for the execution of technology validation and operational demonstrations at a fraction of the cost of a space flight. In particular, planetary entry probe proof-of-concepts can be examined. A typical Rocketball operational scenario consists of a sounding rocket launch and subsequent deployment of a balloon above a desired location. An obvious advantage of this combination is the additional mission 'hang-time' rendered by the balloon once the sounding rocket flight is completed. The system leverages current and emergent technologies at the NASA Goddard Space Flight Center and other organizations.

A multilayer model to simulate rocket exhaust clouds

Directory of Open Access Journals (Sweden)

Davidson Martins Moreira

2011-01-01

Full Text Available This paper presents the MSDEF (Modelo Simulador da Dispersão de Efluentes de Foguetes, in Portuguese model, which represents the solution for time-dependent advection-diffusion equation applying the Laplace transform considering the Atmospheric Boundary Layer as a multilayer system. This solution allows a time evolution description of the concentration field emitted from a source during a release lasting time tr , and it takes into account deposition velocity, first-order chemical reaction, gravitational settling, precipitation scavenging, and plume rise effect. This solution is suitable for describing critical events relative to accidental release of toxic, flammable, or explosive substances. A qualitative evaluation of the model to simulate rocket exhaust clouds is showed.

Research Technology (ASTP) Rocket Based Combined Cycle (RBCC) Engine

Science.gov (United States)

2004-01-01

Pictured is an artist's concept of the Rocket Based Combined Cycle (RBCC) launch. The RBCC's overall objective is to provide a technology test bed to investigate critical technologies associated with opperational usage of these engines. The program will focus on near term technologies that can be leveraged to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsions systems and ultimately a Single Stage To Orbit (SSTO) air breathing propulsion system.

STS-27 Atlantis, Orbiter Vehicle (OV) 104, at KSC Launch Complex (LC) pad 39B

Science.gov (United States)

1988-01-01

STS-27 Atlantis, Orbiter Vehicle (OV) 104, sits atop the mobile launcher platform at Kennedy Space Center (KSC) Launch Complex (LC) pad 39B. Profile of OV-104 mounted on external tank and flanked by solid rocket boosters (SRBs) is obscured by a flock of seagulls in the foreground. The fixed service structure (FSS) with rotating service structure (RSS) retracted appears in the background. Water resevoir is visible at the base of the launch pad concrete structure.

Combustion diagnosis for analysis of solid propellant rocket abort hazards: Role of spectroscopy

Science.gov (United States)

Gill, W.; Cruz-Cabrera, A. A.; Donaldson, A. B.; Lim, J.; Sivathanu, Y.; Bystrom, E.; Haug, A.; Sharp, L.; Surmick, D. M.

2014-11-01

Solid rocket propellant plume temperatures have been measured using spectroscopic methods as part of an ongoing effort to specify the thermal-chemical-physical environment in and around a burning fragment of an exploded solid rocket at atmospheric pressures. Such specification is needed for launch safety studies where hazardous payloads become involved with large fragments of burning propellant. The propellant burns in an off-design condition producing a hot gas flame loaded with burning metal droplets. Each component of the flame (soot, droplets and gas) has a characteristic temperature, and it is only through the use of spectroscopy that their temperature can be independently identified.

Combustion diagnosis for analysis of solid propellant rocket abort hazards: Role of spectroscopy

International Nuclear Information System (INIS)

Gill, W; Cruz-Cabrera, A A; Bystrom, E; Donaldson, A B; Haug, A; Sharp, L; Lim, J; Sivathanu, Y; Surmick, D M

2014-01-01

Solid rocket propellant plume temperatures have been measured using spectroscopic methods as part of an ongoing effort to specify the thermal-chemical-physical environment in and around a burning fragment of an exploded solid rocket at atmospheric pressures. Such specification is needed for launch safety studies where hazardous payloads become involved with large fragments of burning propellant. The propellant burns in an off-design condition producing a hot gas flame loaded with burning metal droplets. Each component of the flame (soot, droplets and gas) has a characteristic temperature, and it is only through the use of spectroscopy that their temperature can be independently identified

The Road from the NASA Access to Space Study to a Reusable Launch Vehicle

Science.gov (United States)

Powell, Richard W.; Cook, Stephen A.; Lockwood, Mary Kae

1998-01-01

NASA is cooperating with the aerospace industry to develop a space transportation system that provides reliable access-to-space at a much lower cost than is possible with today's launch vehicles. While this quest has been on-going for many years it received a major impetus when the U.S. Congress mandated as part of the 1993 NASA appropriations bill that: "In view of budget difficulties, present and future..., the National Aeronautics and Space Administration shall ... recommend improvements in space transportation." NASA, working with other organizations, including the Department of Transportation, and the Department of Defense identified three major transportation architecture options that were to be evaluated in the areas of reliability, operability and cost. These architectural options were: (1) retain and upgrade the Space Shuttle and the current expendable launch vehicles; (2) develop new expendable launch vehicles using conventional technologies and transition to these new vehicles beginning in 2005; and (3) develop new reusable vehicles using advanced technology, and transition to these vehicles beginning in 2008. The launch needs mission model was based on 1993 projections of civil, defense, and commercial payload requirements. This "Access to Space" study concluded that the option that provided the greatest potential for meeting the cost, operability, and reliability goals was a rocket-powered single-stage-to-orbit fully reusable launch vehicle (RLV) fleet designed with advanced technologies.

Observation of solar HLy-α by K-9M-59 rocket

International Nuclear Information System (INIS)

Watanabe, Norihiko; Tono, Ichiro; Koshio, Takafumi.

1978-01-01

The purpose of the observation is to measure the illumination intensity of solar HLy-α (1216 A) and the altitude distribution of O 2 density in earth atmosphere. Since solar HLy-α is strongly absorbed by O 2 in the earth atmosphere, the extinction of HLy-α depends on the altitude distribution of O 2 density. The K-9M-59 rocket was launched on September 3, 1977. The HLy-α detector was a narrow-band ion chamber filled with NO gas. The result showed that the solar HLy-α irradiance at the rocket altitude of 90 km was 3.5 x 10 11 photons/cm 2 s and the altitude distribution of O 2 density was obtained. (Yoshimori, M.)

Flow-Structural Interaction in Solid Rocket Motors

National Research Council Canada - National Science Library

Murdock, John

2004-01-01

.... The static test failure of the Titan solid rocket motor upgrade (SRMU) that occurred on 1 April, 1991, demonstrated the importance of flow-structural modeling in the design of large, solid rocket motors...

Space Launch System for Exploration and Science

Science.gov (United States)

Klaus, K.

2013-12-01

Introduction: The Space Launch System (SLS) is the most powerful rocket ever built and provides a critical heavy-lift launch capability enabling diverse deep space missions. The exploration class vehicle launches larger payloads farther in our solar system and faster than ever before. The vehicle's 5 m to 10 m fairing allows utilization of existing systems which reduces development risks, size limitations and cost. SLS lift capacity and superior performance shortens mission travel time. Enhanced capabilities enable a myriad of missions including human exploration, planetary science, astrophysics, heliophysics, planetary defense and commercial space exploration endeavors. Human Exploration: SLS is the first heavy-lift launch vehicle capable of transporting crews beyond low Earth orbit in over four decades. Its design maximizes use of common elements and heritage hardware to provide a low-risk, affordable system that meets Orion mission requirements. SLS provides a safe and sustainable deep space pathway to Mars in support of NASA's human spaceflight mission objectives. The SLS enables the launch of large gateway elements beyond the moon. Leveraging a low-energy transfer that reduces required propellant mass, components are then brought back to a desired cislunar destination. SLS provides a significant mass margin that can be used for additional consumables or a secondary payloads. SLS lowers risks for the Asteroid Retrieval Mission by reducing mission time and improving mass margin. SLS lift capacity allows for additional propellant enabling a shorter return or the delivery of a secondary payload, such as gateway component to cislunar space. SLS enables human return to the moon. The intermediate SLS capability allows both crew and cargo to fly to translunar orbit at the same time which will simplify mission design and reduce launch costs. Science Missions: A single SLS launch to Mars will enable sample collection at multiple, geographically dispersed locations and a

Contamination-free sounding rocket Langmuir probe

Science.gov (United States)

Amatucci, W. E.; Schuck, P. W.; Walker, D. N.; Kintner, P. M.; Powell, S.; Holback, B.; Leonhardt, D.

2001-04-01

A technique for removing surface contaminants from a sounding rocket spherical Langmuir probe is presented. Contamination layers present on probe surfaces can skew the collected data, resulting in the incorrect determination of plasma parameters. Despite following the usual probe cleaning techniques that are used prior to a launch, the probe surface can become coated with layers of adsorbed neutral gas in less than a second when exposed to atmosphere. The laboratory tests reported here show that by heating the probe from the interior using a small halogen lamp, adsorbed neutral particles can be removed from the probe surface, allowing accurate plasma parameter measurements to be made.

Photogrammetry and ballistic analysis of a high-flying projectile in the STS-124 space shuttle launch

Science.gov (United States)

Metzger, Philip T.; Lane, John E.; Carilli, Robert A.; Long, Jason M.; Shawn, Kathy L.

2010-07-01

A method combining photogrammetry with ballistic analysis is demonstrated to identify flying debris in a rocket launch environment. Debris traveling near the STS-124 Space Shuttle was captured on cameras viewing the launch pad within the first few seconds after launch. One particular piece of debris caught the attention of investigators studying the release of flame trench fire bricks because its high trajectory could indicate a flight risk to the Space Shuttle. Digitized images from two pad perimeter high-speed 16-mm film cameras were processed using photogrammetry software based on a multi-parameter optimization technique. Reference points in the image were found from 3D CAD models of the launch pad and from surveyed points on the pad. The three-dimensional reference points were matched to the equivalent two-dimensional camera projections by optimizing the camera model parameters using a gradient search optimization technique. Using this method of solving the triangulation problem, the xyz position of the object's path relative to the reference point coordinate system was found for every set of synchronized images. This trajectory was then compared to a predicted trajectory while performing regression analysis on the ballistic coefficient and other parameters. This identified, with a high degree of confidence, the object's material density and thus its probable origin within the launch pad environment. Future extensions of this methodology may make it possible to diagnose the underlying causes of debris-releasing events in near-real time, thus improving flight safety.

Electron spectra over discrete auroras as measured by the Substorm-GEOS rockets

International Nuclear Information System (INIS)

Sandahl, I.; Eliasson, L.; Lundin, R.

1980-01-01

Results from the first two Substorm-GEOS rockets are presented. These rockets, as well as the third one, were launched from ESRANGE on January 27, 1979 into different substorm phases. The first rocket went into an active pre-breakup evening arc and the second one into a breakup close to magnetic midnight. Electron spectra of downcoming particles measured by a narrow energy bandwidth detector show very narrow energy peaks as soon as the integral energy fluxes are high. These peaks always show populations of two different characteristic energies above the peak energy. One of the populations has properties similar to those found in the boundary layer plasma and the other one seems to be of plasma sheet origin. The plasma sheet like population is also seen where there are no signs of energy peaks, for example equatorward of the arc. The boundary layer plasma is exclusively connected with the signatures of acceleration. (Auth.)

Rocket science

International Nuclear Information System (INIS)

Upson Sandra

2011-01-01

Expanding across the Solar System will require more than a simple blast off, a range of promising new propulsion technologies are being investigated by ex- NASA shuttle astronaut Chang Diaz. He is developing an alternative to chemical rockets, called VASIMR -Variable Specific Impulse Magnetoplasm Rocket. In 2012 Ad Astra plans to test a prototype, using solar power rather than nuclear, on the International Space Station. Development of this rocket for human space travel is discussed. The nuclear reactor's heat would be converted into electricity in an electric rocket such as VASIMR, and at the peak of nuclear rocket research thrust levels of almost one million newtons were reached.

Air-Powered Rockets.

Science.gov (United States)

Rodriguez, Charley; Raynovic, Jim

This document describes methods for designing and building two types of rockets--rockets from paper and rockets from bottles. Devices used for measuring the heights that the rockets obtain are also discussed. (KHR)

The beginning of Space Life Science in China exploration rockets for biological experiment during 1960's

Science.gov (United States)

Jiang, Peidong; Zhang, Jingxue

The first step of space biological experiment in China was a set of five exploration rockets launched during 1964 to 1966, by Shanghai Institute of Machine and Electricity, and Institute of Biophysics of The Chinese Academy of Sciences. Three T-7AS1rockets for rats, mice and other samples in a biological cabin were launched and recovered safely in July of 1964 and June of 1965. Two T-7AS2rockets for dog, rats, mice and other samples in a biological cabin were launched and recovered safely in July of 1966. Institute of Biophysics in charged of the general design of biological experiments, telemetry of physiological parameters, and selection and training of experiment animals. The samples on-board were: rats, mice, dogs, and test tubes with fruit fly, enzyme, bacteria, E. Coli., lysozyme, bacteriaphage, RNAase, DNAase, crystals of enzyme, etc. Physiological, biochemical, bacte-riological, immunological, genetic, histochemical studies had been conducted, in cellular and sub cellular level. The postures of rat and dog were monitored during flight and under weight-lessness. Physiological parameters of ECG, blood pressure, respiration rate, body temperature were recorded. A dog named"Xiao Bao"was flight in 1966 with video monitor, life support system and conditioned reflex equipment. It flighted for more than 20 minutes and about 70km high. After 40 years, the experimental data recorded of its four physiological parameters during the flight process was reviewed. The change of 4 parameters during various phase of total flight process were compared, analyzed and discussed.

Building and Leading the Next Generation of Exploration Launch Vehicles

Science.gov (United States)

Cook, Stephen A.; Vanhooser, Teresa

2010-01-01

NASA s Constellation Program is depending on the Ares Projects to deliver the crew and cargo launch capabilities needed to send human explorers to the Moon and beyond. Ares I and V will provide the core space launch capabilities needed to continue providing crew and cargo access to the International Space Station (ISS), and to build upon the U.S. history of human spaceflight to the Moon and beyond. Since 2005, Ares has made substantial progress on designing, developing, and testing the Ares I crew launch vehicle and has continued its in-depth studies of the Ares V cargo launch vehicle. In 2009, the Ares Projects plan to: conduct the first flight test of Ares I, test-fire the Ares I first stage solid rocket motor; build the first integrated Ares I upper stage; continue testing hardware for the J-2X upper stage engine, and continue refining the design of the Ares V cargo launch vehicle. These efforts come with serious challenges for the project leadership team as it continues to foster a culture of ownership and accountability, operate with limited funding, and works to maintain effective internal and external communications under intense external scrutiny.

US Rocket Propulsion Industrial Base Health Metrics

Science.gov (United States)

Doreswamy, Rajiv

2013-01-01

The number of active liquid rocket engine and solid rocket motor development programs has severely declined since the "space race" of the 1950s and 1960s center dot This downward trend has been exacerbated by the retirement of the Space Shuttle, transition from the Constellation Program to the Space launch System (SLS) and similar activity in DoD programs center dot In addition with consolidation in the industry, the rocket propulsion industrial base is under stress. To Improve the "health" of the RPIB, we need to understand - The current condition of the RPIB - How this compares to past history - The trend of RPIB health center dot This drives the need for a concise set of "metrics" - Analogous to the basic data a physician uses to determine the state of health of his patients - Easy to measure and collect - The trend is often more useful than the actual data point - Can be used to focus on problem areas and develop preventative measures The nation's capability to conceive, design, develop, manufacture, test, and support missions using liquid rocket engines and solid rocket motors that are critical to its national security, economic health and growth, and future scientific needs. center dot The RPIB encompasses US government, academic, and commercial (including industry primes and their supplier base) research, development, test, evaluation, and manufacturing capabilities and facilities. center dot The RPIB includes the skilled workforce, related intellectual property, engineering and support services, and supply chain operations and management. This definition touches the five main segments of the U.S. RPIB as categorized by the USG: defense, intelligence community, civil government, academia, and commercial sector. The nation's capability to conceive, design, develop, manufacture, test, and support missions using liquid rocket engines and solid rocket motors that are critical to its national security, economic health and growth, and future scientific needs

Deep Impact Delta II Launch Vehicle Cracked Thick Film Coating on Electronic Packages Technical Consultation Report

Science.gov (United States)

Cameron, Kenneth D.; Kichak, Robert A.; Piascik, Robert S.; Leidecker, Henning W.; Wilson, Timmy R.

2009-01-01

The Deep Impact spacecraft was launched on a Boeing Delta II rocket from Cape Canaveral Air Force Station (CCAFS) on January 12, 2005. Prior to the launch, the Director of the Office of Safety and Mission Assurance (OS&MA) requested the NASA Engineering and Safety Center (NESC) lead a team to render an independent opinion on the rationale for flight and the risk code assignments for the hazard of cracked Thick Film Assemblies (TFAs) in the E-packages of the Delta II launch vehicle for the Deep Impact Mission. The results of the evaluation are contained in this report.

Rocket measurement of auroral partial parallel distribution functions

Science.gov (United States)

Lin, C.-A.

1980-01-01

The auroral partial parallel distribution functions are obtained by using the observed energy spectra of electrons. The experiment package was launched by a Nike-Tomahawk rocket from Poker Flat, Alaska over a bright auroral band and covered an altitude range of up to 180 km. Calculated partial distribution functions are presented with emphasis on their slopes. The implications of the slopes are discussed. It should be pointed out that the slope of the partial parallel distribution function obtained from one energy spectra will be changed by superposing another energy spectra on it.

Measurement of IR atmospheric band dayglow by S-520-4 rocket

International Nuclear Information System (INIS)

Makino, Tadao; Yamamoto, Hiromasa; Sekiguchi, Hiroyuki

1984-01-01

The measurement of IR atmospheric band dayglow was made by rocket S-520-4 flown from Uchinoura at 1000 JST on Sept. 5, 1981. The instrument loaded on the rocket was the same type as the one loaded on EXOS-C satellite which will be launched in 1984 in order to observe the mesospheric ozone. This rocket experiment was performed for the purpose of testing the functions of this instrument in flight. The 1.27 μm filter radiometer consisted of three plane mirros, a camera lens, a chopper and a PbS detector array. The PbS array (4x5=20 elements) was operated at about -4 0 C with a thermoelectric cooler. We obtained the following results from the rocket experiment: (i) this instrument worked well during the flight, (ii) the intensities of the solar radiation scattered by the sea and clouds were obtained at 1.27 μm, and (iii) the baffle designed to permit the daytime measurement of the atmospheric emission could attenuate the off-axis radiation as weak as possible. The altitude distribution of the daytime mesospheric ozone density derived from the downleg data was in agreement with the previous profile obtained in twilight condition. (author)

STS-91 Launch of Discovery from Launch Pad 39-A

Science.gov (United States)

1998-01-01

Searing the early evening sky with its near sun-like rocket exhaust, the Space Shuttle Discovery lifts off from Launch Pad 39A at 6:06:24 p.m. EDT June 2 on its way to the Mir space station. On board Discovery are Mission Commander Charles J. Precourt; Pilot Dominic L. Gorie; and Mission Specialists Wendy B. Lawrence, Franklin R. Chang-Diaz, Janet Lynn Kavandi and Valery Victorovitch Ryumin. The nearly 10-day mission will feature the ninth and final Shuttle docking with the Russian space station Mir, the first Mir docking for the Space Shuttle orbiter Discovery, the first on-orbit test of the Alpha Magnetic Spectrometer (AMS), and the first flight of the new Space Shuttle super lightweight external tank. Astronaut Andrew S. W. Thomas will be returning to Earth as a STS-91 crew member after living more than four months aboard Mir.

Evaluation of the Effect of Exhausts from Liquid and Solid Rockets on Ozone Layer

Science.gov (United States)

Yamagiwa, Yoshiki; Ishimaki, Tetsuya

This paper reports the analytical results of the influences of solid rocket and liquid rocket exhausts on ozone layer. It is worried about that the exhausts from solid propellant rockets cause the ozone depletion in the ozone layer. Some researchers try to develop the analytical model of ozone depletion by rocket exhausts to understand its physical phenomena and to find the effective design of rocket to minimize its effect. However, these models do not include the exhausts from liquid rocket although there are many cases to use solid rocket boosters with a liquid rocket at the same time in practical situations. We constructed combined analytical model include the solid rocket exhausts and liquid rocket exhausts to analyze their effects. From the analytical results, we find that the exhausts from liquid rocket suppress the ozone depletion by solid rocket exhausts.

Small Rocket/Spacecraft Technology (SMART) Platform

Science.gov (United States)

Esper, Jaime; Flatley, Thomas P.; Bull, James B.; Buckley, Steven J.

2011-01-01

The NASA Goddard Space Flight Center (GSFC) and the Department of Defense Operationally Responsive Space (ORS) Office are exercising a multi-year collaborative agreement focused on a redefinition of the way space missions are designed and implemented. A much faster, leaner and effective approach to space flight requires the concerted effort of a multi-agency team tasked with developing the building blocks, both programmatically and technologically, to ultimately achieve flights within 7-days from mission call-up. For NASA, rapid mission implementations represent an opportunity to find creative ways for reducing mission life-cycle times with the resulting savings in cost. This in tum enables a class of missions catering to a broader audience of science participants, from universities to private and national laboratory researchers. To that end, the SMART (Small Rocket/Spacecraft Technology) micro-spacecraft prototype demonstrates an advanced avionics system with integrated GPS capability, high-speed plug-and-playable interfaces, legacy interfaces, inertial navigation, a modular reconfigurable structure, tunable thermal technology, and a number of instruments for environmental and optical sensing. Although SMART was first launched inside a sounding rocket, it is designed as a free-flyer.

Ionospheric disturbances induced by a missile launched from North Korea on 12 December 2012

Science.gov (United States)

Kakinami, Yoshihiro; Yamamoto, Masayuki; Chen, Chia-Hung; Watanabe, Shigeto; Lin, Charles; Liu, Jenn-Yanq; Habu, Hiroto

2013-08-01

disturbances caused by a missile launched from North Korea on 12 December 2012 were investigated by using the GPS total electron content (TEC). The spatial characteristic of the front edge of V-shaped disturbances produced by missiles and rockets was first determined. Considering the launch direction and the height of estimated ionospheric points at which GPS radio signal pierces the ionosphere, the missile passed through the ionosphere at heights of 391, 425, and 435 km at 0056:30, 0057:00, and 0057:30 UT, respectively. The observed velocities of the missile were 2.8 and 3.2 km/s at that time, which was estimated from the traveling speed of the front edge of V-shaped disturbances. Westward and eastward V-shaped disturbances propagated at 1.8-2.6 km/s. The phase velocities of the westward and eastward V-shaped disturbances were much faster than the speed of acoustic waves reported in previous studies, suggesting that sources other than acoustic waves may have played an important role. Furthermore, the plasma density depletion that is often observed following missile and rocket launches was not found. This suggests that the depletion resulting from the missile's exhaust was not strong enough to be observed in the TEC distribution in the topside ionosphere.

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...

Investigation of Advanced Propellants to Enable Single Stage to Orbit Launch Vehicles

Science.gov (United States)

2006-10-30

ERS-PAS-2006-205) 13. SUPPLEMENTARY NOTES Graduate work for California State University, Fresno 14. ABSTRACT Single-Stage-To-Orbit ( SSTO ...and maintained. Despite well-funded development efforts, no SSTO vehicles have been fielded to date. Existing chemical rocket and vehicle...technologies do not enable feasible SSTO designs. In the future, new propellants with advanced properties could enable SSTO launch vehicles. A parametric

Telemetry Boards Interpret Rocket, Airplane Engine Data

Science.gov (United States)

2009-01-01

For all the data gathered by the space shuttle while in orbit, NASA engineers are just as concerned about the information it generates on the ground. From the moment the shuttle s wheels touch the runway to the break of its electrical umbilical cord at 0.4 seconds before its next launch, sensors feed streams of data about the status of the vehicle and its various systems to Kennedy Space Center s shuttle crews. Even while the shuttle orbiter is refitted in Kennedy s orbiter processing facility, engineers constantly monitor everything from power levels to the testing of the mechanical arm in the orbiter s payload bay. On the launch pad and up until liftoff, the Launch Control Center, attached to the large Vehicle Assembly Building, screens all of the shuttle s vital data. (Once the shuttle clears its launch tower, this responsibility shifts to Mission Control at Johnson Space Center, with Kennedy in a backup role.) Ground systems for satellite launches also generate significant amounts of data. At Cape Canaveral Air Force Station, across the Banana River from Kennedy s location on Merritt Island, Florida, NASA rockets carrying precious satellite payloads into space flood the Launch Vehicle Data Center with sensor information on temperature, speed, trajectory, and vibration. The remote measurement and transmission of systems data called telemetry is essential to ensuring the safe and successful launch of the Agency s space missions. When a launch is unsuccessful, as it was for this year s Orbiting Carbon Observatory satellite, telemetry data also provides valuable clues as to what went wrong and how to remedy any problems for future attempts. All of this information is streamed from sensors in the form of binary code: strings of ones and zeros. One small company has partnered with NASA to provide technology that renders raw telemetry data intelligible not only for Agency engineers, but also for those in the private sector.

Simulation and experimental research on line throwing rocket with flight

Directory of Open Access Journals (Sweden)

Wen-bin Gu

2014-06-01

Full Text Available The finite segment method is used to model the line throwing rocket system. A dynamic model of line throwing rocket with flight motion based on Kane's method is presented by the kinematics description of the system and the consideration of the forces acting on the system. The experiment designed according to the parameters of the dynamic model is made. The simulation and experiment results, such as range, velocity and flight time, are compared and analyzed. The simulation results are basically agreed with the test data, which shows that the flight motion of the line throwing rocket can be predicted by the dynamic model. A theoretical model and guide for the further research on the disturbance of rope and the guidance, flight control of line throwing rocket are provided by the dynamic modeling.

Computer Program for Analysis, Design and Optimization of Propulsion, Dynamics, and Kinematics of Multistage Rockets

Science.gov (United States)

Lali, Mehdi

2009-03-01

A comprehensive computer program is designed in MATLAB to analyze, design and optimize the propulsion, dynamics, thermodynamics, and kinematics of any serial multi-staging rocket for a set of given data. The program is quite user-friendly. It comprises two main sections: "analysis and design" and "optimization." Each section has a GUI (Graphical User Interface) in which the rocket's data are entered by the user and by which the program is run. The first section analyzes the performance of the rocket that is previously devised by the user. Numerous plots and subplots are provided to display the performance of the rocket. The second section of the program finds the "optimum trajectory" via billions of iterations and computations which are done through sophisticated algorithms using numerical methods and incremental integrations. Innovative techniques are applied to calculate the optimal parameters for the engine and designing the "optimal pitch program." This computer program is stand-alone in such a way that it calculates almost every design parameter in regards to rocket propulsion and dynamics. It is meant to be used for actual launch operations as well as educational and research purposes.

Coupled Lagrangian impingement spray model for doublet impinging injectors under liquid rocket engine operating conditions

Directory of Open Access Journals (Sweden)

Qiang WEI

2017-08-01

Full Text Available To predict the effect of the liquid rocket engine combustion chamber conditions on the impingement spray, the conventional uncoupled spray model for impinging injectors is extended by considering the coupling of the jet impingement process and the ambient gas field. The new coupled model consists of the plain-orifice sub-model, the jet-jet impingement sub-model and the droplet collision sub-model. The parameters of the child droplet are determined with the jet-jet impingement sub-model using correlations about the liquid jet parameters and the chamber conditions. The overall model is benchmarked under various impingement angles, jet momentum and off-center ratios. Agreement with the published experimental data validates the ability of the model to predict the key spray characteristics, such as the mass flux and mixture ratio distributions in quiescent air. Besides, impinging sprays under changing ambient pressure and non-uniform gas flow are investigated to explore the effect of liquid rocket engine chamber conditions. First, a transient impingement spray during engine start-up phase is simulated with prescribed pressure profile. The minimum average droplet diameter is achieved when the orifices work in cavitation state, and is about 30% smaller than the steady single phase state. Second, the effect of non-uniform gas flow produces off-center impingement and the rotated spray fan by 38°. The proposed model suggests more reasonable impingement spray characteristics than the uncoupled one and can be used as the first step in the complex simulation of coupling impingement spray and combustion in liquid rocket engines.

Coupled Lagrangian impingement spray model for doublet impinging injectors under liquid rocket engine operating conditions

Institute of Scientific and Technical Information of China (English)

Qiang WEI; Guozhu LIANG

2017-01-01

To predict the effect of the liquid rocket engine combustion chamber conditions on the impingement spray,the conventional uncoupled spray model for impinging injectors is extended by considering the couplingof the jet impingement process and the ambient gas field.The new coupled model consists of the plain-orifice sub-model,the jet-jet impingement sub-model and the droplet collision sub-model.The parameters of the child droplet are determined with the jet-jet impingement sub-model using correlations about the liquid jet parameters and the chamber conditions.The overall model is benchmarked under various impingement angles,jet momentum and offcenter ratios.Agreement with the published experimental data validates the ability of the model to predict the key spray characteristics,such as the mass flux and mixture ratio distributions in quiescent air.Besides,impinging sprays under changing ambient pressure and non-uniform gas flow are investigated to explore the effect of liquid rocket engine chamber conditions.First,a transient impingement spray during engine start-up phase is simulated with prescribed pressure profile.The minimum average droplet diameter is achieved when the orifices work in cavitation state,and is about 30％ smaller than the steady single phase state.Second,the effect of non-uniform gas flow produces off-center impingement and the rotated spray fan by 38°.The proposed model suggests more reasonable impingement spray characteristics than the uncoupled one and can be used as the first step in the complex simulation of coupling impingement spray and combustion in liquid rocket engines.

Ares V: Game Changer for National Security Launch

Science.gov (United States)

Sumrall, Phil; Morris, Bruce

2009-01-01

NASA is designing the Ares V cargo launch vehicle to vastly expand exploration of the Moon begun in the Apollo program and enable the exploration of Mars and beyond. As the largest launcher in history, Ares V also represents a national asset offering unprecedented opportunities for new science, national security, and commercial missions of unmatched size and scope. The Ares V is the heavy-lift component of NASA's dual-launch architecture that will replace the current space shuttle fleet, complete the International Space Station, and establish a permanent human presence on the Moon as a stepping-stone to destinations beyond. During extensive independent and internal architecture and vehicle trade studies as part of the Exploration Systems Architecture Study (ESAS), NASA selected the Ares I crew launch vehicle and the Ares V to support future exploration. The smaller Ares I will launch the Orion crew exploration vehicle with four to six astronauts into orbit. The Ares V is designed to carry the Altair lunar lander into orbit, rendezvous with Orion, and send the mated spacecraft toward lunar orbit. The Ares V will be the largest and most powerful launch vehicle in history, providing unprecedented payload mass and volume to establish a permanent lunar outpost and explore significantly more of the lunar surface than was done during the Apollo missions. The Ares V consists of a Core Stage, two Reusable Solid Rocket Boosters (RSRBs), Earth Departure Stage (EDS), and a payload shroud. For lunar missions, the shroud would cover the Lunar Surface Access Module (LSAM). The Ares V Core Stage is 33 feet in diameter and 212 feet in length, making it the largest rocket stage ever built. It is the same diameter as the Saturn V first stage, the S-IC. However, its length is about the same as the combined length of the Saturn V first and second stages. The Core Stage uses a cluster of five Pratt & Whitney Rocketdyne RS-68B rocket engines, each supplying about 700,000 pounds of thrust

Performance of an Axisymmetric Rocket Based Combined Cycle Engine During Rocket Only Operation Using Linear Regression Analysis

Science.gov (United States)

Smith, Timothy D.; Steffen, Christopher J., Jr.; Yungster, Shaye; Keller, Dennis J.

1998-01-01

The all rocket mode of operation is shown to be a critical factor in the overall performance of a rocket based combined cycle (RBCC) vehicle. An axisymmetric RBCC engine was used to determine specific impulse efficiency values based upon both full flow and gas generator configurations. Design of experiments methodology was used to construct a test matrix and multiple linear regression analysis was used to build parametric models. The main parameters investigated in this study were: rocket chamber pressure, rocket exit area ratio, injected secondary flow, mixer-ejector inlet area, mixer-ejector area ratio, and mixer-ejector length-to-inlet diameter ratio. A perfect gas computational fluid dynamics analysis, using both the Spalart-Allmaras and k-omega turbulence models, was performed with the NPARC code to obtain values of vacuum specific impulse. Results from the multiple linear regression analysis showed that for both the full flow and gas generator configurations increasing mixer-ejector area ratio and rocket area ratio increase performance, while increasing mixer-ejector inlet area ratio and mixer-ejector length-to-diameter ratio decrease performance. Increasing injected secondary flow increased performance for the gas generator analysis, but was not statistically significant for the full flow analysis. Chamber pressure was found to be not statistically significant.

Designing Liquid Rocket Engine Injectors for Performance, Stability, and Cost

Science.gov (United States)

Westra, Douglas G.; West, Jeffrey S.

2014-01-01

NASA is developing the Space Launch System (SLS) for crewed exploration missions beyond low Earth orbit. Marshall Space Flight Center (MSFC) is designing rocket engines for the SLS Advanced Booster (AB) concepts being developed to replace the Shuttle-derived solid rocket boosters. One AB concept uses large, Rocket-Propellant (RP)-fueled engines that pose significant design challenges. The injectors for these engines require high performance and stable operation while still meeting aggressive cost reduction goals for access to space. Historically, combustion stability problems have been a critical issue for such injector designs. Traditional, empirical injector design tools and methodologies, however, lack the ability to reliably predict complex injector dynamics that often lead to combustion stability. Reliance on these tools alone would likely result in an unaffordable test-fail-fix cycle for injector development. Recently at MSFC, a massively parallel computational fluid dynamics (CFD) program was successfully applied in the SLS AB injector design process. High-fidelity reacting flow simulations were conducted for both single-element and seven-element representations of the full-scale injector. Data from the CFD simulations was then used to significantly augment and improve the empirical design tools, resulting in a high-performance, stable injector design.

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.

Launch Services, a Proven Model

Science.gov (United States)

Trafton, W. C.; Simpson, J.

2002-01-01

From a commercial perspective, the ability to justify "leap frog" technology such as reusable systems has been difficult to justify because the estimated 5B to 10B investment is not supported in the current flat commercial market coupled with an oversupply of launch service suppliers. The market simply does not justify investment of that magnitude. Currently, next generation Expendable Launch Systems, including Boeing's Delta IV, Lockheed Martin's Atlas 5, Ariane V ESCA and RSC's H-IIA are being introduced into operations signifying that only upgrades to proven systems are planned to meet the changes in anticipated satellite demand (larger satellites, more lifetime, larger volumes, etc.) in the foreseeable future. We do not see a new fleet of ELVs emerging beyond that which is currently being introduced, only continuous upgrades of the fleet to meet the demands. To induce a radical change in the provision of launch services, a Multinational Government investment must be made and justified by World requirements. The commercial market alone cannot justify such an investment. And if an investment is made, we cannot afford to repeat previous mistakes by relying on one system such as shuttle for commercial deployment without having any back-up capability. Other issues that need to be considered are national science and security requirements, which to a large extent fuels the Japanese, Chinese, Indian, Former Soviet Union, European and United States space transportation entries. Additionally, this system must support or replace current Space Transportation Economies with across-the-board benefits. For the next 10 to 20 years, Multinational cooperation will be in the form of piecing together launch components and infrastructure to supplement existing launch systems and reducing the amount of non-recurring investment while meeting the future requirements of the End-User. Virtually all of the current systems have some form of multinational participation: Sea Launch

The Extended Duration Sounding Rocket (EDSR): Low Cost Science and Technology Missions

Science.gov (United States)

Cruddace, R. G.; Chakrabarti, S.; Cash, W.; Eberspeaker, P.; Figer, D.; Figueroa, O.; Harris, W.; Kowalski, M.; Maddox, R.; Martin, C.; McCammon, D.; Nordsieck, K.; Polidan, R.; Sanders, W.; Wilkinson, E.; Asrat

2011-12-01

The 50-year old NASA sounding rocket (SR) program has been successful in launching scientific payloads into space frequently and at low cost with a 85% success rate. In 2008 the NASA Astrophysics Sounding Rocket Assessment Team (ASRAT), set up to review the future course of the SR program, made four major recommendations, one of which now called Extended Duration Sounding Rocket (EDSR). ASRAT recommended a system capable of launching science payloads (up to 420 kg) into low Earth orbit frequently (1/yr) at low cost, with a mission duration of approximately 30 days. Payload selection would be based on meritorious high-value science that can be performed by migrating sub-orbital payloads to orbit. Establishment of this capability is a essential for NASA as it strives to advance technical readiness and lower costs for risk averse Explorers and flagship missions in its pursuit of a balanced and sustainable program and achieve big science goals within a limited fiscal environment. The development of a new generation of small, low-cost launch vehicles (SLV), primarily the SpaceX Falcon 1 and the Orbital Sciences Minotaur I has made this concept conceivable. The NASA Wallops Flight Facility (WFF)conducted a detailed engineering concept study, aimed at defining the technical characteristics of all phases of a mission, from design, procurement, assembly, test, integration and mission operations. The work was led by Dr. Raymond Cruddace, a veteran of the SR program and the prime mover of the EDSR concept. The team investigated details such as, the "FAA licensed contract" for launch service procurement, with WFF and NASA SMD being responsible for mission assurance which results in a factor of two cost savings over the current approach. These and other creative solutions resulted in a proof-of-concept Class D mission design that could have a sustained launch rate of at least 1/yr, a mission duration of up to about 3 months, and a total cost of $25-30 million for each mission

Statistical methods for launch vehicle guidance, navigation, and control (GN&C) system design and analysis

Science.gov (United States)

Rose, Michael Benjamin

A novel trajectory and attitude control and navigation analysis tool for powered ascent is developed. The tool is capable of rapid trade-space analysis and is designed to ultimately reduce turnaround time for launch vehicle design, mission planning, and redesign work. It is streamlined to quickly determine trajectory and attitude control dispersions, propellant dispersions, orbit insertion dispersions, and navigation errors and their sensitivities to sensor errors, actuator execution uncertainties, and random disturbances. The tool is developed by applying both Monte Carlo and linear covariance analysis techniques to a closed-loop, launch vehicle guidance, navigation, and control (GN&C) system. The nonlinear dynamics and flight GN&C software models of a closed-loop, six-degree-of-freedom (6-DOF), Monte Carlo simulation are formulated and developed. The nominal reference trajectory (NRT) for the proposed lunar ascent trajectory is defined and generated. The Monte Carlo truth models and GN&C algorithms are linearized about the NRT, the linear covariance equations are formulated, and the linear covariance simulation is developed. The performance of the launch vehicle GN&C system is evaluated using both Monte Carlo and linear covariance techniques and their trajectory and attitude control dispersion, propellant dispersion, orbit insertion dispersion, and navigation error results are validated and compared. Statistical results from linear covariance analysis are generally within 10% of Monte Carlo results, and in most cases the differences are less than 5%. This is an excellent result given the many complex nonlinearities that are embedded in the ascent GN&C problem. Moreover, the real value of this tool lies in its speed, where the linear covariance simulation is 1036.62 times faster than the Monte Carlo simulation. Although the application and results presented are for a lunar, single-stage-to-orbit (SSTO), ascent vehicle, the tools, techniques, and mathematical

Next Generation Launch Technology Program Lessons Learned

Science.gov (United States)

Cook, Stephen; Tyson, Richard

2005-01-01

In November 2002, NASA revised its Integrated Space Transportation Plan (ISTP) to evolve the Space Launch Initiative (SLI) to serve as a theme for two emerging programs. The first of these, the Orbital Space Plane (OSP), was intended to provide crew-escape and crew-transfer functions for the ISS. The second, the NGLT Program, developed technologies needed for safe, routine space access for scientific exploration, commerce, and national defense. The NGLT Program was comprised of 12 projects, ranging from fundamental high-temperature materials research to full-scale engine system developments (turbine and rocket) to scramjet flight test. The Program included technology advancement activities with a broad range of objectives, ultimate applications/timeframes, and technology maturity levels. An over-arching Systems Engineering and Analysis (SE&A) approach was employed to focus technology advancements according to a common set of requirements. Investments were categorized into three segments of technology maturation: propulsion technologies, launch systems technologies, and SE&A.

Transfer Matrix Method for the Determination of the Natural Vibration Characteristics of Realistic Thrusting Launch Vehicle—Part I

Directory of Open Access Journals (Sweden)

Laith K. Abbas

2013-01-01

Full Text Available The feasibility of using the transfer matrix method (TMM to compute the natural vibration characteristics of a flexible rocket/satellite launch vehicle is explored theoretically. In the approach to the problem, a nonuniform free-free Timoshenko and Euler-Bernoulli beamlike structure is modeled. A provision is made to take into consideration the effects of shear deformation and rotary inertia. Large thrust-to-weight ratio leads to large axial accelerations that result in an axial inertia load distribution from nose to tail which causes the development of significant compressive forces along the length of the launch vehicle. Therefore, it is important to take into account this effect in the transverse vibration model. Once the transfer matrix of a single component has been obtained, the product of all component matrices composes the matrix of the entire structure. The frequency equation and mode shape are formulated in terms of the elements of the structural matrices. Flight test and analytical results validate the present TMM formulas.

A Model for the Sounding Rocket Measurement on an Ionospheric E-F Valley at the Hainan Low Latitude Station

International Nuclear Information System (INIS)

Wang Zheng; Shi Jiankui; Guan Yibing; Liu Chao; Zhu Guangwu; Torkar Klaus; Fredrich Martin

2014-01-01

To understand the physics of an ionospheric E-F valley, a new overlapping three-Chapman-layer model is developed to interpret the sounding rocket measurement in the morning (sunrise) on May 7, 2011 at the Hainan low latitude ionospheric observation station (19.5°N, 109.1°E). From our model, the valley width, depth and height are 43.0 km, 62.9% and 121.0 km, respectively. From the sounding rocket observation, the valley width, depth and height are 42.2 km, 47.0% and 123.5 km, respectively. The model results are well consistent with the sounding rocket observation. The observed E-F valley at Hainan station is very wide and deep, and rapid development of the photochemical process in the ionosphere should be the underlying reason. (astrophysics and space plasma)

Three Dimensional Numerical Simulation of Rocket-based Combined-cycle Engine Response During Mode Transition Events

Science.gov (United States)

Edwards, Jack R.; McRae, D. Scott; Bond, Ryan B.; Steffan, Christopher (Technical Monitor)

2003-01-01

The GTX program at NASA Glenn Research Center is designed to develop a launch vehicle concept based on rocket-based combined-cycle (RBCC) propulsion. Experimental testing, cycle analysis, and computational fluid dynamics modeling have all demonstrated the viability of the GTX concept, yet significant technical issues and challenges still remain. Our research effort develops a unique capability for dynamic CFD simulation of complete high-speed propulsion devices and focuses this technology toward analysis of the GTX response during critical mode transition events. Our principal attention is focused on Mode 1/Mode 2 operation, in which initial rocket propulsion is transitioned into thermal-throat ramjet propulsion. A critical element of the GTX concept is the use of an Independent Ramjet Stream (IRS) cycle to provide propulsion at Mach numbers less than 3. In the IRS cycle, rocket thrust is initially used for primary power, and the hot rocket plume is used as a flame-holding mechanism for hydrogen fuel injected into the secondary air stream. A critical aspect is the establishment of a thermal throat in the secondary stream through the combination of area reduction effects and combustion-induced heat release. This is a necessity to enable the power-down of the rocket and the eventual shift to ramjet mode. Our focus in this first year of the grant has been in three areas, each progressing directly toward the key initial goal of simulating thermal throat formation during the IRS cycle: CFD algorithm development; simulation of Mode 1 experiments conducted at Glenn's Rig 1 facility; and IRS cycle simulations. The remainder of this report discusses each of these efforts in detail and presents a plan of work for the next year.

Holographic elements and curved slit used to enlarge field of view in rocket detection system

Science.gov (United States)

Breton, Mélanie; Fortin, Jean; Lessard, Roger A.; Châteauneuf, Marc

2006-09-01

Rocket detection over a wide field of view is an important issue in the protection of light armored vehicle. Traditionally, the detection occurs in UV band, but recent studies have shown the existence of significant emission peaks in the visible and near infrared at rocket launch time. The use of the visible region is interesting in order to reduce the weight and cost of systems. Current methods to detect those specific peaks involve use of interferometric filters. However, they fail to combine wide angle with wavelength selectivity. A linear array of volume holographic elements combined with a curved exit slit is proposed for the development of a wide field of view sensor for the detection of solid propellant motor launch flash. The sensor is envisaged to trigger an active protection system. On the basis of geometric theory, a system has been designed. It consists of a collector, a linear array of holographic elements, a curved slit and a detector. The collector is an off-axis parabolic mirror. Holographic elements are recorded subdividing a hologram film in regions, each individually exposed with a different incidence angle. All regions have a common diffraction angle. The incident angle determines the instantaneous field of view of the elements. The volume hologram performs the function of separating and focusing the diffracted beam on an image plane to achieve wavelength filtering. Conical diffraction property is used to enlarge the field of view in elevation. A curved slit was designed to correspond to oblique incidence of the holographic linear array. It is situated at the image plane and filters the diffracted spectrum toward the sensor. The field of view of the design was calculated to be 34 degrees. This was validated by a prototype tested during a field trial. Results are presented and analyzed. The system succeeded in detecting the rocket launch flash at desired fields of view.

Rocket and ground-based study of an auroral breakup event

International Nuclear Information System (INIS)

Marklund, G.

1982-02-01

On 27 January, 1979 the substorm-GEOS rocket S23H was launched from ESRANGE, Kiruna, shortly after the onset of an intense magnetospheric substorm over northern Scandinavia. Rocket electric field and particle observations have been used to calculate ionospheric currents and heating rates. These results are generally consistent with the ground magnetic and optical observations. An important finding emerging from a comparison of this event with a pre-breakup event earlier on this day is that the ionospheric substorm-related electric field could be split up into two parts, namely: 1) an ambient LT dependent field, probably of magnetospheric origin 2) superimposed on this a small-scale electric field associated with the bright auroral structures, being southward for both events. This is shown to have important consequences on the location of the ionospheric currents and the Joule energy discussion relative to the auroral forms. (Author)

Observation of cosmic hard x-ray by L-3H-9 rocket

International Nuclear Information System (INIS)

Hayakawa, Sachio; Makino, Fumiyoshi; Matsui, Yutaka; Fukada, Yutaka.

1978-01-01

It has been considered that the isotropic constituents of cosmic hard X-ray have their origins outside the galactic system. As the spectra are uncertain, the generation mechanism of X-ray has not been clearly known yet. It was attempted to make more reliable observation by shutter method and the technique removing charged particles, using the L-3H-8 rocket. The equipment consists of NaI scintillation counter, a front counter, a Xenon counter, a UV sensor, a collimator, a shutter and a shutter-driving device. The L-3H-9 rocket was launched on August 16, 1977, and reached height of 310 km in about 300 seconds. Then the observation was started, but it was not able to observe the isotropic constituents of hard X-ray which were aimed at, as the shutter didn't work normally. It is expected to make another observation with the K-9M-64 rocket in August, 1978, after investigating the action of the shutter and employing and improved driving device. (Kobatake, H.)

Air liquefaction and enrichment system propulsion in reusable launch vehicles

Science.gov (United States)

Bond, W. H.; Yi, A. C.

1994-07-01

A concept is shown for a fully reusable, Earth-to-orbit launch vehicle with horizontal takeoff and landing, employing an air-turborocket for low speed and a rocket for high-speed acceleration, both using liquid hydrogen for fuel. The turborocket employs a modified liquid air cycle to supply the oxidizer. The rocket uses 90% pure liquid oxygen as its oxidizer that is collected from the atmosphere, separated, and stored during operation of the turborocket from about Mach 2 to 5 or 6. The takeoff weight and the thrust required at takeoff are markedly reduced by collecting the rocket oxidizer in-flight. This article shows an approach and the corresponding technology needs for using air liquefaction and enrichment system propulsion in a single-stage-to-orbit (SSTO) vehicle. Reducing the trajectory altitude at the end of collection reduces the wing area and increases payload. The use of state-of-the-art materials, such as graphite polyimide, in a direct substitution for aluminum or aluminum-lithium alloy, is critical to meet the structure weight objective for SSTO. Configurations that utilize 'waverider' aerodynamics show great promise to reduce the vehicle weight.

Experimental investigation of solid rocket motors for small sounding rockets

Science.gov (United States)

Suksila, Thada

2018-01-01

Experimentation and research of solid rocket motors are important subjects for aerospace engineering students. However, many institutes in Thailand rarely include experiments on solid rocket motors in research projects of aerospace engineering students, mainly because of the complexity of mixing the explosive propellants. This paper focuses on the design and construction of a solid rocket motor for total impulse in the class I-J that can be utilised as a small sounding rocket by researchers in the near future. Initially, the test stands intended for measuring the pressure in the combustion chamber and the thrust of the solid rocket motor were designed and constructed. The basic design of the propellant configuration was evaluated. Several formulas and ratios of solid propellants were compared for achieving the maximum thrust. The convenience of manufacturing and casting of the fabricated solid rocket motors were a critical consideration. The motor structural analysis such as the combustion chamber wall thickness was also discussed. Several types of nozzles were compared and evaluated for ensuring the maximum thrust of the solid rocket motors during the experiments. The theory of heat transfer analysis in the combustion chamber was discussed and compared with the experimental data.

Proposed UK high-latitude rocket campaign in late 1976/early 1977

International Nuclear Information System (INIS)

Thomas, G.R.; Bryant, D.A.

1975-01-01

The second major UK high-latitude rocket campaign is scheduled for late 1976/early 19777 at Andoya. The proposed experiments provide a comprehensive set of measurements of high-latitude phenomena and include studies of the sources and acceleration of auroral particles, the stability of plasma flow, wave-particle interactions, and the response of the atmosphere and ionosphere to enhanced geomagnetic activity. These experiments require co-ordinated launching of high-latitude (740-950 km) and small, medium-altitude (320-370km) rockets. The provisional campaign plan includes four Skylark 12's (with Skylark 11 as a possible substitute), one Skylark 7 (with Skylark 6 as a possible substitute), and five Fulmars (with Skylark 10A as a possible substitute). Some of the experiments require simultaneous measurements by GEOS in the European sector (early 1977), but the remainder could be carried out in late 1976

Performance and technological feasibility of rocket powered HTHL-SSTO with take-off assist (aerospace plane/ekranoplane)

Science.gov (United States)

Tomita, Nobuyuki; Nebylov, Alexander V.; Sokolov, Victor V.; Ohkami, Yoshiaki

It might be said that it is common understanding that rocket-powered single stage to orbit (SSTO) aerospace planes will become feasible with near-term technology as described in [1] (Koelle, D. E. Survey and comparison of winged launch vehicle options, ISTS 94-g-11 V 1994) and [2] (Bekey, I. Why SSTO rocket launch vehicles are now feasible and practical, IAF-94-V.1.524 1994). Among two methods of launching aerospace planes into orbit, vertical take-off (VT) and horizontal take-off (HT), it seems that VT takes the lead from HT [1, 2]. The decision for the X-33 program by NASA, also, seems to favor VT. In retrospect, almost all of the launch vehicles in the past have been VT, mainly because VT solved the problem of exit from atmosphere to space. However, broadening the range of requirements for space transportation systems from military to commercial and unmanned to manned seems to favor the need for HT. In this paper, the authors are going to prove that aerospace plane/ekranoplane system, which is a reusable launch vehicle system based on the HT concept, with ekranoplane as a take-off and possibly, landing assist, could be competitive with the VT concept from both technological and economical view points. Ekranoplane is a wing-in-ground-effect craft (WIG), which moves at a speed of approximately 0.5 M, carrying heavy loads above the sea surface. Combination of high initial velocity and high performance tri-propellant engine for aerospace plane makes it possible to configure an aerospace plane which is competitive with VT. Other specific features of HT in comparison with VT are discussed.

Evaluation of Alternative Refractory Materials for the Main Flame Deflectors at KSC Launch Complexes

Science.gov (United States)

Calle, Luz Marina; Trejo, David; Rutkowsky, Justin

2006-01-01

The deterioration of the refractory materials used to protect the KSC launch complex steel base structures from the high temperatures during launches results in frequent and costly repairs and safety hazards. KSC-SPEC-P-0012, Specification for Refractory Concrete, is ineffective in qualifying refractory materials. This study of the specification and of alternative refractory materials recommends a complete revision of the specification and further investigation of materials that were found to withstand the environment of the Solid Rocket Booster main flame deflector better than the refractory materials in current use in terms of compressive strength, tensile strength, modulus of rupture, shrinkage, and abrasion.

Dynamic modeling and ascent flight control of Ares-I Crew Launch Vehicle

Science.gov (United States)

Du, Wei

This research focuses on dynamic modeling and ascent flight control of large flexible launch vehicles such as the Ares-I Crew Launch Vehicle (CLV). A complete set of six-degrees-of-freedom dynamic models of the Ares-I, incorporating its propulsion, aerodynamics, guidance and control, and structural flexibility, is developed. NASA's Ares-I reference model and the SAVANT Simulink-based program are utilized to develop a Matlab-based simulation and linearization tool for an independent validation of the performance and stability of the ascent flight control system of large flexible launch vehicles. A linearized state-space model as well as a non-minimum-phase transfer function model (which is typical for flexible vehicles with non-collocated actuators and sensors) are validated for ascent flight control design and analysis. This research also investigates fundamental principles of flight control analysis and design for launch vehicles, in particular the classical "drift-minimum" and "load-minimum" control principles. It is shown that an additional feedback of angle-of-attack can significantly improve overall performance and stability, especially in the presence of unexpected large wind disturbances. For a typical "non-collocated actuator and sensor" control problem for large flexible launch vehicles, non-minimum-phase filtering of "unstably interacting" bending modes is also shown to be effective. The uncertainty model of a flexible launch vehicle is derived. The robust stability of an ascent flight control system design, which directly controls the inertial attitude-error quaternion and also employs the non-minimum-phase filters, is verified by the framework of structured singular value (mu) analysis. Furthermore, nonlinear coupled dynamic simulation results are presented for a reference model of the Ares-I CLV as another validation of the feasibility of the ascent flight control system design. Another important issue for a single main engine launch vehicle is

Determination of the Flow Field in the Propellant Tank of a Rocket Engine on Completion of the Mission

Science.gov (United States)

Fedorov, A. V.; Bedarev, I. A.; Lavruk, S. A.; Trushlyakov, V. I.; Kudentsov, V. Yu.

2018-03-01

In the present work, a method of mathematical simulation is employed to describe processes occurring in the specimens of new equipment and using the remaining propellant in rocket-engine tanks. Within the framework of certain turbulence models, the authors perform a calculation of the flow field in the volume of the tank of the launch-vehicle stage when a hot gas jet is injected into it. A vortex flow structure is revealed; the characteristics of heat transfer for different angles of injection of the jet are determined. The obtained correlation Nu = Nu(Re) satisfactorily describes experimental data.

The flight of uncontrolled rockets

CERN Document Server

Gantmakher, F R; Dryden, H L

1964-01-01

International Series of Monographs on Aeronautics and Astronautics, Division VII, Volume 5: The Flight of Uncontrolled Rockets focuses on external ballistics of uncontrolled rockets. The book first discusses the equations of motion of rockets. The rocket as a system of changing composition; application of solidification principle to rockets; rotational motion of rockets; and equations of motion of the center of mass of rockets are described. The text looks at the calculation of trajectory of rockets and the fundamentals of rocket dispersion. The selection further focuses on the dispersion of f

Modeling Rocket Flight in the Low-Friction Approximation

Directory of Open Access Journals (Sweden)

Logan White

2014-09-01

Full Text Available In a realistic model for rocket dynamics, in the presence of atmospheric drag and altitude-dependent gravity, the exact kinematic equation cannot be integrated in closed form; even when neglecting friction, the exact solution is a combination of elliptic functions of Jacobi type, which are not easy to use in a computational sense. This project provides a precise analysis of the various terms in the full equation (such as gravity, drag, and exhaust momentum, and the numerical ranges for which various approximations are accurate to within 1%. The analysis leads to optimal approximations expressed through elementary functions, which can be implemented for efficient flight prediction on simple computational devices, such as smartphone applications.

Ionospheric response to the shock and acoustic waves excited by the launch of the Shenzhou 10 spacecraft

Science.gov (United States)

Ding, Feng; Wan, Weixing; Mao, Tian; Wang, Min; Ning, Baiqi; Zhao, Biqiang; Xiong, Bo

2014-05-01

We used a dense GPS network in China to track the ionospheric response to waves excited by the launch of the rocket that carried Shenzhou 10 spacecraft on 11 June 2013. The long-distance propagation of shock and acoustic waves were observed on both sides of the rocket's trajectory. On the southern side, the wave structures (characterized by a horizontal extension of ~1400 km and initial amplitudes of 0.3 total electron content unit (TECU) and 0.1 TECU for the shock and acoustic waves, respectively), traveled southwestward a distance of ~1500 km at mean velocities of 1011 m s-1 and 709 m s-1, respectively. On the northern side, northward propagating waves were seen to travel a distance of ~600 km with much smaller amplitudes of less than 0.05 TECU. Subsequent waves with amplitudes of less than 0.02 TECU could also be seen. Clear wave structures were found at a distance of ~600-2000 km from launch site.

The Soyuz launch vehicle the two lives of an engineering triumph

CERN Document Server

Lardier, Christian

2013-01-01

The Soyuz launch vehicle has had a long and illustrious history. Built as the world's first intercontinental missile, it took the first man into space in April 1961, before becoming the workhorse of Russian spaceflight, launching satellites, interplanetary probes, every cosmonaut from Gagarin onwards, and, now, the multinational crews of the International Space Station. This remarkable book gives a complete and accurate description of the two lives of Soyuz, chronicling the cooperative space endeavor of Europe and Russia. First, it takes us back to the early days of astronautics, when technology served politics. From archives found in the Soviet Union the authors describe the difficulty of designing a rocket in the immediate post-war period. Then, in Soyuz's golden age, it launched numerous scientific missions and manned flights which were publicized worldwide while the many more numerous military missions were kept highly confidential! The second part of the book tells the contemporary story of the second li...

Benefits to the Europa Clipper Mission Provided by the Space Launch System

Science.gov (United States)

Creech, Stephen D.; Patel, Keyur

2013-01-01

The National Aeronautics and Space Administration's (NASA's) proposed Europa Clipper mission would provide an unprecedented look at the icy Jovian moon, and investigate its environment to determine the possibility that it hosts life. Focused on exploring the water, chemistry, and energy conditions on the moon, the spacecraft would examine Europa's ocean, ice shell, composition and geology by performing 32 low-altitude flybys of Europa from Jupiter orbit over 2.3 years, allowing detailed investigations of globally distributed regions of Europa. In hopes of expediting the scientific program, mission planners at NASA's Jet Propulsion Laboratory are working with the Space Launch System (SLS) program, managed at Marshall Space Flight Center. Designed to be the most powerful launch vehicle ever flown, SLS is making progress toward delivering a new capability for exploration beyond Earth orbit. The SLS rocket will offer an initial low-Earth-orbit lift capability of 70 metric tons (t) beginning with a first launch in 2017 and will then evolve into a 130 t Block 2 version. While the primary focus of the development of the initial version of SLS is on enabling human exploration missions beyond low Earth orbit using the Orion Multi-Purpose Crew Vehicle, the rocket offers unique benefits to robotic planetary exploration missions, thanks to the high characteristic energy it provides. This paper will provide an overview of both the proposed Europa Clipper mission and the Space Launch System vehicle, and explore options provided to the Europa Clipper mission for a launch within a decade by a 70 t version of SLS with a commercially available 5-meter payload fairing, through comparison with a baseline of current Evolved Expendable Launch Vehicle (EELV) capabilities. Compared to that baseline, a mission to the Jovian system could reduce transit times to less than half, or increase mass to more than double, among other benefits. In addition to these primary benefits, the paper will

Nuclear rockets

Energy Technology Data Exchange (ETDEWEB)

Sarram, M [Teheran Univ. (Iran). Inst. of Nuclear Science and Technology

1972-02-01

Nuclear energy has found many applications in space projects. This article deals with these applications. The first application is the use of nuclear energy for the production of electricity in space and the second main application is the use of nuclear energy for propulsion purposes in space flight. The main objective is to develop a 75000 pound thrust flight engine called NERVA by heating liquid hydrogen in a nuclear reactor. The paper describes in detail the salient features of the NERVA rocket as well as its comparison with the conventional chemical rockets. It is shown that a nuclear rocket using liquid hydrogen as medium is at least 85% more efficient as compared with the chemical rockets such as those used for the APOLLO moon flight.

On transient electric fields observed in chemical release experiments by rockets

International Nuclear Information System (INIS)

Marklund, G.; Brenning, N.; Holmgren, G.; Haerendel, G.

1986-06-01

As a follow-up to the successful chemical release experiment Trigger in 1977, the TOR (Trigger Optimized Repetition) rocket was launched from Esrange on Oct. 24, 1984. Like in the Trigger experiment a large amplitude electric field pulse of 200 mV/m was detected shortly after the explosion. The central part of the pulse was found to be clearly correlated with an intense layer of swept up ambient particles behind a propagating shock-front. The field was directed towards the centre of the expanding ionized cloud, which is indicative of a polarisation electric field source. Expressions for this radial polarisation field and the much weaker azimuthal induced electric field are derived from a simple cylindrical model for the field and the expanding neutral cloud. Time profiles of the radial electric field are shown to be in good agreement with observations. (authors)

Flight demonstration of flight termination system and solid rocket motor ignition using semiconductor laser initiated ordnance

Science.gov (United States)

Schulze, Norman R.; Maxfield, B.; Boucher, C.

1995-01-01

Solid State Laser Initiated Ordnance (LIO) offers new technology having potential for enhanced safety, reduced costs, and improved operational efficiency. Concerns over the absence of programmatic applications of the technology, which has prevented acceptance by flight programs, should be abated since LIO has now been operationally implemented by the Laser Initiated Ordnance Sounding Rocket Demonstration (LOSRD) Program. The first launch of solid state laser diode LIO at the NASA Wallops Flight Facility (WFF) occurred on March 15, 1995 with all mission objectives accomplished. This project, Phase 3 of a series of three NASA Headquarters LIO demonstration initiatives, accomplished its objective by the flight of a dedicated, all-LIO sounding rocket mission using a two-stage Nike-Orion launch vehicle. LIO flight hardware, made by The Ensign-Bickford Company under NASA's first Cooperative Agreement with Profit Making Organizations, safely initiated three demanding pyrotechnic sequence events, namely, solid rocket motor ignition from the ground and in flight, and flight termination, i.e., as a Flight Termination System (FTS). A flight LIO system was designed, built, tested, and flown to support the objectives of quickly and inexpensively putting LIO through ground and flight operational paces. The hardware was fully qualified for this mission, including component testing as well as a full-scale system test. The launch accomplished all mission objectives in less than 11 months from proposal receipt. This paper concentrates on accomplishments of the ordnance aspects of the program and on the program's implementation and results. While this program does not generically qualify LIO for all applications, it demonstrated the safety, technical, and operational feasibility of those two most demanding applications, using an all solid state safe and arm system in critical flight applications.

Rocket measurements of electric fields, electron density and temperature during the three phases of auroral substorms

International Nuclear Information System (INIS)

Marklund, G.; Block, L.; Lindqvist, P.-A.

1979-12-01

On Jan. 27, 1979, three rocket payloads were launched from Kiruna, Sweden, into different phases of two successive auroral substorms. Among other experiments, the payloads carried the RIT double probe electric field experiments, providing electric field, electron density and temperature data, which are presented here. These are discussed in association with observations of particles, ionospheric drifts (STARE) and electric fields in the equatorial plane (GEOS). The motions of the auroral forms, as obtained from auroral pictures are compared with the E x B/B 2 drifts and the currents calculated from the rocket electric field and density measurements with the equivalent current system deduced from ground based magnetometer data (SMA). (Auth.)

A computational model for thermal fluid design analysis of nuclear thermal rockets

International Nuclear Information System (INIS)

Given, J.A.; Anghaie, S.

1997-01-01

A computational model for simulation and design analysis of nuclear thermal propulsion systems has been developed. The model simulates a full-topping expander cycle engine system and the thermofluid dynamics of the core coolant flow, accounting for the real gas properties of the hydrogen propellant/coolant throughout the system. Core thermofluid studies reveal that near-wall heat transfer models currently available may not be applicable to conditions encountered within some nuclear rocket cores. Additionally, the possibility of a core thermal fluid instability at low mass fluxes and the effects of the core power distribution are investigated. Results indicate that for tubular core coolant channels, thermal fluid instability is not an issue within the possible range of operating conditions in these systems. Findings also show the advantages of having a nonflat centrally peaking axial core power profile from a fluid dynamic standpoint. The effects of rocket operating conditions on system performance are also investigated. Results show that high temperature and low pressure operation is limited by core structural considerations, while low temperature and high pressure operation is limited by system performance constraints. The utility of these programs for finding these operational limits, optimum operating conditions, and thermal fluid effects is demonstrated

Nuclear thermal rockets using indigenous extraterrestrial propellants

International Nuclear Information System (INIS)

Zubrin, R.M.

1990-01-01

A preliminary examination of a concept for a Mars and outer solar system exploratory vehicle is presented. Propulsion is provided by utilizing a nuclear thermal reactor to heat a propellant volatile indigenous to the destination world to form a high thrust rocket exhaust. Candidate propellants, whose performance, materials compatibility, and ease of acquisition are examined and include carbon dioxide, water, methane, nitrogen, carbon monoxide, and argon. Ballistics and winged supersonic configurations are discussed. It is shown that the use of this method of propulsion potentially offers high payoff to a manned Mars mission. This is accomplished by sharply reducing the initial mission mass required in low earth orbit, and by providing Mars explorers with greatly enhanced mobility in traveling about the planet through the use of a vehicle that can refuel itself each time it lands. Thus, the nuclear landing craft is utilized in combination with a hydrogen-fueled nuclear-thermal interplanetary launch. By utilizing such a system in the outer solar system, a low level aerial reconnaissance of Titan combined with a multiple sample return from nearly every satellite of Saturn can be accomplished in a single launch of a Titan 4 or the Space Transportation System (STS). Similarly a multiple sample return from Callisto, Ganymede, and Europa can also be accomplished in one launch of a Titan 4 or the STS

Using Innovative Technologies for Manufacturing Rocket Engine Hardware

Science.gov (United States)

Betts, E. M.; Eddleman, D. E.; Reynolds, D. C.; Hardin, N. A.

2011-01-01

Many of the manufacturing techniques that are currently used for rocket engine component production are traditional methods that have been proven through years of experience and historical precedence. As the United States enters into the next space age where new launch vehicles are being designed and propulsion systems are being improved upon, it is sometimes necessary to adopt innovative techniques for manufacturing hardware. With a heavy emphasis on cost reduction and improvements in manufacturing time, rapid manufacturing techniques such as Direct Metal Laser Sintering (DMLS) are being adopted and evaluated for their use on NASA s Space Launch System (SLS) upper stage engine, J-2X, with hopes of employing this technology on a wide variety of future projects. DMLS has the potential to significantly reduce the processing time and cost of engine hardware, while achieving desirable material properties by using a layered powder metal manufacturing process in order to produce complex part geometries. Marshall Space Flight Center (MSFC) has recently hot-fire tested a J-2X gas generator (GG) discharge duct that was manufactured using DMLS. The duct was inspected and proof tested prior to the hot-fire test. Using a workhorse gas generator (WHGG) test fixture at MSFC's East Test Area, the duct was subjected to extreme J-2X hot gas environments during 7 tests for a total of 537 seconds of hot-fire time. The duct underwent extensive post-test evaluation and showed no signs of degradation. DMLS manufacturing has proven to be a viable option for manufacturing rocket engine hardware, and further development and use of this manufacturing method is recommended.

Observation of galactic soft x-ray by the rocket K-10-14

International Nuclear Information System (INIS)

Sato, Yasue; Higuchi, Masuto; Kitamoto, Shunji; Miyamoto, Shigenori; Kawabata, Akira.

1981-01-01

An experiment was performed to observe the soft X-ray from Cygnus Loop. Since the attitude control system of the rocket did not work well, the observed area was around Hercules HI min. The observation was made with the HTXT on K-10-14. The local hot plasma in this region was able to be evaluated. The spectrum analysis was made on the data (region 1) taken from 114 to 219 second after the launch of the rocket and the data (region 2) taken from 226 to 300 second. In these two periods, the direction of field of view was stable. The counting rates of M band (more than 0.5 KeV) and L band (less than 0.3 KeV) are shown in figures. The counting rate in the region 2 was stronger than that of the region 1. By Hadamard transformation, images were constructed from the obtained data. The intensity of the diffuse component of soft X-ray was isotropic. This result proved the correct operation of HTXT. The chi-square fit of the energy spectrum in the region 2 showed a good agreement with a high temperature plasma model. In the region 1, the simple high temperature plasma model was able to be rejected with 90 percent confidence. It can be said that the Hercules HI min region is filled with local hot plasma, and this reaches to the end of direction of thickness of galaxy disk. (Kato, T.)

Performance and technical feasibility comparison of reusable launch systems: A synthesis of the ESA winged launcher studies

Science.gov (United States)

Berry, W.; Grallert, H.

1996-02-01

The paper presents a synthesis of the performance and technical feasibility assessment of 7 reusable launcher types, comprising 13 different vehicles, studied by European Industry for ESA in the ESA Winged Launcher Study in the period January 1988 to May 1994. The vehicles comprised single-stage-to-orbit (SSTO) and two-stage-to-orbit (TSTO) vehicles, propelled by either air-breathing/rocket propulsion or entirely by rocket propulsion. The results showed that an SSTO vehicle of the HOTOL-type, propelled by subsonic combustion air-breathing/rocket engines could barely deliver the specified payload mass and was aerodynamically unstable; that a TSTO vehicle of the Saenger type, employing subsonic combustion airbreathing propulsion in its first stage and rocket propulsion in its second stage, could readily deliver the specified payload mass and was found to be technically feasible and versatile; that an SSTO vehicle of the NASP type, propelled by supersonic combustion airbreathing/rocket propulsion was able to deliver a reduced payload mass, was very complex and required very advanced technologies; that an air-launched rocket propelled vehicle of the Interim HOTOL type, although technically feasible, could deliver only a reduced payload mass, being constrained by the lifting capability of the carrier airplane; that three different, entirely rocket-propelled vehicles could deliver the specified payload mass, were technically feasible but required relatively advanced technologies.

Nuclear rockets

International Nuclear Information System (INIS)

Sarram, M.

1972-01-01

Nuclear energy has found many applications in space projects. This article deals with these applications. The first application is the use of nuclear energy for the production of electricity in space and the second main application is the use of nuclear energy for propulsion purposes in space flight. The main objective is to develop a 75000 pound thrust flight engine call NERVA by heating liquid hydrogen, in a nuclear reactor, from 420F to 4000 0 F. The paper describes in detail the salient features of the NERVA rocket as well as its comparison with the conventional chemical rockets. It is shown that a nuclear rocket using liquid hydrogen as medium is at least 85% more efficient as compared with the chemical rockets such as those used for the APOLLO moon flight

Experimental reslts from the HERO project: In situ measurements of ionospheric modifications using sounding rockets

International Nuclear Information System (INIS)

Rose, G.; Grandal, B.; Neske, E.; Ott, W.; Spenner, K.; Maseide, K.; Troim, J.

1985-01-01

The Heating Rocket project HERO comprised the first in situ experiments to measure artifical ionospheric modifications at F layer heights set up by radio waves transmitted from the Heating facility at Ramfjord near Tromso in Northern Norway. Four instrumented payloads were launched on sounding rockets from Andoya Rocket Range during the autumn of 1982 into a sunlit ionosphere with the sun close to the horizon. The payloads recorded modifications, in particular, the presence of electron plasma waves near the reflection level of the heating wave. The amplitude and phase of the three components of the electric and magnetic fields of the heating wave were measured simultaneously as a function of altitude. Coherent spectra of the three electric field components of the locally generated electron plasma waves were obtained in a 50-kHz-wide band. At the same time quasi-continuous measurements were made on several fixed frequencies from 4 kHz to 16 kHz below the heating frequency and in the VLF-range using linear dipole antennas. Moreover, measurements were made of electron temperature, suprathermal electrons and local electron density along the rocket trajectory. The experimental results will be presented and discussed

Echo 2: a study of electron beams injected into the high-latitude ionosphere from a large sounding rocket

International Nuclear Information System (INIS)

Winckler, J.R.; Arnoldy, R.L.; Hendrickson, R.A.

1975-01-01

The Black Brant V-C Echo 2 rocket was launched at Fort Churchill on September 25, 1972, and it injected 64-ms pulses of electron beams of 80-mA current and 45-keV voltage into the ionosphere. This paper studies the responses of on-board electrostatic deflection and solid state detectors to injected electrons after motion in the near ionosphere and atmosphere. It is shown that it was only through some form of scattering that the detectors could sense the injected beam electrons. By means of 'phase maps' of injection and detection pitch angles a number of distinct regions are found corresponding to a rocket scattering halo, an atmospheric scattering halo, a region of weak responses, and a source of strong scattering above the rocket. The atmospheric scattering has been compared with the theoretical and experimental results of the Echo 1 experiment, and it is found to be in reasonable agreement. The rocket halo is discussed qualitatively; but no explanation is found for the backscatter from above the rocket, which may be associated with an occasional violent beam instability. This analysis has been carried out to better understand the complexities of electron motion observed near large rockets carrying artifical electron accelerators as a guide in the planning of future experiments

The microwave thermal thruster and its application to the launch problem

Science.gov (United States)

Parkin, Kevin L. G.

Nuclear thermal thrusters long ago bypassed the 50-year-old specific impulse (Isp) limitation of conventional thrusters, using nuclear powered heat exchangers in place of conventional combustion to heat a hydrogen propellant. These heat exchanger thrusters experimentally achieved an Isp of 825 seconds, but with a thrust-to-weight ratio (T/W) of less than ten they have thus far been too heavy to propel rockets into orbit. This thesis proposes a new idea to achieve both high Isp and high T/W The Microwave Thermal Thruster. This thruster covers the underside of a rocket aeroshell with a lightweight microwave absorbent heat exchange layer that may double as a re-entry heat shield. By illuminating the layer with microwaves directed from a ground-based phased array, an Isp of 700--900 seconds and T/W of 50--150 is possible using a hydrogen propellant. The single propellant simplifies vehicle design, and the high Isp increases payload fraction and structural margins. These factors combined could have a profound effect on the economics of building and reusing rockets. A laboratory-scale microwave thermal heat exchanger is constructed using a single channel in a cylindrical microwave resonant cavity, and new type of coupled electromagnetic-conduction-convection model is developed to simulate it. The resonant cavity approach to small-scale testing reveals several drawbacks, including an unexpected oscillatory behavior. Stable operation of the laboratory-scale thruster is nevertheless successful, and the simulations are consistent with the experimental results. In addition to proposing a new type of propulsion and demonstrating it, this thesis provides three other principal contributions: The first is a new perspective on the launch problem, placing it in a wider economic context. The second is a new type of ascent trajectory that significantly reduces the diameter, and hence cost, of the ground-based phased array. The third is an eclectic collection of data, techniques, and

Sounding rocket flight report, MUMP 9 and MUMP 10

Science.gov (United States)

Grassl, H. J.

1971-01-01

The results of the launching of two-Marshall-University of Michigan Probes (MUMP 9 and MUMP 10), Nike-Tomahawk sounding rocket payloads, are summarized. The MUMP is similar to the thermosphere probe, an ejectable instrument package for studying the variability of the earth's atmospheric parameters. The MUMP 9 payload included an omegatron mass analyzer, a molecular fluorescence densitometer, a mini-tilty filter, and a lunar position sensor. This complement of instruments permitted the determination of the molecular nitrogen density and temperature in the altitude range from approximately 143 to 297 km over Wallops Island, Virginia, during January 1971. The MUMP 10 payload included an omegatron mass analyzer, an electron temperature probe, a cryogenic densitometer, and a solar position sensor. These instruments permitted the determination of the molecular nitrogen density and temperature and the charged particle density and temperature in the altitude range from approximately 145 to 290 km over Wallops Island during the afternoon preceding the MUMP 9 launch.

A Change of Inertia-Supporting the Thrust Vector Control of the Space Launch System

Science.gov (United States)

Dziubanek, Adam J.

2012-01-01

The Space Launch System (SLS) is America's next launch vehicle. To utilize the vehicle more economically, heritage hardware from the Space Transportation System (STS) will be used when possible. The Solid Rocket Booster (SRB) actuators could possibly be used in the core stage of the SLS. The dynamic characteristics of the SRB actuator will need to be tested on an Inertia Load Stand (ILS) that has been converted to Space Shuttle Main Engine (SSME). The inertia on the pendulum of the ILS will need to be changed to match the SSME inertia. In this testing environment an SRB actuator can be tested with the equivalent resistence of an SSME.

Optical tools and techniques for aligning solar payloads with the SPARCS control system. [Solar Pointing Aerobee Rocket Control System

Science.gov (United States)

Thomas, N. L.; Chisel, D. M.

1976-01-01

The success of a rocket-borne experiment depends not only on the pointing of the attitude control system, but on the alignment of the attitude control system to the payload. To ensure proper alignment, special optical tools and alignment techniques are required. Those that were used in the SPARCS program are described and discussed herein. These tools include theodolites, autocollimators, a 38-cm diameter solar simulator, a high-performance 1-m heliostat to provide a stable solar source during the integration of the rocket payload, a portable 75-cm sun tracker for use at the launch site, and an innovation called the Solar Alignment Prism. Using the real sun as the primary reference under field conditions, the Solar Alignment Prism facilitates the coalignment of the attitude sun sensor with the payload. The alignment techniques were developed to ensure the precise alignment of the solar payloads to the SPARCS attitude sensors during payload integration and to verify the required alignment under field conditions just prior to launch.

Modeling of Uneven Flow and Electromagnetic Field Parameters in the Combustion Chamber of Liquid Rocket Engine with a Near-wall Layer Available

Directory of Open Access Journals (Sweden)

A. V. Rudinskii

2015-01-01

Full Text Available The paper concerns modeling of an uneven flow and electromagnetic field parameters in the combustion chamber of the liquid rocket engine with a near-wall layer available.The research objective was to evaluate quantitatively influence of changing model chamber mode of the liquid rocket engine on the electro-physical characteristics of the hydrocarbon fuel combustion by-products.The main method of research was based on development of a final element model of the flowing path of the rocket engine chamber and its adaptation to the boundary conditions.The paper presents a developed two-dimensional non-stationary mathematical model of electro-physical processes in the liquid rocket engine chamber using hydrocarbon fuel. The model takes into consideration the features of a gas-dynamic contour of the engine chamber and property of thermo-gas-dynamic characteristics of the ionized products of combustion of hydrocarbonic fuel. Distributions of magnetic field intensity and electric conductivity received and analyzed taking into account a low-temperature near-wall layer. Special attention is paid to comparison of obtained calculation values of the electric current, which is taken out from intrachamber space of the engine with earlier published data of other authors.

Reusable launch vehicles, enabling technology for the development of advanced upper stages and payloads

International Nuclear Information System (INIS)

Metzger, John D.

1998-01-01

In the near future there will be classes of upper stages and payloads that will require initial operation at a high-earth orbit to reduce the probability of an inadvertent reentry that could result in a detrimental impact on humans and the biosphere. A nuclear propulsion system, such as was being developed under the Space Nuclear Thermal Propulsion (SNTP) Program, is an example of such a potential payload. This paper uses the results of a reusable launch vehicle (RLV) study to demonstrate the potential importance of a Reusable Launch Vehicle (RLV) to test and implement an advanced upper stage (AUS) or payload in a safe orbit and in a cost effective and reliable manner. The RLV is a horizontal takeoff and horizontal landing (HTHL), two-stage-to-orbit (TSTO) vehicle. The results of the study shows that an HTHL is cost effective because it implements airplane-like operation, infrastructure, and flight operations. The first stage of the TSTO is powered by Rocket-Based-Combined-Cycle (RBCC) engines, the second stage is powered by a LOX/LH rocket engine. The TSTO is used since it most effectively utilizes the capability of the RBCC engine. The analysis uses the NASA code POST (Program to Optimize Simulated Trajectories) to determine trajectories and weight in high-earth orbit for AUS/advanced payloads. Cost and reliability of an RLV versus current generation expandable launch vehicles are presented

A Shuttle Derived Vehicle launch system

Science.gov (United States)

Tewell, J. R.; Buell, D. N.; Ewing, E. S.

1982-01-01

This paper describes a Shuttle Derived Vehicle (SDV) launch system presently being studied for the NASA by Martin Marietta Aerospace which capitalizes on existing Shuttle hardware elements to provide increased accommodations for payload weight, payload volume, or both. The SDV configuration utilizes the existing solid rocket boosters, external tank and the Space Shuttle main engines but replaces the manned orbiter with an unmanned, remotely controlled cargo carrier. This cargo carrier substitution more than doubles the performance capability of the orbiter system and is realistically achievable for minimal cost. The advantages of the SDV are presented in terms of performance and economics. Based on these considerations, it is concluded that an unmanned SDV offers a most attractive complement to the present Space Transportation System.

The Dual-channel Extreme Ultraviolet Continuum Experiment: Sounding Rocket EUV Observations of Local B Stars to Determine Their Potential for Supplying Intergalactic Ionizing Radiation

Science.gov (United States)

Erickson, Nicholas; Green, James C.; France, Kevin; Stocke, John T.; Nell, Nicholas

2018-06-01

We describe the scientific motivation and technical development of the Dual-channel Extreme Ultraviolet Continuum Experiment (DEUCE). DEUCE is a sounding rocket payload designed to obtain the first flux-calibrated spectra of two nearby B stars in the EUV 650-1150Å bandpass. This measurement will help in understanding the ionizing flux output of hot B stars, calibrating stellar models and commenting on the potential contribution of such stars to reionization. DEUCE consists of a grazing incidence Wolter II telescope, a normal incidence holographic grating, and the largest (8” x 8”) microchannel plate detector ever flown in space, covering the 650-1150Å band in medium and low resolution channels. DEUCE will launch on December 1, 2018 as NASA/CU sounding rocket mission 36.331 UG, observing Epsilon Canis Majoris, a B2 II star.

Yes--This is Rocket Science: MMCs for Liquid Rocket Engines

National Research Council Canada - National Science Library

Shelley, J

2001-01-01

The Air Force's Integrated High-Payoff Rocket Propulsion Technologies (IHPRPT) Program has established aggressive goals for both improved performance and reduced cost of rocket engines and components...

Peer Review of Launch Environments

Science.gov (United States)

Wilson, Timmy R.

2011-01-01

Catastrophic failures of launch vehicles during launch and ascent are currently modeled using equivalent trinitrotoluene (TNT) estimates. This approach tends to over-predict the blast effect with subsequent impact to launch vehicle and crew escape requirements. Bangham Engineering, located in Huntsville, Alabama, assembled a less-conservative model based on historical failure and test data coupled with physical models and estimates. This white paper summarizes NESC's peer review of the Bangham analytical work completed to date.

Two-Dimensional Motions of Rockets

Science.gov (United States)

Kang, Yoonhwan; Bae, Saebyok

2007-01-01

We analyse the two-dimensional motions of the rockets for various types of rocket thrusts, the air friction and the gravitation by using a suitable representation of the rocket equation and the numerical calculation. The slope shapes of the rocket trajectories are discussed for the three types of rocket engines. Unlike the projectile motions, the…

Rocket propulsion elements - An introduction to the engineering of rockets (6th revised and enlarged edition)

Science.gov (United States)

Sutton, George P.

The subject of rocket propulsion is treated with emphasis on the basic technology, performance, and design rationale. Attention is given to definitions and fundamentals, nozzle theory and thermodynamic relations, heat transfer, flight performance, chemical rocket propellant performance analysis, and liquid propellant rocket engine fundamentals. The discussion also covers solid propellant rocket fundamentals, hybrid propellant rockets, thrust vector control, selection of rocket propulsion systems, electric propulsion, and rocket testing.

Experimental Performance Evaluation of a Supersonic Turbine for Rocket Engine Applications

Science.gov (United States)

Snellgrove, Lauren M.; Griffin, Lisa W.; Sieja, James P.; Huber, Frank W.

2003-01-01

In order to mitigate the risk of rocket propulsion development, efficient, accurate, detailed fluid dynamics analysis and testing of the turbomachinery is necessary. To support this requirement, a task was developed at NASA Marshall Space Flight Center (MSFC) to improve turbine aerodynamic performance through the application of advanced design and analysis tools. These tools were applied to optimize a supersonic turbine design suitable for a reusable launch vehicle (RLV). The hot gas path and blading were redesigned-to obtain an increased efficiency. The goal of the demonstration was to increase the total-to- static efficiency of the turbine by eight points over the baseline design. A sub-scale, cold flow test article modeling the final optimized turbine was designed, manufactured, and tested in air at MSFC s Turbine Airflow Facility. Extensive on- and off- design point performance data, steady-state data, and unsteady blade loading data were collected during testing.

Rocket Ozone Data Recovery for Digital Archival

Science.gov (United States)

Hwang, S. H.; Krueger, A. J.; Hilsenrath, E.; Haffner, D. P.; Bhartia, P. K.

2014-12-01

Ozone distributions in the photochemically-controlled upper stratosphere and mesosphere were first measured using spectrometers on V-2 rockets after WWII. The IGY(1957-1958) spurred development of new optical and chemical instruments for flight on meteorological and sounding rockets. In the early 1960's, the US Navy developed an Arcas rocket-borne optical ozonesonde and NASA GSFC developed chemiluminescent ozonesonde onboard Nike_Cajun and Arcas rocket. The Navy optical ozone program was moved in 1969 to GSFC where rocket ozone research was expanded and continued until 1994 using Super Loki-Dart rocket at 11 sites in the range of 0-65N and 35W-160W. Over 300 optical ozone soundings and 40 chemiluminescent soundings were made. The data have been used to produce the US Standard Ozone Atmosphere, determine seasonal and diurnal variations, and validate early photochemical models. The current effort includes soundings conducted by Australia, Japan, and Korea using optical techniques. New satellite ozone sounding techniques were initially calibrated and later validated using the rocket ozone data. As satellite techniques superseded the rocket methods, the sponsoring agencies lost interest in the data and many of those records have been discarded. The current task intends to recover as much of the data as possible from the private records of the experimenters and their publications, and to archive those records in the WOUDC (World Ozone and Ultraviolet Data Centre). The original data records are handwritten tabulations, computer printouts that are scanned with OCR techniques, and plots digitized from publications. This newly recovered digital rocket ozone profile data from 1965 to 2002 could make significant contributions to the Earth science community in atmospheric research including long-term trend analysis.

JANNAF "Test and Evaluation Guidelines for Liquid Rocket Engines": Status and Application

Science.gov (United States)

Parkinson, Douglas; VanLerberghe, Wayne M.; Rahman, Shamim A.

2017-01-01

For many decades, the U.S. rocket propulsion industrial base has performed remarkably in developing complex liquid rocket engines that can propel critical payloads into service for the nation, as well as transport people and hardware for missions that open the frontiers of space exploration for humanity. This has been possible only at considerable expense given the lack of detailed guidance that captures the essence of successful practices and knowledge accumulated over five decades of liquid rocket engine development. In an effort to provide benchmarks and guidance for the next generation of rocket engineers, the Joint Army Navy NASA Air Force (JANNAF) Interagency Propulsion Committee published a liquid rocket engine (LRE) test and evaluation (T&E) guideline document in 2012 focusing on the development challenges and test verification considerations for liquid rocket engine systems. This document has been well received and applied by many current LRE developers as a benchmark and guidance tool, both for government-driven applications as well as for fully commercial ventures. The USAF Space and Missile Systems Center (SMC) has taken an additional near-term step and is directing activity to adapt and augment the content from the JANNAF LRE T&E guideline into a standard for potential application to future USAF requests for proposals for LRE development initiatives and launch vehicles for national security missions. A draft of this standard was already sent out for review and comment, and is intended to be formally approved and released towards the end of 2017. The acceptance and use of the LRE T&E guideline is possible through broad government and industry participation in the JANNAF liquid propulsion committee and associated panels. The sponsoring JANNAF community is expanding upon this initial baseline version and delving into further critical development aspects of liquid rocket propulsion testing at the integrated stage level as well as engine component level, in

NASA Space Rocket Logistics Challenges

Science.gov (United States)

Neeley, James R.; Jones, James V.; Watson, Michael D.; Bramon, Christopher J.; Inman, Sharon K.; Tuttle, Loraine

2014-01-01

The Space Launch System (SLS) is the new NASA heavy lift launch vehicle and is scheduled for its first mission in 2017. The goal of the first mission, which will be uncrewed, is to demonstrate the integrated system performance of the SLS rocket and spacecraft before a crewed flight in 2021. SLS has many of the same logistics challenges as any other large scale program. Common logistics concerns for SLS include integration of discreet programs geographically separated, multiple prime contractors with distinct and different goals, schedule pressures and funding constraints. However, SLS also faces unique challenges. The new program is a confluence of new hardware and heritage, with heritage hardware constituting seventy-five percent of the program. This unique approach to design makes logistics concerns such as commonality especially problematic. Additionally, a very low manifest rate of one flight every four years makes logistics comparatively expensive. That, along with the SLS architecture being developed using a block upgrade evolutionary approach, exacerbates long-range planning for supportability considerations. These common and unique logistics challenges must be clearly identified and tackled to allow SLS to have a successful program. This paper will address the common and unique challenges facing the SLS programs, along with the analysis and decisions the NASA Logistics engineers are making to mitigate the threats posed by each.

Simultaneous rocket and radar measurements of currents in an auroral arc

International Nuclear Information System (INIS)

Robinson, R.M.; Bering, E.A.; Vondrak, R.R.; Anderson, H.R.; Cloutier, P.A.

1981-01-01

A detailed study of electric field, current and conductivities associated with an auroral arc was made in a coordinated rocket and radar experiment in Alaska on March 9, 1978. The payload, designated 29.007 UE, was launched at 1013 p.m. local time. It penetrated the diffuse aurora on the upleg and at apogee traversed field lines connected to a stable auroral arc of 40 kR intensity. Among the instruments carried by the payload were a vector magnetometer, a set of electrostatic double probes and a set of electron and proton spectrometers. Simultaneous electron density and line-of-sight velocity measurements were made by Chatanika radar operating in an elevation scan mode in the magnetic meridian plane. Both the radar and rocket measurements indicated that the zonal electric field was westward and approximately constant across the arc with a magnitude of about 7 mV/m. Small differences between the rocket and radar zonal electric field measurements indicated the presence of upward drifting ions in the region of the arc. The meridional field was large and northward equatorward of the arc, but negligible within the arc. Conductivities computed from measured fluxes of energetic electrons agreed well with the conductivities derived from the radar measureements of electron density. The electric field and conductivity measurements indicated that the zonal currents were eastward equatorward of the arc and westward within the arc. These electrojet currents agreed well with those inferred from the rocket magnetometer data. Better agreement was obtained when a westward neutral wind was added. The westward wind was also consistent with differences between the rocket and radar meridional electric fields. The meridional currents computed from the electric field measurements were northward over the entire region

Advanced transportation system study: Manned launch vehicle concepts for two way transportation system payloads to LEO

Science.gov (United States)

Duffy, James B.

1993-01-01

The purpose of the Advanced Transportation System Study (ATSS) task area 1 study effort is to examine manned launch vehicle booster concepts and two-way cargo transfer and return vehicle concepts to determine which of the many proposed concepts best meets NASA's needs for two-way transportation to low earth orbit. The study identified specific configurations of the normally unmanned, expendable launch vehicles (such as the National Launch System family) necessary to fly manned payloads. These launch vehicle configurations were then analyzed to determine the integrated booster/spacecraft performance, operations, reliability, and cost characteristics for the payload delivery and return mission. Design impacts to the expendable launch vehicles which would be required to perform the manned payload delivery mission were also identified. These impacts included the implications of applying NASA's man-rating requirements, as well as any mission or payload unique impacts. The booster concepts evaluated included the National Launch System (NLS) family of expendable vehicles and several variations of the NLS reference configurations to deliver larger manned payload concepts (such as the crew logistics vehicle (CLV) proposed by NASA JSC). Advanced, clean sheet concepts such as an F-1A engine derived liquid rocket booster (LRB), the single stage to orbit rocket, and a NASP-derived aerospace plane were also included in the study effort. Existing expendable launch vehicles such as the Titan 4, Ariane 5, Energia, and Proton were also examined. Although several manned payload concepts were considered in the analyses, the reference manned payload was the NASA Langley Research Center's HL-20 version of the personnel launch system (PLS). A scaled up version of the PLS for combined crew/cargo delivery capability, the HL-42 configuration, was also included in the analyses of cargo transfer and return vehicle (CTRV) booster concepts. In addition to strictly manned payloads, two-way cargo

ROCOZ-A (improved rocket launched ozone sensor) for middle atmosphere ozone measurements

International Nuclear Information System (INIS)

Lee, H.S.; Parsons, C.L.

1987-01-01

An improved interference filter based ultraviolet photometer (ROCOZ-A) for measuring stratospheric ozone is discussed. The payload is launched aboard a Super-Loki to a typical apogee of 70 km. The instrument measures the solar ultraviolet irradiance as it descends on a parachute. The total cumulative ozone is then calculated based on the Beer-Lambert law. The cumulative ozone precision measured in this way is 2.0% to 2.5% over an altitude range of 20 and 55 km. Results of the intercomparison with the SBUV overpass data and ROCOZ-A data are also discussed

A study of performance and cost improvement potential of the 120 inch (3.05 m) diameter solid rocket motor. Volume 1: Summary report

Science.gov (United States)

Backlund, S. J.; Rossen, J. N.

1971-01-01

A parametric study of ballistic modifications to the 120 inch diameter solid propellant rocket engine which forms part of the Air Force Titan 3 system is presented. 576 separate designs were defined and 24 were selected for detailed analysis. Detailed design descriptions, ballistic performance, and mass property data were prepared for each design. It was determined that a relatively simple change in design parameters could provide a wide range of solid propellant rocket engine ballistic characteristics for future launch vehicle applications.

Lunar mission design using nuclear thermal rockets

International Nuclear Information System (INIS)

Stancati, M.L.; Collins, J.T.; Borowski, S.K.

1991-01-01

The NERVA-class Nuclear Thermal Rocket (NTR), with performance nearly double that of advanced chemical engines, has long been considered an enabling technology for human missions to Mars. NTR engines address the demanding trip time and payload delivery needs of both cargo-only and piloted flights. But NTR can also reduce the Earth launch requirements for manned lunar missions. First use of NTR for the Moon would be less demanding and would provide a test-bed for early operations experience with this powerful technology. Study of application and design options indicates that NTR propulsion can be integrated with the Space Exploration Initiative scenarios to deliver performance gains while managing controlled, long-term disposal of spent reactors to highly stable orbits

Sporadic sodium and E layers observed during the summer 2002 MaCWAVE/MIDAS rocket campaign

Directory of Open Access Journals (Sweden)

B. P. Williams

2006-07-01

Full Text Available On 5 July 2002, a MaCWAVE (Mountain and Convective Waves Ascending VErtically payload launched from Andøya Rocket Range, Norway, observed narrow enhanced layers of electron density that were nearly coincident with sporadic sodium layers measured by the Weber sodium lidar at the nearby ALOMAR Observatory. We investigate the formation mechanism of these layers using the neutral wind and temperature profiles measured directly by the lidar and the vertical motion deduced from the sodium mixing ratio. Through comparisons of the lidar data to the sporadic E in situ data, we find support for the concentration and downward motion of ions to an altitude where chemical models predict the rapid conversion of sodium ions to neutral sodium.

Acoustic-Modal Testing of the Ares I Launch Abort System Attitude Control Motor Valve

Science.gov (United States)

Davis, R. Benjamin; Fischbach, Sean R.

2010-01-01

The Attitude Control Motor (ACM) is being developed for use in the Launch Abort System (LAS) of NASA's Ares I launch vehicle. The ACM consists of a small solid rocket motor and eight actuated pintle valves that directionally allocate.thrust_- 1t.has-been- predicted-that significant unsteady. pressure.fluctuations.will.exist. inside the-valves during operation. The dominant frequencies of these oscillations correspond to the lowest several acoustic natural frequencies of the individual valves. An acoustic finite element model of the fluid volume inside the valve has been critical to the prediction of these frequencies and their associated mode shapes. This work describes an effort to experimentally validate the acoustic finite model of the valve with an acoustic modal test. The modal test involved instrumenting a flight-like valve with six microphones and then exciting the enclosed air with a loudspeaker. The loudspeaker was configured to deliver broadband noise at relatively high sound pressure levels. The aquired microphone signals were post-processed and compared to results generated from the acoustic finite element model. Initial comparisons between the test data and the model results revealed that additional model refinement was necessary. Specifically, the model was updated to implement a complex impedance boundary condition at the entrance to the valve supply tube. This boundary condition models the frequency-dependent impedance that an acoustic wave will encounter as it reaches the end of the supply tube. Upon invoking this boundary condition, significantly improved agreement between the test data and the model was realized.

A Historical Systems Study of Liquid Rocket Engine Throttling Capabilities

Science.gov (United States)

Betts, Erin M.; Frederick, Robert A., Jr.

2010-01-01

This is a comprehensive systems study to examine and evaluate throttling capabilities of liquid rocket engines. The focus of this study is on engine components, and how the interactions of these components are considered for throttling applications. First, an assessment of space mission requirements is performed to determine what applications require engine throttling. A background on liquid rocket engine throttling is provided, along with the basic equations that are used to predict performance. Three engines are discussed that have successfully demonstrated throttling. Next, the engine system is broken down into components to discuss special considerations that need to be made for engine throttling. This study focuses on liquid rocket engines that have demonstrated operational capability on American space launch vehicles, starting with the Apollo vehicle engines and ending with current technology demonstrations. Both deep throttling and shallow throttling engines are discussed. Boost and sustainer engines have demonstrated throttling from 17% to 100% thrust, while upper stage and lunar lander engines have demonstrated throttling in excess of 10% to 100% thrust. The key difficulty in throttling liquid rocket engines is maintaining an adequate pressure drop across the injector, which is necessary to provide propellant atomization and mixing. For the combustion chamber, cooling can be an issue at low thrust levels. For turbomachinery, the primary considerations are to avoid cavitation, stall, surge, and to consider bearing leakage flows, rotordynamics, and structural dynamics. For valves, it is necessary to design valves and actuators that can achieve accurate flow control at all thrust levels. It is also important to assess the amount of nozzle flow separation that can be tolerated at low thrust levels for ground testing.

Sabots, Obturator and Gas-In-Launch Tube Techniques for Heat Flux Models in Ballistic Ranges

Science.gov (United States)

Bogdanoff, David W.; Wilder, Michael C.

2013-01-01

For thermal protection system (heat shield) design for space vehicle entry into earth and other planetary atmospheres, it is essential to know the augmentation of the heat flux due to vehicle surface roughness. At the NASA Ames Hypervelocity Free Flight Aerodynamic Facility (HFFAF) ballistic range, a campaign of heat flux studies on rough models, using infrared camera techniques, has been initiated. Several phenomena can interfere with obtaining good heat flux data when using this measuring technique. These include leakage of the hot drive gas in the gun barrel through joints in the sabot (model carrier) to create spurious thermal imprints on the model forebody, deposition of sabot material on the model forebody, thereby changing the thermal properties of the model surface and unknown in-barrel heating of the model. This report presents developments in launch techniques to greatly reduce or eliminate these problems. The techniques include the use of obturator cups behind the launch package, enclosed versus open front sabot designs and the use of hydrogen gas in the launch tube. Attention also had to be paid to the problem of the obturator drafting behind the model and impacting the model. Of the techniques presented, the obturator cups and hydrogen in the launch tube were successful when properly implemented

Dr. von Braun With a Model of a Launch Vehicle

Science.gov (United States)

1950-01-01

Dr. von Braun stands beside a model of the upper stage (Earth-returnable stage) of the three-stage launch vehicle built for the series of the motion picture productions of space flight produced by Walt Disney in the mid-1950's.

Performance analysis of an IMU-augmented GNSS tracking system on board the MAIUS-1 sounding rocket

Science.gov (United States)

Braun, Benjamin; Grillenberger, Andreas; Markgraf, Markus

2018-05-01

Satellite navigation receivers are adequate tracking sensors for range safety of both orbital launch vehicles and suborbital sounding rockets. Due to high accuracy and its low system complexity, satellite navigation is seen as well-suited supplement or replacement of conventional tracking systems like radar. Having the well-known shortcomings of satellite navigation like deliberate or unintentional interferences in mind, it is proposed to augment the satellite navigation receiver by an inertial measurement unit (IMU) to enhance continuity and availability of localization. The augmented receiver is thus enabled to output at least an inertial position solution in case of signal outages. In a previous study, it was shown by means of simulation using the example of Ariane 5 that the performance of a low-grade microelectromechanical IMU is sufficient to bridge expected outages of some ten seconds, and still meeting the range safety requirements in effect. In this publication, these theoretical findings shall be substantiated by real flight data that were recorded on MAIUS-1, a sounding rocket launched from Esrange, Sweden, in early 2017. The analysis reveals that the chosen representative of a microelectromechanical IMU is suitable to bridge outages of up to thirty seconds.

FAA's Implementation of the Commercial Space Launch Amendments Act of 2004- The Experimental Permit

Science.gov (United States)

Repcheck, J. Randall

2005-12-01

A number of entrepreneurs are committed to the goal of developing and operating reusable launch vehicles for private human space travel. In order to promote this emerging industry, and to create a clear legal, regulatory, and safety regime, the United States (U.S.) Congress passed the Commercial Space Launch Amendments Act of 2004 (CSLAA). Signed on December 23, 2004 by U.S. President George W. Bush, the CSLAA makes the Federal Aviation Administration (FAA) responsible for regulating human spaceflight. The CSLAA, among other things, establishes an experimental permit regime for developmental reusable suborbital rockets. This paper describes the FAA's approach in developing guidelines for obtaining and maintaining an experimental permit, and describes the core safety elements of those guidelines.

The comparative analysis of the forecasts of development of rocket propulsion in past and now

Science.gov (United States)

Nedaivoda, A.; Prisniakov, V.

2001-03-01

Consideration is being given to use the known long and short forecasts of development of rocket engines in past - at the beginning of development of a missile engineering (K. Tsiolkovsky etc. pioneers of rocket propulsion); on the eve of launching of the artificial satellite of Earth (A. Blagonravov); after manned flight of Yu. Gagarin (V. Gluchko); after manned flight on Moon (" The Forecasts on 2001 " on materials of readings R. Goddard in USA); in middle of 70-s' years (D. Sevruk, V. Prisniakov) and at the end of 20 centure. Last years under the initiative R. Beichel and M. Pouliquen IAA. Advanced Propulsion Working Group carries out large researches on definition of the tendencies of development of rocket propulsion for the next forty years, the outcomes which one will be used in the report. The comparison of development of rocket propulsion expected to the end of 20 century and real-life is given. The report analyses the errors of the forecasts of the past - the absence reliable prognostic procedure; the euphoria of the maiden successes of conquest of space; dominance of military and political- propaganda motives of implementation of the space programs before economical; to keep developments secret; competition of two super-powers USSR and USA etc.

Evaluation of undeveloped rocket engine cycle applications to advanced transportation

Science.gov (United States)

1990-01-01

Undeveloped pump-fed, liquid propellant rocket engine cycles were assessed and evaluated for application to Next Manned Transportation System (NMTS) vehicles, which would include the evolving Space Transportation System (STS Evolution), the Personnel Launch System (PLS), and the Advanced Manned Launch System (AMLS). Undeveloped engine cycles selected for further analysis had potential for increased reliability, more maintainability, reduced cost, and improved (or possibly level) performance when compared to the existing SSME and proposed STME engines. The split expander (SX) cycle, the full flow staged combustion (FFSC) cycle, and a hybrid version of the FFSC, which has a LOX expander drive for the LOX pump, were selected for definition and analysis. Technology requirements and issues were identified and analyses of vehicle systems weight deltas using the SX and FFSC cycles in AMLS vehicles were performed. A strawman schedule and cost estimate for FFSC subsystem technology developments and integrated engine system demonstration was also provided.

Eddie Rocket's Franchise

OpenAIRE

Vahter, Jenni

2008-01-01

Eddie Rocket's Franchise - Setting up a franchise restaurant in Helsinki. TIIVISTELMÄ: Eddie Rocket's on menestynyt amerikkalaistyylinen 1950-luvun ”diner” franchiseravintolaketju Irlannista. Ravintoloita on perustettu viimeisen 18 vuoden aikana 28 kappaletta Irlantiin ja Isoon Britanniaan sekä yksi Espanjaan. Tämän tutkimuksen tarkoitus on tutkia onko Eddie Rocket'silla potentiaalia menestyä Helsingissä, Suomessa. Tutkimuskysymystä on lähestytty toimiala-analyysin, markkinatutkimuksen j...

Development of Kabila rocket: A radioisotope heated thermionic plasma rocket engine

Directory of Open Access Journals (Sweden)

Kalomba Mboyi

2015-04-01

Full Text Available A new type of plasma rocket engine, the Kabila rocket, using a radioisotope heated thermionic heating chamber instead of a conventional combustion chamber or catalyst bed is introduced and it achieves specific impulses similar to the ones of conventional solid and bipropellant rockets. Curium-244 is chosen as a radioisotope heat source and a thermal reductive layer is also used to obtain precise thermionic emissions. The self-sufficiency principle is applied by simultaneously heating up the emitting material with the radioisotope decay heat and by powering the different valves of the plasma rocket engine with the same radioisotope decay heat using a radioisotope thermoelectric generator. This rocket engine is then benchmarked against a 1 N hydrazine thruster configuration operated on one of the Pleiades-HR-1 constellation spacecraft. A maximal specific impulse and power saving of respectively 529 s and 32% are achieved with helium as propellant. Its advantages are its power saving capability, high specific impulses and simultaneous ease of storage and restart. It can however be extremely voluminous and potentially hazardous. The Kabila rocket is found to bring great benefits to the existing spacecraft and further research should optimize its geometric characteristics and investigate the physical principals of its operation.

EUVS Sounding Rocket Payload

Science.gov (United States)

Stern, Alan S.

1996-01-01

During the first half of this year (CY 1996), the EUVS project began preparations of the EUVS payload for the upcoming NASA sounding rocket flight 36.148CL, slated for launch on July 26, 1996 to observe and record a high-resolution (approx. 2 A FWHM) EUV spectrum of the planet Venus. These preparations were designed to improve the spectral resolution and sensitivity performance of the EUVS payload as well as prepare the payload for this upcoming mission. The following is a list of the EUVS project activities that have taken place since the beginning of this CY: (1) Applied a fresh, new SiC optical coating to our existing 2400 groove/mm grating to boost its reflectivity; (2) modified the Ranicon science detector to boost its detective quantum efficiency with the addition of a repeller grid; (3) constructed a new entrance slit plane to achieve 2 A FWHM spectral resolution; (4) prepared and held the Payload Initiation Conference (PIC) with the assigned NASA support team from Wallops Island for the upcoming 36.148CL flight (PIC held on March 8, 1996; see Attachment A); (5) began wavelength calibration activities of EUVS in the laboratory; (6) made arrangements for travel to WSMR to begin integration activities in preparation for the July 1996 launch; (7) paper detailing our previous EUVS Venus mission (NASA flight 36.117CL) published in Icarus (see Attachment B); and (8) continued data analysis of the previous EUVS mission 36.137CL (Spica occultation flight).

Design and qualification of an UHV system for operation on sounding rockets

Energy Technology Data Exchange (ETDEWEB)

Grosse, Jens, E-mail: jens.grosse@dlr.de; Braxmaier, Claus [Center of Applied Space Technology and Microgravity (ZARM), University of Bremen, Bremen, 28359, Germany and German Aerospace Center (DLR) Bremen, Bremen, 28359 (Germany); Seidel, Stephan Tobias; Becker, Dennis; Lachmann, Maike Diana [Institute of Quantum Optics, Leibniz University Hanover, Hanover, 30167 (Germany); Scharringhausen, Marco [German Aerospace Center (DLR) Bremen, Bremen, 28359 (Germany); Rasel, Ernst Maria [Institute of Quantum Optics, Leibniz University Hanover, Hanover, 30167, Bremen (Germany)

2016-05-15

The sounding rocket mission MAIUS-1 has the objective to create the first Bose–Einstein condensate in space; therefore, its scientific payload is a complete cold atom experiment built to be launched on a VSB-30 sounding rocket. An essential part of the setup is an ultrahigh vacuum system needed in order to sufficiently suppress interactions of the cooled atoms with the residual background gas. Contrary to vacuum systems on missions aboard satellites or the international space station, the required vacuum environment has to be reached within 47 s after motor burn-out. This paper contains a detailed description of the MAIUS-1 vacuum system, as well as a description of its qualification process for the operation under vibrational loads of up to 8.1 g{sub RMS} (where RMS is root mean square). Even though a pressure rise dependent on the level of vibration was observed, the design presented herein is capable of regaining a pressure of below 5 × 10{sup −10} mbar in less than 40 s when tested at 5.4 g{sub RMS}. To the authors' best knowledge, it is the first UHV system qualified for operation on a sounding rocket.

Structured waves near the plasma frequency observed in three auroral rocket flights

Directory of Open Access Journals (Sweden)

M. Samara

2006-11-01

Full Text Available We present observations of waves at and just above the plasma frequency (fpe from three high frequency electric field experiments on three recent rockets launched to altitudes of 300–900 km in active aurora. The predominant observed HF waves just above fpe are narrowband, short-lived emissions with amplitudes ranging from <1 mV/m to 20 mV/m, often associated with structured electron density. The nature of these HF waves, as determined from frequency-time spectrograms, is highly variable: in some cases, the frequency decreases monotonically with time as in the "HF-chirps" previously reported (McAdams and LaBelle, 1999, but in other cases rising frequencies are observed, or features which alternately rise and fall in frequency. They exhibit two timescales of amplitude variation: a short timescale, typically 50–100 ms, associated with individual discrete features, and a longer timescale associated with the general decrease in the amplitudes of the emissions as the rocket moves away from where the condition f~fpe holds. The latter timescale ranges from 0.6 to 6.0 s, corresponding to distances of 2–7 km, assuming the phenomenon to be stationary and using the rocket velocity to convert time to distance.

Development of CFD model for augmented core tripropellant rocket engine

Science.gov (United States)

Jones, Kenneth M.

1994-10-01

The Space Shuttle era has made major advances in technology and vehicle design to the point that the concept of a single-stage-to-orbit (SSTO) vehicle appears more feasible. NASA presently is conducting studies into the feasibility of certain advanced concept rocket engines that could be utilized in a SSTO vehicle. One such concept is a tripropellant system which burns kerosene and hydrogen initially and at altitude switches to hydrogen. This system will attain a larger mass fraction because LOX-kerosene engines have a greater average propellant density and greater thrust-to-weight ratio. This report describes the investigation to model the tripropellant augmented core engine. The physical aspects of the engine, the CFD code employed, and results of the numerical model for a single modular thruster are discussed.

Integral performance optimum design for multistage solid propellant rocket motors

Energy Technology Data Exchange (ETDEWEB)

Zhang, Hongtao (Shaanxi Power Machinery Institute (China))

1989-04-01

A mathematical model for integral performance optimization of multistage solid propellant rocket motors is presented. A calculation on a three-stage, volume-fixed, solid propellant rocket motor is used as an example. It is shown that the velocity at burnout of intermediate-range or long-range ballistic missile calculated using this model is four percent greater than that using the usual empirical method.

Two-Rockets Thought Experiment

Science.gov (United States)

Smarandache, Florentin

2014-03-01

Let n>=2 be identical rockets: R1 ,R2 , ..., Rn. Each of them moving at constant different velocities respectively v1, v2, ..., vn on parallel directions in the same sense. In each rocket there is a light clock, the observer on earth also has a light clock. All n + 1 light clocks are identical and synchronized. The proper time Δt' in each rocket is the same. Let's focus on two arbitrary rockets Ri and Rjfrom the previous n rockets. Let's suppose, without loss of generality, that their speeds verify virocket Rj is contracted with the factor C(vj -vi) , i.e. Lj =Lj' C(vj -vi) .(2) But in the reference frame of the astronaut in Rjit is like rocket Rjis stationary andRi moves with the speed vj -vi in opposite direction. Therefore, similarly, the non-proper time interval as measured by the astronaut inRj with respect to the event inRi is dilated with the same factor D(vj -vi) , i.e. Δtj . i = Δt' D(vj -vi) , and rocketRi is contracted with the factor C(vj -vi) , i.e. Li =Li' C(vj -vi) .But it is a contradiction to have time dilations in both rockets. (3) Varying i, j in {1, 2, ..., n} in this Thought Experiment we get again other multiple contradictions about time dilations. Similarly about length contractions, because we get for a rocket Rj, n-2 different length contraction factors: C(vj -v1) , C(vj -v2) , ..., C(vj -vj - 1) , C(vj -vj + 1) , ..., C(vj -vn) simultaneously! Which is abnormal.

Rocketing into the future the history and technology of rocket planes

CERN Document Server

van Pelt, Michel

2012-01-01

Rocketing into the Future journeys into the exciting world of rocket planes, examining the exotic concepts and actual flying vehicles that have been devised over the last one hundred years. Lavishly illustrated with over 150 photographs, it recounts the history of rocket planes from the early pioneers who attached simple rockets on to their wooden glider airplanes to the modern world of high-tech research vehicles. The book then looks at the possibilities for the future. The technological and economic challenges of the Space Shuttle proved insurmountable, and thus the program was unable to fulfill its promise of low-cost access to space. However, the burgeoning market of suborbital space tourism may yet give the necessary boost to the development of a truly reusable spaceplane.

An analysis of the orbital distribution of solid rocket motor slag

Science.gov (United States)

Horstman, Matthew F.; Mulrooney, Mark

2009-01-01

The contribution by solid rocket motors (SRMs) to the orbital debris environment is potentially significant and insufficiently studied. Design and combustion processes can lead to the emission of enough by-products to warrant assessment of their contribution to orbital debris. These particles are formed during SRM tail-off, or burn termination, by the rapid solidification of molten Al2O3 slag accumulated during the burn. The propensity of SRMs to generate particles larger than 100μm raises concerns regarding the debris environment. Sizes as large as 1 cm have been witnessed in ground tests, and comparable sizes have been estimated via observations of sub-orbital tail-off events. Utilizing previous research we have developed more sophisticated size distributions and modeled the time evolution of resultant orbital populations using a historical database of SRM launches, propellant, and likely location and time of tail-off. This analysis indicates that SRM ejecta is a significant component of the debris environment.

Inter-Hemispheric Coupling During Recent North Polar Summer Periods as Predicted by MaCWAVE/MIDAS Rocket Data and Traced by TIMED/SABER Measurements

Science.gov (United States)

Goldberg, Richard A.; Feofilov, Artem G.; Kutepov, Alexander A.; Pesnell W. Dean; Schmidlin, Francis J.

2011-01-01

In July, 2002, the MaCWAVE-MIDAS Rocket Program was launched from Andoya Rocket Range (ARR) in Norway. Data from these flights demonstrated that the polar summer mesosphere during this period was unusual, at least above ARR. Theoretical studies have since been published that imply that the abnormal characteristics of this polar summer were generated by dynamical processes occurring in the southern polar winter hemisphere. We have used data from the SABER instrument aboard the NASA Thermosphere Ionosphere Mesosphere Energetics and Dynamics (TIMED) Satellite to study these characteristics and compare them with the features observed in the ensuing eight years. For background, the TIMED Satellite was launched on December 7,2001 to study the dynamics and energy of the mesosphere and lower thermosphere. The SABER instrument is a limb scanning infrared radiometer designed to measure temperature of the region as well as a large number of minor constituents. In this study, we review the MaCWAVE rocket results. Next, we investigate the temperature characteristics of the polar mesosphere as a function of spatial and temporal considerations. We have used the most recent SABER dataset (1.07). Weekly averages are used to make comparisons between the winter and summer hemispheres. Furthermore, the data analysis agrees with recent theoretical studies showing that this behavior is a result of anomalous dynamical events in the southern hemisphere. The findings discussed here clearly show the value of scientific rocket flights used in a discovery mode.

Environmental Conditions and Threatened and Endangered Species Populations near the Titain, Atlas, and Delta Launch Complexes, Cape Canaveral Air Station

Science.gov (United States)

Oddy, Donna M.; Stolen, Eric D.; Schmalzer, Paul A.; Hensley, Melissa A.; Hall, Patrice; Larson, Vickie L.; Turek, Shannon R.

1999-01-01

Launches of Delta, Atlas, and Titan rockets from Cape Canaveral Air Station (CCAS) have potential environmental effects. These could occur from direct impacts of launches or indirectly from habitat alterations. This report summarizes a three-year study (1995-1998) characterizing the environment, with particular attention to threatened and endangered species, near Delta, Atlas, and Titan launch facilities. Cape Canaveral has been modified by Air Force development and by 50 years of fire suppression. The dominant vegetation type around the Delta and Atlas launch complexes is coastal oak hammock forest. Oak scrub is the predominant upland vegetation type near the Titan launch complexes. Compositionally, these are coastal scrub communities that has been unburned for greater than 40 years and have developed into closed canopy, low-stature forests. Herbaceous vegetation around active and inactive facilities, coastal strand and dune vegetation near the Atlantic Ocean, and exotic vegetation in disturbed areas are common. Marsh and estuarine vegetation is most common west of the Titan complexes. Launch effects to vegetation include scorch, acid, and particulate deposition. Discernable, cumulative effects are limited to small areas near the launch complexes. Water quality samples were collected at the Titan, Atlas, and Delta launch complexes in September 1995 (wet season) and January 1996 (dry season). Samples were analyzed for heavy metals, chloride, total organic carbon, calcium, iron, magnesium, sodium, total alkalinity, pH, and conductivity. Differences between fresh, brackish, and saline surface waters were evident. The natural buffering capacity of the environment surrounding the CCAS launch complexes is adequate for neutralizing acid deposition in rainfall and launch deposition. Populations of the Florida Scrub-Jay (Aphelocoma coerulescens), a Federally- listed, threatened species, reside near the launch complexes. Thirty-seven to forty-one scrub-jay territories were

Environmental Conditions and Threatened and Endangered Species Populations near the Titan, Atlas, and Delta Launch Complexes, Cape Canaveral Air Station

Science.gov (United States)

Oddy, Donna M.; Stolen, Eric D.; Schmalzer, Paul A.; Hensley, Melissa A.; Hall, Patrice; Larson, Vickie L.; Turek, Shannon R.

1999-01-01

Launches of Delta, Atlas, and Titan rockets from Cape Canaveral Air Station (CCAS) have potential environmental effects. These could occur from direct impacts of launches or indirectly from habitat alterations. This report summarizes a three-year study (1 995-1 998) characterizing the environment, with particular attention to threatened and endangered species, near Delta, Atlas, and Titan launch facilities. Cape Canaveral has been modified by Air Force development and by 50 years of fire suppression. The dominant vegetation type around the Delta and Atlas launch complexes is coastal oak hammock forest. Oak scrub is the predominant upland vegetation type near the Titan launch complexes. Compositionally, these are coastal scrub communities that has been unburned for > 40 years and have developed into closed canopy, low-stature forests. Herbaceous vegetation around active and inactive facilities, coastal strand and dune vegetation near the Atlantic Ocean, and exotic vegetation in disturbed areas are common. Marsh and estuarine vegetation is most common west of the Titan complexes. Launch effects to vegetation include scorch, acid, and particulate deposition. Discernable, cumulative effects are limited to small areas near the launch complexes. Water quality samples were collected at the Titan, Atlas, and Delta launch complexes in September 1995 (wet season) and January 1996 (dry season). Samples were analyzed for heavy metals, chloride, total organic carbon, calcium, iron, magnesium, sodium, total alkalinity, pH, and conductivity. Differences between fresh, brackish, and saline surface waters were evident. The natural buffering capacity of the environment surrounding the CCAS launch complexes is adequate for neutralizing acid deposition in rainfall and launch deposition. Populations of the Florida Scrub-Jay (Aphelocoma coerulescens), a Federally-listed, threatened species, reside near the launch complexes. Thirty-seven to forty-one scrub-jay territories were located at

Modeling of Rocket Fuel Heating and Cooling Processes in the Interior Receptacle Space of Ground-Based Systems

Directory of Open Access Journals (Sweden)

K. I. Denisova

2016-01-01

Full Text Available The propellant to fill the fuel tanks of the spacecraft, upper stages, and space rockets on technical and ground-based launch sites before fueling should be prepared to ensure many of its parameters, including temperature, in appropriate condition. Preparation of fuel temperature is arranged through heating and cooling the rocket propellants (RP in the tanks of fueling equipment. Processes of RP temperature preparation are the most energy-intensive and timeconsuming ones, which require that a choice of sustainable technologies and modes of cooling (heating RP provided by the ground-based equipment has been made through modeling of the RP [1] temperature preparation processes at the stage of design and operation of the groundbased fueling equipment.The RP temperature preparation in the tanks of the ground-based systems can be provided through the heat-exchangers built-in the internal space and being external with respect to the tank in which antifreeze, air or liquid nitrogen may be used as the heat transfer media. The papers [1-12], which note a promising use of the liquid nitrogen to cool PR, present schematic diagrams and modeling systems for the RP temperature preparation in the fueling equipment of the ground-based systems.We consider the RP temperature preparation using heat exchangers to be placed directly in RP tanks. Feeding the liquid nitrogen into heat exchanger with the antifreeze provides the cooling mode of PR while a heated air fed there does that of heating. The paper gives the systems of equations and results of modeling the processes of RP temperature preparation, and its estimated efficiency.The systems of equations of cooling and heating RP are derived on the assumption that the heat exchange between the fuel and the antifreeze, as well as between the storage tank and the environment is quasi-stationary.The paper presents calculation results of the fuel temperature in the tank, and coolant temperature in the heat exchanger, as

Sounding rocket flight report: MUMP 9 and MUMP 10

Science.gov (United States)

Grassl, H. J.

1971-01-01

The results of the launching of two Marshall-University of Michigan Probes (MUMP 9 and MUMP 10), Nike-Tomahawk sounding rocket payloads, are summarized. The MUMP 9 paylaod included an omegatron mass analyzer, a molecular fluorescence densitometer, a mini-tilty filter, and a lunar position sensor. This complement of instruments permitted the determination of the molecular nitrogen density and temperature in the altitude range from approximately 143 to 297 km over Wallops Island, Virginia, during January 1971. The MUMP 10 payload included an omegatron mass analyzer, an electron temperature probe (Spencer, Brace, and Carignan, 1962), a cryogenic densitometer, and a solar position sensor. This complement of instruments permitted the determination of the molecular nitrogen density and temperature and the charged particle density and temperature in the altitude range from approximately 145 to 290 km over Wallops Island, Virginia, during the afternoon preceding the MUMP 9 launch in January 1971. A general description of the payload kinematics, orientation analysis, and the technique for the reduction and analysis of the data is given.

Origin of how steam rockets can reduce space transport cost by orders of magnitude

International Nuclear Information System (INIS)

Zuppero, A.; Larson, T.K.; Schnitzler, B.G.; Rice, J.W.; Hill, T.J.; Richins, W.D.; Parlier, L.; Werner, J.E.

1999-01-01

A brief sketch shows the origin of why and how thermal rocket propulsion has the unique potential to dramatically reduce the cost of space transportation for most inner solar system missions of interest. Orders of magnitude reduction in cost are apparently possible when compared to all processes requiring electrolysis for the production of rocket fuels or propellants and to all electric propulsion systems. An order of magnitude advantage can be attributed to rocket propellant tank factors associated with storing water propellant, compared to cryogenic liquids. An order of magnitude can also be attributed to the simplicity of the extraction and processing of ice on the lunar surface, into an easily stored, non-cryogenic rocket propellant (water). A nuclear heated thermal rocket can deliver thousands of times its mass to Low Earth Orbit from the Lunar surface, providing the equivalent to orders of magnitude drop in launch cost for mass in Earth orbit. Mass includes water ice. These cost reductions depend (exponentially) on the mission delta-v requirements being less than about 6 km/s, or about 3 times the specific velocity of steam rockets (2 km/s, from Isp 200 sec). Such missions include: from the lunar surface to Low Lunar Orbit, (LLO), from LLO to lunar escape, from Low Earth Orbit (LEO) to Geosynchronous Orbit (GEO), from LEO to Earth Escape, from LEO to Mars Transfer Orbit, from LLO to GEO, missions returning payloads from about 10% of the periodic comets using propulsive capture to orbits around Earth itself, and fast, 100 day missions from Lunar Escape to Mars. All the assertions depend entirely and completely on the existence of abundant, nearly pure ice at the permanently dark North and South Poles of the Moon. copyright 1999 American Institute of Physics

Rocket observations at the northern edge of the eastward electrojet

International Nuclear Information System (INIS)

Cahill, L.J. Jr.; Arnoldy, R.L.; Taylor, W.W.L.

1980-01-01

A Nike-Tomahawk rocket was launched north over two quiet, late evening auroral arcs in March 1975. A northward magnetic disturbance was observed on the ground under the rocket trajectory. South of the arcs the northward electric field was 60 mV/m, indicating strong westward plasma flow. An eastward electrojet current layer was penetrated in the upward flight. Precipitating electrons were observed over each arc. The electric field decreased to below 20 mV/m over each arc and recovered to 50 mV/m between the arcs. Using the observed electron flux and a model of the ionosphere, the Hall and Pedersen conductivities were calculated. These conductivities were used, with the observed electric field, to calculate the eastward and northward components of the horizontal ionospheric currents. The eastward current calculated south of the first arc agreed well with the observed eastward electrojet current. The power dissipated by the Pedersen current Σ/sub p/EPSILON 2 was also calculated and compared with the power carried by the precipitating electrons. The Joule power decreased abruptly in the auroral arcs, as the pricipitating electron power increased. The total dissipated power was approximately the same inside the arcs, between them, and south of the luminosity. North of the aurora the electric field and dissipated power remained low. Field-aligned currents carried by the observed electrons were about a factor of 3 lower than those inferred from the magnetic field measurements. Likewise, current continuity arguments to keep the auroral current system divergence free required a larger field-aligned current than that obtained from the particle detectors

Mean Flow Augmented Acoustics in Rocket Systems

Science.gov (United States)

Fischbach, Sean R.

2014-01-01

Oscillatory motion in solid rocket motors and liquid engines has long been a subject of concern. Many rockets display violent fluctuations in pressure, velocity, and temperature originating from the complex interactions between the combustion process and gas dynamics. The customary approach to modeling acoustic waves inside a rocket chamber is to apply the classical inhomogeneous wave equation to the combustion gas. The assumption of a linear, non-dissipative wave in a quiescent fluid remains valid while the acoustic amplitudes are small and local gas velocities stay below Mach 0.2. The converging section of a rocket nozzle, where gradients in pressure, density, and velocity become large, is a notable region where this approach is not applicable. The expulsion of unsteady energy through the nozzle of a rocket is identified as the predominate source of acoustic damping for most rocket systems. An accurate model of the acoustic behavior within this region where acoustic modes are influenced by the presence of a steady mean flow is required for reliable stability predictions. Recently, an approach to address nozzle damping with mean flow effects was implemented by French [1]. This new approach extends the work originated by Sigman and Zinn [2] by solving the acoustic velocity potential equation (AVPE) formulated by perturbing the Euler equations [3]. The acoustic velocity potential (psi) describing the acoustic wave motion in the presence of an inhomogeneous steady high-speed flow is defined by, (del squared)(psi) - (lambda/c)(exp 2)(psi) - M(dot)[M(dot)(del)(del(psi))] - 2(lambda(M/c) + (M(dot)del(M))(dot)del(psi)-2(lambda)(psi)[M(dot)del(1/c)]=0 (1) with M as the Mach vector, c as the speed of sound, and lambda as the complex eigenvalue. French apply the finite volume method to solve the steady flow field within the combustion chamber and nozzle with inviscid walls. The complex eigenvalues and eigenvector are determined with the use of the ARPACK eigensolver. The

Launch Vehicle Demonstrator Using Shuttle Assets

Science.gov (United States)

Threet, Grady E., Jr.; Creech, Dennis M.; Philips, Alan D.; Water, Eric D.

2011-01-01

The Marshall Space Flight Center Advanced Concepts Office (ACO) has the leading role for NASA s preliminary conceptual launch vehicle design and performance analysis. Over the past several years the ACO Earth-to-Orbit Team has evaluated thousands of launch vehicle concept variations for a multitude of studies including agency-wide efforts such as the Exploration Systems Architecture Study (ESAS), Constellation, Heavy Lift Launch Vehicle (HLLV), Heavy Lift Propulsion Technology (HLPT), Human Exploration Framework Team (HEFT), and Space Launch System (SLS). NASA plans to continue human space exploration and space station utilization. Launch vehicles used for heavy lift cargo and crew will be needed. One of the current leading concepts for future heavy lift capability is an inline one and a half stage concept using solid rocket boosters (SRB) and based on current Shuttle technology and elements. Potentially, the quickest and most cost-effective path towards an operational vehicle of this configuration is to make use of a demonstrator vehicle fabricated from existing shuttle assets and relying upon the existing STS launch infrastructure. Such a demonstrator would yield valuable proof-of-concept data and would provide a working test platform allowing for validated systems integration. Using shuttle hardware such as existing RS-25D engines and partial MPS, propellant tanks derived from the External Tank (ET) design and tooling, and four-segment SRB s could reduce the associated upfront development costs and schedule when compared to a concept that would rely on new propulsion technology and engine designs. There are potentially several other additional benefits to this demonstrator concept. Since a concept of this type would be based on man-rated flight proven hardware components, this demonstrator has the potential to evolve into the first iteration of heavy lift crew or cargo and serve as a baseline for block upgrades. This vehicle could also serve as a demonstration

Evaluation of thermal sprayed metallic coatings for use on the structures at Launch Complex 39

Science.gov (United States)

Welch, Peter J.

1990-01-01

The current status of the evaluation program is presented. The objective was to evaluate the applicability of Thermal Sprayed Coatings (TSC) to protect the structures in the high temperature acid environment produced by exhaust of the Solid Rocket Boosters during the launches of the Shuttle Transportation System. Only the relatively low cost aluminum TSC which provides some cathodic protection for steel appears to be a practical candidate for further investigation.

Propulsion/ASME Rocket-Based Combined Cycle Activities in the Advanced Space Transportation Program Office

Science.gov (United States)

Hueter, Uwe; Turner, James

1998-01-01

NASA's Office Of Aeronautics and Space Transportation Technology (OASTT) has establish three major coals. "The Three Pillars for Success". The Advanced Space Transportation Program Office (ASTP) at the NASA's Marshall Space Flight Center in Huntsville,Ala. focuses on future space transportation technologies under the "Access to Space" pillar. The Advanced Reusable Technologies (ART) Project, part of ASTP, focuses on the reusable technologies beyond those being pursued by X-33. The main activity over the past two and a half years has been on advancing the rocket-based combined cycle (RBCC) technologies. In June of last year, activities for reusable launch vehicle (RLV) airframe and propulsion technologies were initiated. These activities focus primarily on those technologies that support the year 2000 decision to determine the path this country will take for Space Shuttle and RLV. In February of this year, additional technology efforts in the reusable technologies were awarded. The RBCC effort that was completed early this year was the initial step leading to flight demonstrations of the technology for space launch vehicle propulsion. Aerojet, Boeing-Rocketdyne and Pratt & Whitney were selected for a two-year period to design, build and ground test their RBCC engine concepts. In addition, ASTROX, Pennsylvania State University (PSU) and University of Alabama in Huntsville also conducted supporting activities. The activity included ground testing of components (e.g., injectors, thrusters, ejectors and inlets) and integrated flowpaths. An area that has caused a large amount of difficulty in the testing efforts is the means of initiating the rocket combustion process. All three of the prime contractors above were using silane (SiH4) for ignition of the thrusters. This follows from the successful use of silane in the NASP program for scramjet ignition. However, difficulties were immediately encountered when silane (an 80/20 mixture of hydrogen/silane) was used for rocket

Another Look at Rocket Thrust

Science.gov (United States)

Hester, Brooke; Burris, Jennifer

2012-01-01

Rocket propulsion is often introduced as an example of Newton's third law. The rocket exerts a force on the exhaust gas being ejected; the gas exerts an equal and opposite force--the thrust--on the rocket. Equivalently, in the absence of a net external force, the total momentum of the system, rocket plus ejected gas, remains constant. The law of…

'RCHX-1-STORM' first Slovenian meteorological rocket program

Science.gov (United States)

Kerstein, Aleksander; Matko, Drago; Trauner, Amalija; Britovšek, Zvone

2004-08-01

Astronautic and Rocket Society Celje (ARSC) formed a special working team for research and development of a small meteorological hail suppression rocket in the 70th. The hail suppression system was established in former Yugoslavia in the late 60th as an attempt to protect important agricultural regions from one of the summer's most vicious storm. In this time Slovenia was a part of Yugoslavia as one of the federal republic with relative high developed agricultural region production. The Rocket program 'RCHX-STORM' was a second attempt, for Slovenia indigenously developed in the production of meteorological hail suppression rocket. ARSC has designed a family of small sounding rocket that were based on highly promising hybrid propellant propulsion. Hybrid propulsion was selected for this family because it was offering low cost, save production and operation and simple logistics. Conventional sounding rockets use solid propellant motor for their propulsion. The introduction of hybrid motors has enabled a considerable decrease in overall cost. The transportation handling and storage procedures were greatly simplified due to the fact that a hybrid motor was not considered as explosive matter. A hybrid motor may also be designed to stand a severe environment without resorting to conditioning arrangements. The program started in the late 70th when the team ARSC was integrated in the Research and Development Institute in Celje (RDIC). The development program aimed to produce three types of meteorological rockets with diameters 76, 120 and 160 mm. Development of the RCHX-76 engine and rocket vehicle including flight certification has been undertaken by a joint team comprising of the ARCS, RDIC and the company Cestno podjetje Celje (CPC), Road building company Celje. Many new techniques and methods were used in this program such as computer simulation of external and internal ballistics, composite materials for rocket construction, intensive static testing of models and

Stochastic rocket dynamics under random nozzle side loads: Ornstein-Uhlenbeck boundary layer separation and its coarse grained connection to side loading and rocket response

Energy Technology Data Exchange (ETDEWEB)

Keanini, R.G.; Srivastava, N.; Tkacik, P.T. [Department of Mechanical Engineering, University of North Carolina at Charlotte, 9201 University City Blvd., Charlotte, NC 28078 (United States); Weggel, D.C. [Department of Civil and Environmental Engineering, University of North Carolina at Charlotte, 9201 University City Blvd., Charlotte, NC 28078 (United States); Knight, P.D. [Mitchell Aerospace and Engineering, Statesville, North Carolina 28677 (United States)

2011-06-15

A long-standing, though ill-understood problem in rocket dynamics, rocket response to random, altitude-dependent nozzle side-loads, is investigated. Side loads arise during low altitude flight due to random, asymmetric, shock-induced separation of in-nozzle boundary layers. In this paper, stochastic evolution of the in-nozzle boundary layer separation line, an essential feature underlying side load generation, is connected to random, altitude-dependent rotational and translational rocket response via a set of simple analytical models. Separation line motion, extant on a fast boundary layer time scale, is modeled as an Ornstein-Uhlenbeck process. Pitch and yaw responses, taking place on a long, rocket dynamics time scale, are shown to likewise evolve as OU processes. Stochastic, altitude-dependent rocket translational motion follows from linear, asymptotic versions of the full nonlinear equations of motion; the model is valid in the practical limit where random pitch, yaw, and roll rates all remain small. Computed altitude-dependent rotational and translational velocity and displacement statistics are compared against those obtained using recently reported high fidelity simulations [Srivastava, Tkacik, and Keanini, J. Appl. Phys. 108, 044911 (2010)]; in every case, reasonable agreement is observed. As an important prelude, evidence indicating the physical consistency of the model introduced in the above article is first presented: it is shown that the study's separation line model allows direct derivation of experimentally observed side load amplitude and direction densities. Finally, it is found that the analytical models proposed in this paper allow straightforward identification of practical approaches for: (i) reducing pitch/yaw response to side loads, and (ii) enhancing pitch/yaw damping once side loads cease. (Copyright copyright 2011 WILEY-VCH Verlag GmbH and Co. KGaA, Weinheim)

Chinese modify CZ-2/3 rocket boosters, focus on commercial launch market

Science.gov (United States)

Covault, C.

1985-07-01

A program underway in the People's Republic of China to modify the Titan-class CZ-2/3 satellite-launch and ICBM boosters is described on the basis of a recent visit to the manufacturing plant in Shanghai. The present two-stage CZ-2 and three-stage CZ-3 can place 5000 lbs in LEO or 3080 lbs in GEO, respectively, and are produced on a custom basis with a delivery time of about 2 yrs. Modifications introduced include 4 x 6-ft fins and a pogo-suppression system for the four-engine first stage and a steel support band for the combustion chamber of the 80-ton-thrust second-stage main engine.

Nuclear Thermal Rocket Simulation in NPSS

Science.gov (United States)

Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas M.

2013-01-01

Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic-metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

THE ADIABATIC DEMAGNETIZATION REFRIGERATOR FOR THE MICRO-X SOUNDING ROCKET TELESCOPE

International Nuclear Information System (INIS)

Wikus, P.; Bagdasarova, Y.; Figueroa-Feliciano, E.; Leman, S. W.; Rutherford, J. M.; Trowbridge, S. N.; Adams, J. S.; Bandler, S. R.; Eckart, M. E.; Kelley, R. L.; Kilbourne, C. A.; Porter, F. S.; Doriese, W. B.; McCammon, D.

2010-01-01

The Micro-X Imaging X-ray Spectrometer is a sounding rocket payload slated for launch in 2011. An array of Transition Edge Sensors, which is operated at a bath temperature of 50 mK, will be used to obtain a high resolution spectrum of the Puppis-A supernova remnant. An Adiabatic Demagnetization Refrigerator (ADR) with a 75 gram Ferric Ammonium Alum (FAA) salt pill in the bore of a 4 T superconducting magnet provides a stable heat sink for the detector array only a few seconds after burnout of the rocket motors. This requires a cold stage design with very short thermal time constants. A suspension made from Kevlar strings holds the 255 gram cold stage in place. It is capable of withstanding loads in excess of 200 g. Stable operation of the TES array in proximity to the ADR magnet is ensured by a three-stage magnetic shielding system which consists of a superconducting can, a high-permeability shield and a bucking coil. The development and testing of the Micro-X payload is well underway.

Rocket Engine Innovations Advance Clean Energy

Science.gov (United States)

2012-01-01

During launch countdown, at approximately T-7 seconds, the Space Shuttle Main Engines (SSMEs) roar to life. When the controllers indicate normal operation, the solid rocket boosters ignite and the shuttle blasts off. Initially, the SSMEs throttle down to reduce stress during the period of maximum dynamic pressure, but soon after, they throttle up to propel the orbiter to 17,500 miles per hour. In just under 9 minutes, the three SSMEs burn over 1.6 million pounds of propellant, and temperatures inside the main combustion chamber reach 6,000 F. To cool the engines, liquid hydrogen circulates through miles of tubing at -423 F. From 1981to 2011, the Space Shuttle fleet carried crew and cargo into orbit to perform a myriad of unprecedented tasks. After 30 years and 135 missions, the feat of engineering known as the SSME boasted a 100-percent flight success rate.

Dynamic Analysis of Sounding Rocket Pneumatic System Revision

Science.gov (United States)

Armen, Jerald

2010-01-01

The recent fusion of decades of advancements in mathematical models, numerical algorithms and curve fitting techniques marked the beginning of a new era in the science of simulation. It is becoming indispensable to the study of rockets and aerospace analysis. In pneumatic system, which is the main focus of this paper, particular emphasis will be placed on the efforts of compressible flow in Attitude Control System of sounding rocket.

Simulation and experimental research on line throwing rocket with flight

OpenAIRE

Wen-bin Gu; Ming Lu; Jian-qing Liu; Qin-xing Dong; Zhen-xiong Wang; Jiang-hai Chen

2014-01-01

The finite segment method is used to model the line throwing rocket system. A dynamic model of line throwing rocket with flight motion based on Kane's method is presented by the kinematics description of the system and the consideration of the forces acting on the system. The experiment designed according to the parameters of the dynamic model is made. The simulation and experiment results, such as range, velocity and flight time, are compared and analyzed. The simulation results are basicall...

Experimental and Computational Modal Analyses for Launch Vehicle Models considering Liquid Propellant and Flange Joints

Directory of Open Access Journals (Sweden)

Chang-Hoon Sim

2018-01-01

Full Text Available In this research, modal tests and analyses are performed for a simplified and scaled first-stage model of a space launch vehicle using liquid propellant. This study aims to establish finite element modeling techniques for computational modal analyses by considering the liquid propellant and flange joints of launch vehicles. The modal tests measure the natural frequencies and mode shapes in the first and second lateral bending modes. As the liquid filling ratio increases, the measured frequencies decrease. In addition, as the number of flange joints increases, the measured natural frequencies increase. Computational modal analyses using the finite element method are conducted. The liquid is modeled by the virtual mass method, and the flange joints are modeled using one-dimensional spring elements along with the node-to-node connection. Comparison of the modal test results and predicted natural frequencies shows good or moderate agreement. The correlation between the modal tests and analyses establishes finite element modeling techniques for modeling the liquid propellant and flange joints of space launch vehicles.

Reusable launch vehicle model uncertainties impact analysis

Science.gov (United States)

Chen, Jiaye; Mu, Rongjun; Zhang, Xin; Deng, Yanpeng

2018-03-01

Reusable launch vehicle(RLV) has the typical characteristics of complex aerodynamic shape and propulsion system coupling, and the flight environment is highly complicated and intensely changeable. So its model has large uncertainty, which makes the nominal system quite different from the real system. Therefore, studying the influences caused by the uncertainties on the stability of the control system is of great significance for the controller design. In order to improve the performance of RLV, this paper proposes the approach of analyzing the influence of the model uncertainties. According to the typical RLV, the coupling dynamic and kinematics models are built. Then different factors that cause uncertainties during building the model are analyzed and summed up. After that, the model uncertainties are expressed according to the additive uncertainty model. Choosing the uncertainties matrix's maximum singular values as the boundary model, and selecting the uncertainties matrix's norm to show t how much the uncertainty factors influence is on the stability of the control system . The simulation results illustrate that the inertial factors have the largest influence on the stability of the system, and it is necessary and important to take the model uncertainties into consideration before the designing the controller of this kind of aircraft( like RLV, etc).

This "Is" Rocket Science!

Science.gov (United States)

Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

2013-01-01

Using model rockets to teach physics can be an effective way to engage students in learning. In this paper, we present a curriculum developed in response to an expressed need for helping high school students review physics equations in preparation for a state-mandated exam. This required a mode of teaching that was more advanced and analytical…

Corrosion Protection of Launch Infrastructure and Hardware Through the Space Shuttle Program

Science.gov (United States)

Calle, L. M.

2011-01-01

Corrosion, the environmentally induced degradation of materials, has been a challenging and costly problem that has affected NASA's launch operations since the inception of the Space Program. Corrosion studies began at NASA's Kennedy Space Center (KSC) in 1966 during the Gemini/Apollo Programs with the evaluation of long-term protective coatings for the atmospheric protection of carbon steel. NASA's KSC Beachside Corrosion Test Site, which has been documented by the American Society of Materials (ASM) as one of the most corrosive, naturally occurring environments in the world, was established at that time. With the introduction of the Space Shuttle in 1981, the already highly corrosive natural conditions at the launch pad were rendered even more severe by the acidic exhaust from the solid rocket boosters. In the years that followed, numerous efforts at KSC identified materials, coatings, and maintenance procedures for launch hardware and equipment exposed to the highly corrosiye environment at the launch pads. Knowledge on materials degradation, obtained by facing the highly corrosive conditions of the Space Shuttle launch environment, as well as limitations imposed by the environmental impact of corrosion control, have led researchers at NASA's Corrosion Technology Laboratory to establish a new technology development capability in the area of corrosion prevention, detection, and mitigation at KSC that is included as one of the "highest priority" technologies identified by NASA's integrated technology roadmap. A historical perspective highlighting the challenges encountered in protecting launch infrastructure and hardware from corrosion during the life of the Space Shuttle program and the new technological advances that have resulted from facing the unique and highly corrosive conditions of the Space Shuttle launch environment will be presented.

14 CFR 101.25 - Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets.

Science.gov (United States)

2010-01-01

... Power Rockets and Class 3-Advanced High Power Rockets. 101.25 Section 101.25 Aeronautics and Space... OPERATING RULES MOORED BALLOONS, KITES, AMATEUR ROCKETS AND UNMANNED FREE BALLOONS Amateur Rockets § 101.25 Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets. When operating...

Preliminary Report: DESiGN and Test Result of KSR-3 Rocket Magnetometers

Directory of Open Access Journals (Sweden)

Hyo-Min Kim

2000-12-01

Full Text Available The solar wind contributes to the formation of unique space environment called the Earth's magnetosphere by various interactions with the Earth's magnetic field. Thus the solar-terrestrial environment affects the Earth's magnetic field, which can be observed with an instrument for the magnetic field measurement, the magnetometer usually mounted on the rocket and the satellite and based on the ground observatory. The magnetometer is a useful instrument for the spacecraft attitude control as well as the Earth's magnetic field measurements for a scientific purpose. In this paper, we present the preliminary design and test results of the two onboard magnetometers of KARI's (Korea Aerospace Research Institute sounding rocket, KSR-3, which will be launched four times during the period of 2001-02. The KSR-3 magnetometers consist of the fluxgate magnetometer, MAG/AIM (Attitude Information Magnetometer for acquiring the rocket flight attitude information, and of the search-coil magnetometer, MAG/SIM (Scientific Investigation Magnetometer for the observation of the Earth's magnetic field fluctuations. With the MAG/AIM, the 3-axis attitude information can be acquired by the comparison of the resulting dc magnetic vector field with the IGRF (International Geomagnetic Reference Field. The Earth's magnetic field fluctuations ranging from 10 to 1,000 Hz can also be observed with the MAG/SIM measurement.

Structured waves near the plasma frequency observed in three auroral rocket flights

Directory of Open Access Journals (Sweden)

M. Samara

2006-11-01

Full Text Available We present observations of waves at and just above the plasma frequency (f_pe from three high frequency electric field experiments on three recent rockets launched to altitudes of 300–900 km in active aurora. The predominant observed HF waves just above f_pe are narrowband, short-lived emissions with amplitudes ranging from <1 mV/m to 20 mV/m, often associated with structured electron density. The nature of these HF waves, as determined from frequency-time spectrograms, is highly variable: in some cases, the frequency decreases monotonically with time as in the "HF-chirps" previously reported (McAdams and LaBelle, 1999, but in other cases rising frequencies are observed, or features which alternately rise and fall in frequency. They exhibit two timescales of amplitude variation: a short timescale, typically 50–100 ms, associated with individual discrete features, and a longer timescale associated with the general decrease in the amplitudes of the emissions as the rocket moves away from where the condition f~f_pe holds. The latter timescale ranges from 0.6 to 6.0 s, corresponding to distances of 2–7 km, assuming the phenomenon to be stationary and using the rocket velocity to convert time to distance.

Small Launch Vehicle Design Approaches: Clustered Cores Compared with Multi-Stage Inline Concepts

Science.gov (United States)

Waters, Eric D.; Beers, Benjamin; Esther, Elizabeth; Philips, Alan; Threet, Grady E., Jr.

2013-01-01

In an effort to better define small launch vehicle design options two approaches were investigated from the small launch vehicle trade space. The primary focus was to evaluate a clustered common core design against a purpose built inline vehicle. Both designs focused on liquid oxygen (LOX) and rocket propellant grade kerosene (RP-1) stages with the terminal stage later evaluated as a LOX/methane (CH4) stage. A series of performance optimization runs were done in order to minimize gross liftoff weight (GLOW) including alternative thrust levels, delivery altitude for payload, vehicle length to diameter ratio, alternative engine feed systems, re-evaluation of mass growth allowances, passive versus active guidance systems, and rail and tower launch methods. Additionally manufacturability, cost, and operations also play a large role in the benefits and detriments for each design. Presented here is the Advanced Concepts Office's Earth to Orbit Launch Team methodology and high level discussion of the performance trades and trends of both small launch vehicle solutions along with design philosophies that shaped both concepts. Without putting forth a decree stating one approach is better than the other; this discussion is meant to educate the community at large and let the reader determine which architecture is truly the most economical; since each path has such a unique set of limitations and potential payoffs.

NASA's Space Launch System Takes Shape

Science.gov (United States)

Askins, Bruce; Robinson, Kimberly F.

2017-01-01

Major hardware and software for NASA's Space Launch System (SLS) began rolling off assembly lines in 2016, setting the stage for critical testing in 2017 and the launch of a major new capability for deep space human exploration. SLS continues to pursue a 2018 first launch of Exploration Mission 1 (EM-1). At NASA's Michoud Assembly Facility near New Orleans, LA, Boeing completed welding of structural test and flight liquid hydrogen tanks, and engine sections. Test stands for core stage structural tests at NASA's Marshall Space Flight Center, Huntsville, AL. neared completion. The B2 test stand at NASA's Stennis Space Center, MS, completed major structural renovation to support core stage green run testing in 2018. Orbital ATK successfully test fired its second qualification solid rocket motor in the Utah desert and began casting the motor segments for EM-1. Aerojet Rocketdyne completed its series of test firings to adapt the heritage RS-25 engine to SLS performance requirements. Production is under way on the first five new engine controllers. NASA also signed a contract with Aerojet Rocketdyne for propulsion of the RL10 engines for the Exploration Upper Stage. United Launch Alliance delivered the structural test article for the Interim Cryogenic Propulsion Stage to MSFC for tests and construction was under way on the flight stage. Flight software testing at MSFC, including power quality and command and data handling, was completed. Substantial progress is planned for 2017. Liquid oxygen tank production will be completed at Michoud. Structural testing at Marshall will get under way. RS-25 hotfire testing will verify the new engine controllers. Core stage horizontal integration will begin. The core stage pathfinder mockup will arrive at the B2 test stand for fit checks and tests. EUS will complete preliminary design review. This paper will discuss the technical and programmatic successes and challenges of 2016 and look ahead to plans for 2017.

SLS Launched Missions Concept Studies for LUVOIR Mission

Science.gov (United States)

Stahl, H. Philip; Hopkins, Randall C.

2015-01-01

NASA's "Enduring Quests Daring Visions" report calls for an 8- to 16-meter Large UV-Optical-IR (LUVOIR) Surveyor mission to enable ultra-high-contrast spectroscopy and coronagraphy. AURA's "From Cosmic Birth to Living Earth" report calls for a 12-meter class High-Definition Space Telescope to pursue transformational scientific discoveries. The multi-center ATLAST Team is working to meet these needs. The MSFC Team is examining potential concepts that leverage the advantages of the SLS (Space Launch System). A key challenge is how to affordably get a large telescope into space. The JWST design was severely constrained by the mass and volume capacities of its launch vehicle. This problem is solved by using an SLS Block II-B rocket with its 10-m diameter x 30-m tall fairing and 45 mt payload to SE-L2. Previously, two development study cycles produced a detailed concept called ATLAST-8. Using ATLAST-8 as a point of departure, this paper reports on a new ATLAST-12 concept. ATLAST-12 is a 12-meter class segmented aperture LUVOIR with an 8-m class center segment. Thus, ATLAST-8 is now a de-scope option.

SLS launched missions concept studies for LUVOIR mission

Science.gov (United States)

Stahl, H. Philip; Hopkins, Randall C.

2015-09-01

NASA's "Enduring Quests Daring Visions" report calls for an 8- to 16-m Large UV-Optical-IR (LUVOIR) Surveyor mission to enable ultra-high-contrast spectroscopy and coronagraphy. AURA's "From Cosmic Birth to Living Earth" report calls for a 12-m class High-Definition Space Telescope to pursue transformational scientific discoveries. The multi-center ATLAST Team is working to meet these needs. The MSFC Team is examining potential concepts that leverage the advantages of the SLS (Space Launch System). A key challenge is how to affordably get a large telescope into space. The JWST design was severely constrained by the mass and volume capacities of its launch vehicle. This problem is solved by using an SLS Block II-B rocket with its 10-m diameter x 30-m tall fairing and estimated 45 mt payload to SE-L2. Previously, two development study cycles produced a detailed concept called ATLAST-8. Using ATLAST-8 as a point of departure, this paper reports on a new ATLAST-12 concept. ATLAST-12 is a 12-m class segmented aperture LUVOIR with an 8-m class center segment. Thus, ATLAST-8 is now a de-scope option.

Determination of the availability of appropriate aged flight rocket motors. [captive tests to determine case bond separation and grain bore cracking

Science.gov (United States)

Martin, P. J.

1974-01-01

A program to identify surplus solid rocket propellant engines which would be available for a program of functional integrity testing was conducted. The engines are classified as: (1) upper stage and apogee engines, (2) sounding rocket and launch vehicle engines, and (3) jato, sled, and tactical engines. Nearly all the engines were available because their age exceeds the warranted shelf life. The preference for testing included tests at nominal flight conditions, at design limits, and to establish margin limits. The principal failure modes of interest were case bond separation and grain bore cracking. Data concerning the identification and characteristics of each engine are tabulated. Methods for conducting the tests are described.

Simultaneous observations of E- and B-ULF waves aboard a sounding rocket payload

International Nuclear Information System (INIS)

Kloecker, N.; Luehr, H.; Grabowski, R.

1980-01-01

Magnetic and electric field variations in the frequency range of 0.5 to 4 Hz were made on a payload flown within the IMS sounding rocket campaign 'Substormphenomena'. The payload was launched into an auroral break-up. The waves show amplitudes up to 100 nT in B and 100 mV/m in E. Mutual correlation of B and E as well as correlation with electron precipitation are observed. The energy flux of the waves and the particles are equally directed and of the same order of magnitude. (Auth.)

Cambridge Rocketry Simulator – A Stochastic Six-Degrees-of-Freedom Rocket Flight Simulator

OpenAIRE

Eerland, Willem J.; Box, Simon; Sóbester, András

2017-01-01

The Cambridge Rocketry Simulator can be used to simulate the flight of unguided rockets for both design and operational applications. The software consists of three parts: The first part is a GUI that enables the user to design a rocket. The second part is a verified and peer-reviewed physics model that simulates the rocket flight. This includes a Monte Carlo wrapper to model the uncertainty in the rocket’s dynamics and the atmospheric conditions. The third part generates visualizations of th...

The FUSE satellite is encased in a canister before being moved to the Launch Pad.

Science.gov (United States)

1999-01-01

At Hangar AE, Cape Canaveral Air Station (CCAS), the last segment is lifted over the top of NASA's Far Ultraviolet Spectroscopic Explorer (FUSE) satellite already encased in a protective canister. The satellite will next be moved to Launch Pad 17A, CCAS, for its scheduled launch June 23 aboard a Boeing Delta II rocket. FUSE was developed by The Johns Hopkins University under contract to Goddard Space Flight Center, Greenbelt, Md., to investigate the origin and evolution of the lightest elements in the universe - hydrogen and deuterium. In addition, the FUSE satellite will examine the forces and process involved in the evolution of the galaxies, stars and planetary systems by investigating light in the far ultraviolet portion of the electromagnetic spectrum.

Combustion of metal agglomerates in a solid rocket core flow

Science.gov (United States)

Maggi, Filippo; Dossi, Stefano; DeLuca, Luigi T.

2013-12-01

The need for access to space may require the use of solid propellants. High thrust and density are appealing features for different applications, spanning from boosting phase to other service applications (separation, de-orbiting, orbit insertion). Aluminum is widely used as a fuel in composite solid rocket motors because metal oxidation increases enthalpy release in combustion chamber and grants higher specific impulse. Combustion process of metal particles is complex and involves aggregation, agglomeration and evolution of reacting particulate inside the core flow of the rocket. It is always stated that residence time should be enough in order to grant complete metal oxidation but agglomerate initial size, rocket grain geometry, burning rate, and other factors have to be reconsidered. New space missions may not require large rocket systems and metal combustion efficiency becomes potentially a key issue to understand whether solid propulsion embodies a viable solution or liquid/hybrid systems are better. A simple model for metal combustion is set up in this paper. Metal particles are represented as single drops trailed by the core flow and reacted according to Beckstead's model. The fluid dynamics is inviscid, incompressible, 1D. The paper presents parametric computations on ideal single-size particles as well as on experimental agglomerate populations as a function of operating rocket conditions and geometries.

Turbulent energy dissipation rates observed by Doppler MST Radar and by rocket-borne instruments during the MIDAS/MaCWAVE campaign 2002

Directory of Open Access Journals (Sweden)

N. Engler

2005-06-01

Full Text Available During the MIDAS/MaCWAVE campaign in summer 2002 we have observed turbulence using Doppler beam steering measurements obtained from the ALWIN VHF radar at Andøya/Northern Norway. This radar was operated in the Doppler beam steering mode for turbulence investigations during the campaign, as well as in spaced antenna mode, for continuously measuring the background wind field. The real-time data analysis of the Doppler radar backscattering provided the launch conditions for the sounding rockets. The spectral width data observed during the occurrence of PMSE were corrected for beam and shear broadening caused by the background wind field to obtain the turbulent part of the spectral width. The turbulent energy dissipation rates determined from the turbulent spectral width vary between 5 and 100mW kg-1 in the altitude range of 80-92km and increase with altitude. These estimations agree well with the in-situ measurements using the CONE sensor which was launched on 3 sounding rockets during the campaign.

Liquid Rocket Engine Testing

Science.gov (United States)

2016-10-21

Briefing Charts 3. DATES COVERED (From - To) 17 October 2016 – 26 October 2016 4. TITLE AND SUBTITLE Liquid Rocket Engine Testing 5a. CONTRACT NUMBER...298 (Rev. 8-98) Prescribed by ANSI Std. 239.18 Liquid Rocket Engine Testing SFTE Symposium 21 October 2016 Jake Robertson, Capt USAF AFRL...Distribution Unlimited. PA Clearance 16493 Liquid Rocket Engine Testing • Engines and their components are extensively static-tested in development • This

Theoretical Tools and Software for Modeling, Simulation and Control Design of Rocket Test Facilities

Science.gov (United States)

Richter, Hanz

2004-01-01

A rocket test stand and associated subsystems are complex devices whose operation requires that certain preparatory calculations be carried out before a test. In addition, real-time control calculations must be performed during the test, and further calculations are carried out after a test is completed. The latter may be required in order to evaluate if a particular test conformed to specifications. These calculations are used to set valve positions, pressure setpoints, control gains and other operating parameters so that a desired system behavior is obtained and the test can be successfully carried out. Currently, calculations are made in an ad-hoc fashion and involve trial-and-error procedures that may involve activating the system with the sole purpose of finding the correct parameter settings. The goals of this project are to develop mathematical models, control methodologies and associated simulation environments to provide a systematic and comprehensive prediction and real-time control capability. The models and controller designs are expected to be useful in two respects: 1) As a design tool, a model is the only way to determine the effects of design choices without building a prototype, which is, in the context of rocket test stands, impracticable; 2) As a prediction and tuning tool, a good model allows to set system parameters off-line, so that the expected system response conforms to specifications. This includes the setting of physical parameters, such as valve positions, and the configuration and tuning of any feedback controllers in the loop.

Heat Transfer by Thermo-capillary Convection -Sounding Rocket COMPERE Experiment SOURCE

Science.gov (United States)

Dreyer, Michael; Fuhrmann, Eckart

The sounding rocket COMPERE experiment SOURCE was successfully flown on MASER 11, launched in Kiruna (ESRANGE), May 15th, 2008. SOURCE has been intended to partly ful-fill the scientific objectives of the European Space Agency (ESA) Microgravity Applications Program (MAP) project AO-2004-111 (Convective boiling and condensation). Three parties of principle investigators have been involved to design the experiment set-up: ZARM for thermo-capillary flows, IMFT (Toulouse, France) for boiling studies, EADS Astrium (Bremen, Ger-many) for depressurization. The topic of this paper is to study the effect of wall heat flux on the contact line of the free liquid surface and to obtain a correlation for a convective heat trans-fer coefficient. The experiment has been conducted along a predefined time line. A preheating sequence at ground was the first operation to achieve a well defined temperature evolution within the test cell and its environment inside the rocket. Nearly one minute after launch, the pressurized test cell was filled with the test liquid HFE-7000 until a certain fill level was reached. Then the free surface could be observed for 120 s without distortion. Afterwards, the first depressurization was started to induce subcooled boiling, the second one to start saturated boiling. The data from the flight consists of video images and temperature measurements in the liquid, the solid, and the gaseous phase. Data analysis provides the surface shape versus time and the corresponding apparent contact angle. Computational analysis provides information for the determination of the heat transfer coefficient in a compensated gravity environment where a flow is caused by the temperature difference between the hot wall and the cold liquid. The paper will deliver correlations for the effective contact angle and the heat transfer coefficient as a function of the relevant dimensionsless parameters as well as physical explanations for the observed behavior. The data will be used

Distributed Web-Based Expert System for Launch Operations

Science.gov (United States)

Bardina, Jorge E.; Thirumalainambi, Rajkumar

2005-01-01

The simulation and modeling of launch operations is based on a representation of the organization of the operations suitable to experiment of the physical, procedural, software, hardware and psychological aspects of space flight operations. The virtual test bed consists of a weather expert system to advice on the effect of weather to the launch operations. It also simulates toxic gas dispersion model, and the risk impact on human health. Since all modeling and simulation is based on the internet, it could reduce the cost of operations of launch and range safety by conducting extensive research before a particular launch. Each model has an independent decision making module to derive the best decision for launch.

Contamination control requirements implementation for the James Webb Space Telescope (JWST), part 2: spacecraft, sunshield, observatory, and launch

Science.gov (United States)

Wooldridge, Eve M.; Schweiss, Andrea; Henderson-Nelson, Kelly; Woronowicz, Michael; Patel, Jignasha; Macias, Matthew; McGregor, R. Daniel; Farmer, Greg; Schmeitzky, Olivier; Jensen, Peter; Rumler, Peter; Romero, Beatriz; Breton, Jacques

2014-09-01

This paper will continue from Part 1 of JWST contamination control implementation. In addition to optics, instruments, and thermal vacuum testing, JWST also requires contamination control for a spacecraft that must be vented carefully in order to maintain solar array and thermal radiator thermal properties; a tennis court-sized sunshield made with 1-2 mil Kapton™ layers that must be manufactured and maintained clean; an observatory that must be integrated, stowed and transported to South America; and a rocket that typically launches commercial payloads without contamination sensitivity. An overview of plans developed to implement contamination control for the JWST spacecraft, sunshield, observatory and launch vehicle will be presented.

The model of beam-plasma discharge in the rocket environment during an electron beam injection in the ionosphere

International Nuclear Information System (INIS)

Mishin, E.V.; Ruzhin, Yu.Ya.

1980-01-01

The model of beam-plasma discharge in the rocket environment during electron beam injection in the ionosphere is constructed. The discharge plasma density dependence on the neutral gas concentration and the beam parameters is found

Development of small solid rocket boosters for the ILR-33 sounding rocket

Science.gov (United States)

Nowakowski, Pawel; Okninski, Adam; Pakosz, Michal; Cieslinski, Dawid; Bartkowiak, Bartosz; Wolanski, Piotr

2017-09-01

This paper gives an overview of the development of a 6000 Newton-class solid rocket motor for suborbital applications. The design configuration and results of interior ballistics calculations are given. The initial use of the motor as the main propulsion system of the H1 experimental in-flight test platform, within the Polish Small Sounding Rocket Program, is presented. Comparisons of theoretical and experimental performance are shown. Both on-ground and in-flight tests are discussed. A novel composite-case manufacturing technology, which enabled to reach high propellant mass fractions, was validated and significant cost-reductions were achieved. This paper focuses on the process of adapting the design for use as the booster stage of the ILR-33 sounding rocket, under development at the Institute of Aviation in Warsaw, Poland. Parallel use of two of the flight-proven rocket motors along with the main stage is planned. The process of adapting the rocket motor for booster application consists of stage integration, aerothermodynamics and reliability analyses. The separation mechanism and environmental impact are also discussed within this paper. Detailed performance analysis with focus on propellant grain geometry is provided. The evolution of the design since the first flights of the H1 rocket is covered and modifications of the manufacturing process are described. Issues of simultaneous ignition of two motors and their non-identical performance are discussed. Further applications and potential for future development are outlined. The presented results are based on the initial work done by the Rocketry Group of the Warsaw University of Technology Students' Space Association. The continuation of the Polish Small Sounding Rocket Program on a larger scale at the Institute of Aviation proves the value of the outcomes of the initial educational project.

Baking Soda and Vinegar Rockets

Science.gov (United States)

Claycomb, James R.; Zachary, Christopher; Tran, Quoc

2009-01-01

Rocket experiments demonstrating conservation of momentum will never fail to generate enthusiasm in undergraduate physics laboratories. In this paper, we describe tests on rockets from two vendors that combine baking soda and vinegar for propulsion. The experiment compared two analytical approximations for the maximum rocket height to the…

Computer model for the recombination zone of a microwave-plasma electrothermal rocket

Energy Technology Data Exchange (ETDEWEB)

Filpus, J.W.; Hawley, M.C.

1987-01-01

As part of a study of the microwave-plasma electrothermal rocket, a computer model of the flow regime below the plasma has been developed. A second-order model, including axial dispersion of energy and material and boundary conditions at infinite length, was developed to partially reproduce the absence of mass-flow rate dependence that was seen in experimental temperature profiles. To solve the equations of the model, a search technique was developed to find the initial derivatives. On integrating with a trial set of initial derivatives, the values and their derivatives were checked to judge whether the values were likely to attain values outside the practical regime, and hence, the boundary conditions at infinity were likely to be violated. Results are presented and directions for further development are suggested. 17 references.

Multi-Rocket Thought Experiment

Science.gov (United States)

Smarandache, Florentin

2014-03-01

We consider n>=2 identical rockets: R1 ,R2 , ..., Rn. Each of them moving at constant different velocities respectively v1 ,v2 , ..., vn on parallel directions in the same sense. In each rocket there is a light clock, the observer on earth also has a light clock. All n + 1 light clocks are identical and synchronized. The proper time Δt' in each rocket is the same. (1) If we consider the observer on earth and the first rocket R1, then the non-proper time Δt of the observer on earth is dilated with the factor D(v1) : or Δt = Δt' D(v1) (1) But if we consider the observer on earth and the second rocket R2 , then the non-proper time Δt of the observer on earth is dilated with a different factor D(v2) : or Δt = Δt' D(v2) And so on. Therefore simultaneously Δt is dilated with different factors D(v1) , D(v2), ..., D(vn) , which is a multiple contradiction.

Rocket Science 101 Interactive Educational Program

Science.gov (United States)

Armstrong, Dennis; Funkhouse, Deborah; DiMarzio, Donald

2007-01-01

To better educate the public on the basic design of NASA s current mission rockets, Rocket Science 101 software has been developed as an interactive program designed to retain a user s attention and to teach about basic rocket parts. This program also has helped to expand NASA's presence on the Web regarding educating the public about the Agency s goals and accomplishments. The software was designed using Macromedia s Flash 8. It allows the user to select which type of rocket they want to learn about, interact with the basic parts, assemble the parts to create the whole rocket, and then review the basic flight profile of the rocket they have built.

New high-resolution rocket-ultraviolet filtergrams of the solar disc

Science.gov (United States)

Foing, B.; Bonnet, R.-M.; Bruner, M.

1986-01-01

A rocket-borne solar ultraviolet telescope named Transition Region Camera was launched successfully for the third on July 13, 1982. High quality calibrated photographic images of the sun were obtained at Lyman alpha and in the continuum at 160 nm and 220 nm. The angular resolution achieved is better than one arcsec. A flare, active regions, sunspots, the 8 Mm mesostructure, the chromospheric network, bright UV grains and coronal loops were observed during the flight. The results are presented and the evolution with height in the solar atmosphere of the various structures observed is followed from one wavelength to the other, showing distinct differences. The value of the field's intensity of magnetic flux tubes is deduced from the observations.

Qualification of Coatings for Launch Facilities and Ground Support Equipment Through the NASA Corrosion Technology Laboratory

Science.gov (United States)

Kolody, Mark R.; Curran, Jerome P.; Calle, Luz Marina

2014-01-01

Corrosion protection at NASA's Kennedy Space Center is a high priority item. The launch facilities at the Kennedy Space Center are located approximately 1000 feet from the Atlantic Ocean where they are exposed to salt deposits, high humidity, high UV degradation, and acidic exhaust from solid rocket boosters. These assets are constructed from carbon steel, which requires a suitable coating to provide long-term protection to reduce corrosion and its associated costs.

Multi-Phase Modeling of Rainbird Water Injection

Science.gov (United States)

Vu, Bruce T.; Moss, Nicholas; Sampson, Zoe

2014-01-01

This paper describes the use of a Volume of Fluid (VOF) multiphase model to simulate the water injected from a rainbird nozzle used in the sound suppression system during launch. The simulations help determine the projectile motion for different water flow rates employed at the pad, as it is critical to know if water will splash on the first-stage rocket engine during liftoff.

Two phase flow combustion modelling of a ducted rocket

NARCIS (Netherlands)

Stowe, R.A.; Dubois, C.; Harris, P.G.; Mayer, A.E.H.J.; Champlain, A. de; Ringuette, S.

2001-01-01

Under a co-operative program, the Defence Research Establishment Valcartier and Université Laval in Canada and the TNO Prins Maurits Laboratory in the Netherlands have studied the use of a ducted rocket for missile propulsion. Hot-flow direct-connect combustion experiments using both simulated and

NASA's Space Launch System: Affordability for Sustainability

Science.gov (United States)

May, Todd A.; Creech, Stephen D.

2012-01-01

The National Aeronautics and Space Administration's (NASA) Space Launch System (SLS) Program, managed at the Marshall Space Flight Center, is charged with delivering a new capability for human exploration beyond Earth orbit in an austere economic climate. But the SLS value is clear and codified in United States (U.S.) budget law. The SLS Program knows that affordability is the key to sustainability and will provide an overview of initiatives designed to fit within the funding guidelines by using existing engine assets and hardware now in testing to meet a first launch by 2017 within the projected budget. It also has a long-range plan to keep the budget flat, yet evolve the 70-tonne (t) initial lift capability to 130-t lift capability after the first two flights. To achieve the evolved configuration, advanced technologies must offer appropriate return on investment to be selected through the competitive process. For context, the SLS will be larger than the Saturn V that took 12 men on 6 trips for a total of 11 days on the lunar surface some 40 years ago. Astronauts train for long-duration voyages on platforms such as the International Space Station, but have not had transportation to go beyond Earth orbit in modern times, until now. To arrive at the launch vehicle concept, the SLS Program conducted internal engineering and business studies that have been externally validated by industry and reviewed by independent assessment panels. In parallel with SLS concept studies, NASA is now refining its mission manifest, guided by U.S. space policy and the Global Exploration Roadmap, which reflects the mutual goals of a dozen member nations. This mission planning will converge with a flexible heavy-lift rocket that can carry international crews and the air, water, food, and equipment they need for extended trips to asteroids and Mars. In addition, the SLS capability will accommodate very large science instruments and other payloads, using a series of modular fairings and

Assessing Upper-Level Winds on Day-of-Launch

Science.gov (United States)

Bauman, William H., III; Wheeler, Mark M.

2012-01-01

On the day-or-launch. the 45th Weather Squadron Launch Weather Officers (LWOS) monitor the upper-level winds for their launch customers to include NASA's Launch Services Program (LSP). During launch operations, the payload launch team sometimes asks the LWO if they expect the upper level winds to change during the countdown but the LWOs did not have the capability to quickly retrieve or display the upper-level observations and compare them to the numerical weather prediction model point forecasts. The LWOs requested the Applied Meteorology Unit (AMU) develop a capability in the form of a graphical user interface (GUI) that would allow them to plot upper-level wind speed and direction observations from the Kennedy Space Center Doppler Radar Wind Profilers and Cape Canaveral Air Force Station rawinsondes and then overlay model point forecast profiles on the observation profiles to assess the performance of these models and graphically display them to the launch team. The AMU developed an Excel-based capability for the LWOs to assess the model forecast upper-level winds and compare them to observations. They did so by creating a GUI in Excel that allows the LWOs to first initialize the models by comparing the O-hour model forecasts to the observations and then to display model forecasts in 3-hour intervals from the current time through 12 hours.