Sample records for low-thrust resistojet reboost

  1. Selection of combined water electrolysis and resistojet propulsion for Space Station Freedom

    Schmidt, George R.


    An analytical rationale is presented for the configuration of the NASA Space Station's two-element propulsion system, and attention is given to the cost benefits accruing to this system over the Space Station's service life. The principal system element uses gaseous oxygen and hydrogen obtained through water electrolysis to furnish attitude control, backup attitude control, and contingency maneuvering. The secondary element uses resistojets to augment Space Station reboost through the acceleration of waste gases in the direction opposite the Station's flight path.

  2. Component Data Base for Space Station Resistojet Auxiliary Propulsion

    Bader, Clayton H.


    The resistojet was baselined for Space Station auxiliary propulsion because of its operational versatility, efficiency, and durability. This report was conceived as a guide to designers and planners of the Space Station auxiliary propulsion system. It is directed to the low thrust resistojet concept, though it should have application to other station concepts or systems such as the Environmental Control and Life Support System (ECLSS), Manufacturing and Technology Laboratory (MTL), and the Waste Fluid Management System (WFMS). The information will likely be quite useful in the same capacity for other non-Space Station systems including satellite, freeflyers, explorers, and maneuvering vehicles. The report is a catalog of the most useful information for the most significant feed system components and is organized for the greatest convenience of the user.

  3. Micro Resistojet for Small Satellites, Phase II

    National Aeronautics and Space Administration — Micro-resistojets offer an excellent combination of simplicity, performance and wet system mass for small satellites (<100 kg, <50 watts) requiring mN level...

  4. Micro Resistojet for Small Satellites, Phase I

    National Aeronautics and Space Administration — Micro-resistojets offer the best combination of simplicity, performance, wet system mass and power consumption for small satellites (<100kg, <50Watts)...

  5. Embodiment design for a multipropellant resistojet

    Bao, Johnny; Chilosi, Thierry; Goodwin, Jason; Mocio, Jim; Yeh, Bruce


    This document presents the design of a multipropellant resistojet to use as an auxiliary propulsion system on the Space Station. Such a system is necessary to counteract atmospheric drag effects encountered by the Station in its orbit. NASA specifications are strictly followed with emphasis on reliability, operating life, multipropellant capability, and exhaust emission control. Several design variants are considered, and the final solution is a resistojet with an electronic pressure regulator, variable control, an internal flow heater, and a conical nozzle. To construct the resistojet, the important components are resolved independently and then integrated with secondary units. The document also includes engineering drawings of the final design with assembly instructions. Before final utilization, a prototype testing is recommended to uncover possible problems.

  6. Shaping low-thrust trajectories with thrust-handling feature

    Taheri, Ehsan; Kolmanovsky, Ilya; Atkins, Ella


    Shape-based methods are becoming popular in low-thrust trajectory optimization due to their fast computation speeds. In existing shape-based methods constraints are treated at the acceleration level but not at the thrust level. These two constraint types are not equivalent since spacecraft mass decreases over time as fuel is expended. This paper develops a shape-based method based on a Fourier series approximation that is capable of representing trajectories defined in spherical coordinates and that enforces thrust constraints. An objective function can be incorporated to minimize overall mission cost, i.e., achieve minimum ΔV . A representative mission from Earth to Mars is studied. The proposed Fourier series technique is demonstrated capable of generating feasible and near-optimal trajectories. These attributes can facilitate future low-thrust mission designs where different trajectory alternatives must be rapidly constructed and evaluated.

  7. Optimization of Low-Thrust Spiral Trajectories by Collocation

    Falck, Robert D.; Dankanich, John W.


    As NASA examines potential missions in the post space shuttle era, there has been a renewed interest in low-thrust electric propulsion for both crewed and uncrewed missions. While much progress has been made in the field of software for the optimization of low-thrust trajectories, many of the tools utilize higher-fidelity methods which, while excellent, result in extremely high run-times and poor convergence when dealing with planetocentric spiraling trajectories deep within a gravity well. Conversely, faster tools like SEPSPOT provide a reasonable solution but typically fail to account for other forces such as third-body gravitation, aerodynamic drag, solar radiation pressure. SEPSPOT is further constrained by its solution method, which may require a very good guess to yield a converged optimal solution. Here the authors have developed an approach using collocation intended to provide solution times comparable to those given by SEPSPOT while allowing for greater robustness and extensible force models.

  8. Orbital and angular motion construction for low thrust interplanetary flight

    Yelnikov, R. V.; Mashtakov, Y. V.; Ovchinnikov, M. Yu.; Tkachev, S. S.


    Low thrust interplanetary flight is considered. Firstly, the fuel-optimal control is found. Then the angular motion is synthesized. This motion provides the thruster tracking of the required by optimal control direction. And, finally, reaction wheel control law for tracking this angular motion is proposed and implemented. The numerical example is given and total operation time for thrusters is found. Disturbances from solar pressure, thrust eccentricity, inaccuracy of reaction wheels installation and errors of inertia tensor are taken into account.

  9. Automated low-thrust guidance for the orbital maneuvering vehicle

    Rose, Richard E.; Schmeichel, Harry; Shortwell, Charles P.; Werner, Ronald A.


    This paper describes the highly autonomous OMV Guidance Navigation and Control system. Emphasis is placed on a key feature of the design, the low thrust guidance algorithm. The two guidance modes, orbit change guidance and rendezvous guidance, are discussed in detail. It is shown how OMV will automatically transfer from its initial orbit to an arbitrary target orbit and reach a specified rendezvous position relative to the target vehicle.

  10. The cislunar low-thrust trajectories via the libration point

    Qu, Qingyu; Xu, Ming; Peng, Kun


    The low-thrust propulsion will be one of the most important propulsion in the future due to its large specific impulse. Different from traditional low-thrust trajectories (LTTs) yielded by some optimization algorithms, the gradient-based design methodology is investigated for LTTs in this paper with the help of invariant manifolds of LL1 point and Halo orbit near the LL1 point. Their deformations under solar gravitational perturbation are also presented to design LTTs in the restricted four-body model. The perturbed manifolds of LL1 point and its Halo orbit serve as the free-flight phase to reduce the fuel consumptions as much as possible. An open-loop control law is proposed, which is used to guide the spacecraft escaping from Earth or captured by Moon. By using a two-dimensional search strategy, the ON/OFF time of the low-thrust engine in the Earth-escaping and Moon-captured phases can be obtained. The numerical implementations show that the LTTs achieved in this paper are consistent with the one adopted by the SMART-1 mission.

  11. Global Optimization of Low-Thrust Interplanetary Trajectories Subject to Operational Constraints

    Englander, Jacob Aldo; Vavrina, Matthew; Hinckley, David


    Low-thrust electric propulsion provides many advantages for mission to difficult targets-Comets and asteroids-Mercury-Outer planets (with sufficient power supply)Low-thrust electric propulsion is characterized by high power requirements but also very high specific impulse (Isp), leading to very good mass fractions. Low-thrust trajectory design is a very different process from chemical trajectory.

  12. Minimum Propellant Low-Thrust Maneuvers near the Libration Points

    Marinescu, A.; Dumitrache, M.

    The impulse technique certainly can bring the vehicle on orbits around the libration points or close to them. The question that aries is, by what means can the vehicle arrive in such cases at the libration points? A first investigation carried out in this paper can give an answer: the use of the technique of low-thrust, which, in addition, can bring the vehicle from the libration points near to or into orbits around these points. This aspect is considered in this present paper where for the applications we have considered the transfer for orbits of the equidistant point L4 and of the collinear point L2, from Earth-moon system. This transfer maneuver can be used to insertion one satellite on libration points orbits. In Earth- moon system the points L 4 and L 5 because an vehicle in on of the equidistant points in quite stable and remains in its vicinity of perturbed, have potential interest for the establishment of transporder satellite for interplanetary tracking. In contrast an vehicle in one of the collinear points is quite instable and it will oscillate along the Earth-moon-axis at increasing amplitude and gradually escape from the libration point. Let use assume that a space vehicle equipped with a low-thrust propulsion is near a libration point L. We consider the planar motion in the restricted frame of the three bodies in the rotating system L, where the Earth-moon distance D=l. The unit of time T is period of the moon's orbit divided by 2 and multiplied by the square root of the quantity one plus the moon/Earth mass ratio, and the unit of mass is the Earth's mass. With these predictions the motion equatios of the vehicle equiped with a low-thrust propulsion installation in the linear approximation near the libration point, have been established. The parameters of the motion at the beginning and the end of these maneuvers are known, the variational problem has been formulated as a Lagrange type problem with fixed extremities. On established the differential

  13. A new method for optimization of low-thrust gravity-assist sequences

    Maiwald, V.


    Recently missions like Hayabusa and Dawn have shown the relevance and benefits of low-thrust spacecraft concerning the exploration of our solar system. In general, the efficiency of low-thrust propulsion is one means of improving mission payload mass. At the same time, gravity-assist maneuvers can serve as mission enablers, as they have the capability to provide "free energy." A combination of both, gravity-assist and low-thrust propulsion, has the potential to generally improve mission performance, i.e. planning and optimization of gravity-assist sequences for low-thrust missions is a desirable asset. Currently no established methods exist to include the gravity-assist partners as optimization variable for low-thrust missions. The present paper explains how gravity-assists are planned and optimized, including the gravity-assist partners, for high-thrust missions and discusses the possibility to transfer the established method, based on the Tisserand Criterion, to low-thrust missions. It is shown how the Tisserand Criterion needs to be adapted using a correction term for the low-thrust situation. It is explained why this necessary correction term excludes an a priori evaluation of sequences and therefore their planning and an alternate approach is proposed. Preliminary results of this method, by application of a Differential Evolution optimization algorithm, are presented and discussed, showing that the method is valid but can be improved. Two constraints on the search space are briefly presented for that aim.

  14. Trajectory Design in a Combined Low-Thrust Multi-Body Environment

    National Aeronautics and Space Administration — Low-thrust propulsion (e.g. electric propulsion, solar sailing, etc.) is a promising new technology that may allow smaller, more efficient spacecraft to participate...

  15. Sensitivity Analysis and Mitigation with Applications to Ballistic and Low-thrust Trajectory Design

    Alizadeh, Iman

    The ever increasing desire to expand space mission capabilities within the limited budgets of space industries requires new approaches to the old problem of spacecraft trajectory design. For example, recent initiatives for space exploration involve developing new tools to design low-cost, fail-safe trajectories to visit several potential destinations beyond our celestial neighborhood such as Jupiter's moons, asteroids, etc. Designing and navigating spacecraft trajectories to reach these destinations safely are complex and challenging. In particular, fundamental questions of orbital stability imposed by planetary protection requirements are not easily taken into account by standard optimal control schemes. The event of temporary engine loss or an unexpected missed thrust can indeed quickly lead to impact with planetary bodies or other unrecoverable trajectories. While electric propulsion technology provides superior efficiency compared to chemical engines, the very low-control authority and engine performance degradation can impose higher risk to the mission in strongly perturbed orbital environments. The risk is due to the complex gravitational field and its associated chaotic dynamics which causes large navigation dispersions in a short time if left un-controlled. Moreover, in these situations it can be outside the low-thrust propulsion system capability to correct the spacecraft trajectory in a reasonable time frame. These concerns can lead to complete or partial mission failure or even an infeasible mission concept at the early design stage. The goal of this research is to assess and increase orbital stability of ballistic and low-thrust transfer trajectories in multi-body systems. In particular, novel techniques are presented to characterize sensitivity and improve recovery characteristics of ballistic and low-thrust trajectories in unstable orbital environments. The techniques developed are based on perturbation analysis around ballistic trajectories to

  16. A Numerical Method to Generate Reference Trajectories for Optimization Methods to Support Low-Thrust Mission Design

    National Aeronautics and Space Administration — The recent success of missions employing low thrust propulsion systems has demonstrated the promise this technology holds for a wide array of future applications,...

  17. Search for hadronic events with low thrust and an isolated lepton with the CELLO detector

    Kroha, H.


    Using the CELLO detector at the PETRA e + e - -storage ring, a search for hadronic events with low thrust and an isolated lepton has been carried out. The rate of such events and the thrust distribution of inclusive lepton events show no evidence for a threshold of new sources of inclusive lepton events at the highest PETRA energies. The CELLO data thus do not support the observations of the MARK-J and JADE collaborations of an excess of muon events of this type at √s > 46.3 GeV. (orig.)

  18. Modeling low-thrust transfers between periodic orbits about five libration points: Manifolds and hierarchical design

    Zeng, Hao; Zhang, Jingrui


    The low-thrust version of the fuel-optimal transfers between periodic orbits with different energies in the vicinity of five libration points is exploited deeply in the Circular Restricted Three-Body Problem. Indirect optimization technique incorporated with constraint gradients is employed to further improve the computational efficiency and accuracy of the algorithm. The required optimal thrust magnitude and direction can be determined to create the bridging trajectory that connects the invariant manifolds. A hierarchical design strategy dividing the constraint set is proposed to seek the optimal solution when the problem cannot be solved directly. Meanwhile, the solution procedure and the value ranges of used variables are summarized. To highlight the effectivity of the transfer scheme and aim at different types of libration point orbits, transfer trajectories between some sample orbits, including Lyapunov orbits, planar orbits, halo orbits, axial orbits, vertical orbits and butterfly orbits for collinear and triangular libration points, are investigated with various time of flight. Numerical results show that the fuel consumption varies from a few kilograms to tens of kilograms, related to the locations and the types of mission orbits as well as the corresponding invariant manifold structures, and indicates that the low-thrust transfers may be a beneficial option for the extended science missions around different libration points.

  19. Heater Chip with Different Microchannels Geometries for a Low Pressure Free Molecular Micro-Resistojet

    Cordeiro Guerrieri, D.; de Athayde Costa e Silva, M.; Zandbergen, B.T.C.; Cervone, A.


    This paper presents a dynamic system approach for the modeling of fluid flow in microchannels to be used in thrust control applications. A micro-resistojet fabricated using MEMS (Microelectromechanical Systems) technology has been selected for the analysis. The device operates by vaporizing a liquid

  20. Combined high and low-thrust geostationary orbit insertion with radiation constraint

    Macdonald, Malcolm; Owens, Steven Robert


    The sequential use of an electric propulsion system is considered in combination with a high-thrust propulsion system for application to the propellant-optimal Geostationary Orbit insertion problem, whilst considering both temporal and radiation flux constraints. Such usage is found to offer a combined propellant mass saving when compared with an equivalent high-thrust only transfer. This propellant mass saving is seen to increase as the allowable transfer duration is increased, and as the thrust from the low-thrust system is increased, assuming constant specific impulse. It was found that the required plane change maneuver is most propellant-efficiently performed by the high-thrust system. The propellant optimal trajectory incurs a significantly increased electron flux when compared to an equivalent high-thrust only transfer. However, the electron flux can be reduced to a similar order of magnitude by increasing the high-thrust propellant consumption, whilst still delivering an improved mass fraction.

  1. A Numerical-Analytical Approach Based on Canonical Transformations for Computing Optimal Low-Thrust Transfers

    da Silva Fernandes, S.; das Chagas Carvalho, F.; Bateli Romão, J. V.


    A numerical-analytical procedure based on infinitesimal canonical transformations is developed for computing optimal time-fixed low-thrust limited power transfers (no rendezvous) between coplanar orbits with small eccentricities in an inverse-square force field. The optimization problem is formulated as a Mayer problem with a set of non-singular orbital elements as state variables. Second order terms in eccentricity are considered in the development of the maximum Hamiltonian describing the optimal trajectories. The two-point boundary value problem of going from an initial orbit to a final orbit is solved by means of a two-stage Newton-Raphson algorithm which uses an infinitesimal canonical transformation. Numerical results are presented for some transfers between circular orbits with moderate radius ratio, including a preliminary analysis of Earth-Mars and Earth-Venus missions.

  2. Coupled Low-thrust Trajectory and System Optimization via Multi-Objective Hybrid Optimal Control

    Vavrina, Matthew A.; Englander, Jacob Aldo; Ghosh, Alexander R.


    The optimization of low-thrust trajectories is tightly coupled with the spacecraft hardware. Trading trajectory characteristics with system parameters ton identify viable solutions and determine mission sensitivities across discrete hardware configurations is labor intensive. Local independent optimization runs can sample the design space, but a global exploration that resolves the relationships between the system variables across multiple objectives enables a full mapping of the optimal solution space. A multi-objective, hybrid optimal control algorithm is formulated using a multi-objective genetic algorithm as an outer loop systems optimizer around a global trajectory optimizer. The coupled problem is solved simultaneously to generate Pareto-optimal solutions in a single execution. The automated approach is demonstrated on two boulder return missions.

  3. Low-Thrust Orbital Transfers in the Two-Body Problem

    A. A. Sukhanov


    Full Text Available Low-thrust transfers between given orbits within the two-body problem are considered; the thrust is assumed power limited. A simple method for obtaining the transfer trajectories based on the linearization of the motion near reference orbits is suggested. Required calculation accuracy can be reached by means of use of a proper number of the reference orbits. The method may be used in the case of a large number of the orbits around the attracting center; no averaging is necessary in this case. The suggested method also is applicable to the cases of partly given final orbit and if there are constraints on the thrust direction. The method gives an optimal solution to the linearized problem which is not optimal for the original nonlinear problem; the difference between the optimal solutions to the original and linearized problems is estimated using a numerical example. Also examples illustrating the method capacities are given.

  4. A search for hadronic events with low thrust and an isolated lepton

    Behrend, H.J.; Buerger, J.; Criegee, L.; Dainton, J.B.; Fenner, H.; Field, J.H.; Franke, G.; Meyer, J.; Schroeder, V.; Timm, U.; Winter, G.G.; Zimmermann, W.; Bussey, P.J.; Buttar, C.; Campbell, A.J.; Hendry, D.; McCurrach, G.; Scarr, J.M.; Skillicorn, I.O.; Smith, K.M.; Ahme, J.; Blobel, V.; Feindt, M.; Harjes, J.; Peters, J.H.; Poppe, M.; Spitzer, H.; Boer, W. de; Buschhorn, G.; Grindhammer, G.; Gunderson, B.; Kiesling, C.; Kotthaus, R.; Kroha, H.; Lueers, D.; Oberlack, H.; Schacht, P.; Shooshtari, G.; Wiedenmann, W.; Davier, M.; Grivaz, J.F.; Haissinski, J.; Janot, P.; Journe, V.; Le Diberder, F.; Spadafora, A.; Veillet, J.J.; Cozzika, G.; Ducros, Y.; Lavagne, Y.; Ould Saada, F.; Zacek, J.; Alexander, G.; Bella, G.; Gnat, Y.; Grunhaus, J.; Klatchko, A.; Levy, A.


    Using the CELLO detector at the PETRA e + e - storage ring a search for hadronic events with low thrust and an isolated lepton has been carried out. The rate of such events and the thrust distribution of inclusive lepton events show no evidence for a threshold of new sources of inclusive lepton events at the highest PETRA energies. For √s > 46.3 GeV we observe one muon event of the above type, compared to an expectation of 0.8±0.2 events from lower energy data. We do not observe any electron event of this class at the highest PETRA energies. The CELLO data thus do not support the observations of the MARK-J and JADE collaborations of an excess of muon events of the above type at √s > 46.3 GeV. (orig.)

  5. An Integrated Tool for Low Thrust Optimal Control Orbit Transfers in Interplanetary Trajectories

    Dargent, T.; Martinot, V.

    In the last recent years a significant progress has been made in optimal control orbit transfers using low thrust electrical propulsion for interplanetary missions. The system objective is always the same: decrease the transfer duration and increase the useful satellite mass. The optimum control strategy to perform the minimum time to orbit or the minimum fuel consumption requires the use of sophisticated mathematical tools, most of the time dedicated to a specific mission and therefore hardly reusable. To improve this situation and enable Alcatel Space to perform rather quick trajectory design as requested by mission analysis, we have developed a software tool T-3D dedicated to optimal control orbit transfers which integrates various initial and terminal rendezvous conditions - e.g. fixed arrival time for planet encounter - and engine thrust profiles -e.g. thrust law variation with respect to the distance to the Sun -. This single and quite versatile tool allows to perform analyses like minimum consumption for orbit insertions around a planet from an hyperbolic trajectory, interplanetary orbit transfers, low thrust minimum time multiple revolution orbit transfers, etc… From a mathematical point of view, the software relies on the minimum principle formulation to find the necessary conditions of optimality. The satellite dynamics is a two body model and relies of an equinoctial formulation of the Gauss equation. This choice has been made for numerical purpose and to solve more quickly the two point boundaries values problem. In order to handle the classical problem of co-state variables initialization, problems simpler than the actual one can be solved straight forward by the tool and the values of the co-state variables are kept as first guess for a more complex problem. Finally, a synthesis of the test cases is presented to illustrate the capacities of the tool, mixing examples of interplanetary mission, orbit insertion, multiple revolution orbit transfers

  6. Green micro-resistojet research at Delft University of Technology: new options for Cubesat propulsion

    Cervone, A.; Zandbergen, B.; Guerrieri, D. C.; De Athayde Costa e Silva, M.; Krusharev, I.; van Zeijl, H.


    The aerospace industry is recently expressing a growing interest in green, safe and non-toxic propellants for the propulsion systems of the new generation of space vehicles, which is especially true in the case of Cubesat micro-propulsion systems. Demanding requirements are associated to the future missions and challenges offered by this class of spacecraft, where the availability of a propulsion system might open new possibilities for a wide range of applications including orbital maintenance and transfer, formation flying and attitude control. To accomplish these requirements, Delft University of Technology is currently developing two different concepts of water-propelled micro-thrusters based on MEMS technologies: a free molecular micro-resistojet operating with sublimating solid water (ice) at low plenum gas pressure of less than 600 Pa, and a more conventional micro-resistojet operating with liquid water heated and vaporized by means of a custom designed silicon heating chamber. In this status review paper, the current design and future expected developments of the two micro-propulsion concepts is presented and discussed, together with an initial analysis of the expected performance and potential operational issues. Results of numerical simulations conducted to optimize the design of the heating and expansion slots, as well as a detailed description of the manufacturing steps for the conventional micro-resistojet concept, are presented. Some intended steps for future research activities, including options for thrust intensity and direction control, are briefly introduced.

  7. Fuel optimization for low-thrust Earth-Moon transfer via indirect optimal control

    Pérez-Palau, Daniel; Epenoy, Richard


    The problem of designing low-energy transfers between the Earth and the Moon has attracted recently a major interest from the scientific community. In this paper, an indirect optimal control approach is used to determine minimum-fuel low-thrust transfers between a low Earth orbit and a Lunar orbit in the Sun-Earth-Moon Bicircular Restricted Four-Body Problem. First, the optimal control problem is formulated and its necessary optimality conditions are derived from Pontryagin's Maximum Principle. Then, two different solution methods are proposed to overcome the numerical difficulties arising from the huge sensitivity of the problem's state and costate equations. The first one consists in the use of continuation techniques. The second one is based on a massive exploration of the set of unknown variables appearing in the optimality conditions. The dimension of the search space is reduced by considering adapted variables leading to a reduction of the computational time. The trajectories found are classified in several families according to their shape, transfer duration and fuel expenditure. Finally, an analysis based on the dynamical structure provided by the invariant manifolds of the two underlying Circular Restricted Three-Body Problems, Earth-Moon and Sun-Earth is presented leading to a physical interpretation of the different families of trajectories.

  8. Low-Thrust Out-of-Plane Orbital Station-Keeping Maneuvers for Satellites

    Vivian M. Gomes


    Full Text Available This paper considers the problem of out of plane orbital maneuvers for station keeping of satellites. The main idea is to consider that a satellite is in an orbit around the Earth and that it has its orbit is disturbed by one or more forces. Then, it is necessary to perform a small amplitude orbital correction to return the satellite to its original orbit, to keep it performing its mission. A low thrust propulsion is used to complete this task. It is important to search for solutions that minimize the fuel consumption to increase the lifetime of the satellite. To solve this problem a hybrid optimal control approach is used. The accuracy of the satisfaction of the constraints is considered, in order to try to decrease the fuel expenditure by taking advantage of this freedom. This type of problem presents numerical difficulties and it is necessary to adjust parameters, as well as details of the algorithm, to get convergence. In this versions of the algorithm that works well for planar maneuvers are usually not adequate for the out of plane orbital corrections. In order to illustrate the method, some numerical results are presented.

  9. Anomalous Thrust Production from an RF Test Device Measured on a Low-Thrust Torsion Pendulum

    Brady, David; White, Harold G.; March, Paul; Lawrence, James T.; Davies, Frank J.


    This paper describes the eight-day August 2013 test campaign designed to investigate and demonstrate viability of using classical magnetoplasmadynamics to obtain a propulsive momentum transfer via the quantum vacuum virtual plasma. This paper will not address the physics of the quantum vacuum plasma thruster, but instead will describe the test integration, test operations, and the results obtained from the test campaign. Approximately 30-50 micro-Newtons of thrust were recorded from an electric propulsion test article consisting primarily of a radio frequency (RF) resonant cavity excited at approximately 935 megahertz. Testing was performed on a low-thrust torsion pendulum that is capable of detecting force at a single-digit micronewton level, within a stainless steel vacuum chamber with the door closed but at ambient atmospheric pressure. Several different test configurations were used, including two different test articles as well as a reversal of the test article orientation. In addition, the test article was replaced by an RF load to verify that the force was not being generated by effects not associated with the test article. The two test articles were designed by Cannae LLC of Doylestown, Pennsylvania. The torsion pendulum was designed, built, and operated by Eagleworks Laboratories at the NASA Johnson Space Center of Houston, Texas. Approximately six days of test integration were required, followed by two days of test operations, during which, technical issues were discovered and resolved. Integration of the two test articles and their supporting equipment was performed in an iterative fashion between the test bench and the vacuum chamber. In other words, the test article was tested on the bench, then moved to the chamber, then moved back as needed to resolve issues. Manual frequency control was required throughout the test. Thrust was observed on both test articles, even though one of the test articles was designed with the expectation that it would not

  10. Numerical and Analytical Study of Optimal Low-Thrust Limited-Power Transfers between Close Circular Coplanar Orbits

    Sandro da Silva Fernandes


    Full Text Available A numerical and analytical study of optimal low-thrust limited-power trajectories for simple transfer (no rendezvous between close circular coplanar orbits in an inverse-square force field is presented. The numerical study is carried out by means of an indirect approach of the optimization problem in which the two-point boundary value problem, obtained from the set of necessary conditions describing the optimal solutions, is solved through a neighboring extremal algorithm based on the solution of the linearized two-point boundary value problem through Riccati transformation. The analytical study is provided by a linear theory which is expressed in terms of nonsingular elements and is determined through the canonical transformation theory. The fuel consumption is taken as the performance criterion and the analysis is carried out considering various radius ratios and transfer durations. The results are compared to the ones provided by a numerical method based on gradient techniques.

  11. A Time-Regularized, Multiple Gravity-Assist Low-Thrust, Bounded-Impulse Model for Trajectory Optimization

    Ellison, Donald H.; Englander, Jacob A.; Conway, Bruce A.


    The multiple gravity assist low-thrust (MGALT) trajectory model combines the medium-fidelity Sims-Flanagan bounded-impulse transcription with a patched-conics flyby model and is an important tool for preliminary trajectory design. While this model features fast state propagation via Keplers equation and provides a pleasingly accurate estimation of the total mass budget for the eventual flight suitable integrated trajectory it does suffer from one major drawback, namely its temporal spacing of the control nodes. We introduce a variant of the MGALT transcription that utilizes the generalized anomaly from the universal formulation of Keplers equation as a decision variable in addition to the trajectory phase propagation time. This results in two improvements over the traditional model. The first is that the maneuver locations are equally spaced in generalized anomaly about the orbit rather than time. The second is that the Kepler propagator now has the generalized anomaly as its independent variable instead of time and thus becomes an iteration-free propagation method. The new algorithm is outlined, including the impact that this has on the computation of Jacobian entries for numerical optimization, and a motivating application problem is presented that illustrates the improvements that this model has over the traditional MGALT transcription.

  12. Some Calculated Research Results of the Working Process Parameters of the Low Thrust Rocket Engine Operating on Gaseous Oxygen-Hydrogen Fuel

    Ryzhkov, V.; Morozov, I.


    The paper presents the calculating results of the combustion products parameters in the tract of the low thrust rocket engine with thrust P ∼ 100 N. The article contains the following data: streamlines, distribution of total temperature parameter in the longitudinal section of the engine chamber, static temperature distribution in the cross section of the engine chamber, velocity distribution of the combustion products in the outlet section of the engine nozzle, static temperature near the inner wall of the engine. The presented parameters allow to estimate the efficiency of the mixture formation processes, flow of combustion products in the engine chamber and to estimate the thermal state of the structure.

  13. A Complete First-Order Analytical Solution for Optimal Low-Thrust Limited-Power Transfers Between Coplanar Orbits with Small Eccentricities

    Da Silva Fernandes, Sandro; Das Chagas Carvalho, Francisco; Vilhena de Moraes, Rodolpho

    The purpose of this work is to present a complete first order analytical solution, which includes short periodic terms, for the problem of optimal low-thrust limited power trajectories with large amplitude transfers (no rendezvous) between coplanar orbits with small eccentricities in Newtonian central gravity field. The study of these transfers is particularly interesting because the orbits found in practice often have a small eccentricity and the problem of transferring a vehicle from a low earth orbit to a high earth orbit is frequently found. Besides, the analysis has been motivated by the renewed interest in the use of low-thrust propulsion systems in space missions verified in the last two decades. Several researchers have obtained numerical and sometimes analytical solutions for a number of specific initial orbits and specific thrust profiles. Averaging methods are also used in such researches. Firstly, the optimization problem associated to the space transfer problem is formulated as a Mayer problem of optimal control with Cartesian elements - position and velocity vectors - as state variables. After applying the Pontryagin Maximum Principle, successive Mathieu transformations are performed and suitable sets of orbital elements are introduced. The short periodic terms are eliminated from the maximum Hamiltonian function through an infinitesimal canonical transformation built through Hori method - a perturbation canonical method based on Lie series. The new Hamiltonian function, which results from the infinitesimal canonical transformation, describes the extremal trajectories for long duration maneuvers. Closed-form analytical solutions are obtained for the new canonical system by solving the Hamilton-Jacobi equation through the separation of variables technique. By applying the transformation equations of the algorithm of Hori method, a first order analytical solution for the problem is obtained in non-singular orbital elements. For long duration maneuvers

  14. A Numerical Study of Low-Thrust Limited Power Trajectories between Coplanar Circular Orbits in an Inverse-Square Force Field

    Sandro da Silva Fernandes


    Full Text Available A numerical study of optimal low-thrust limited power trajectories for simple transfer (no rendezvous between circular coplanar orbits in an inverse-square force field is performed by two different classes of algorithms in optimization of trajectories. This study is carried out by means of a direct method based on gradient techniques and by an indirect method based on the second variation theory. The direct approach of the trajectory optimization problem combines the main positive characteristics of two well-known direct methods in optimization of trajectories: the steepest-descent (first-order gradient method and a direct second variation (second-order gradient method. On the other hand, the indirect approach of the trajectory optimization problem involves two different algorithms of the well-known neighboring extremals method. Several radius ratios and transfer durations are considered, and the fuel consumption is taken as the performance criterion. For small-amplitude transfers, the results are compared to the ones provided by a linear analytical theory.

  15. The Thermal State Computational Research of the Low-Thrust Oxygen-Methane Gaseous-Propellant Rocket Engine in the Pulse Mode of Operation

    O. A. Vorozheeva


    Full Text Available Currently promising development direction of space propulsion engineering is to use, as spacecraft controls, low-thrust rocket engines (RDTM on clean fuels, such as oxygen-methane. Modern RDTM are characterized by a lack regenerative cooling and pulse mode of operation, during which there is accumulation of heat energy to lead to the high thermal stress of RDTM structural elements. To get an idea about the thermal state of its elements, which further will reduce the number of fire tests is therefore necessary in the development phase of a new product. Accordingly, the aim of this work is the mathematical modeling and computational study of the thermal state of gaseous oxygen-methane propellant RDMT operating in pulse mode.In this paper we consider a model RDTM working on gaseous propellants oxygen-methane in pulse mode.To calculate the temperature field of the chamber wall of model RDMT under consideration is used the mathematical model of non-stationary heat conduction in a two-dimensional axisymmetric formulation that takes into account both the axial heat leakages and the nonstationary processes occurring inside the chamber during pulse operation of RDMT.As a result of numerical study of the thermal state of model RDMT, are obtained the temperature fields during engine operation based on convective, conductive, and radiative mechanisms of heat transfer from the combustion products to the wall.It is shown that the elements of flanges of combustion chamber of model RDMT act as heat sinks structural elements. Temperatures in the wall of the combustion chamber during the engine mode of operation are considered relatively low.Raised temperatures can also occur in the mixing head in the feeding area of the oxidant into the combustion chamber.During engine operation in the area forming the critical section, there is an intensive heating of a wall, which can result in its melting, which in turn will increase the minimum nozzle throat area and hence

  16. Low Thrust Trajectory Design for GSFC Missions

    National Aeronautics and Space Administration — The Evolutionary Mission Trajectory Generator (EMTG) is a global trajectory optimization tool. EMTG is intended for use in designing interplanetary missions which...

  17. Low-Thrust Bipropellant Engine Technology.


    Non-Destructive Testing OD Outside Diameter xv tr. GLOSSARY (cont.J ODE One Dimensional Equilibrium ODK One Dimensional Kinetics Pc Thrust Chamber...performance (280 sec steady- state, 220 sec pulsing) have not yet been collectively achieved, but should be obtainable with further development activities...even at nozzle area ratios up to 400:1. The influence of nozzle kinetics (i.e., equilibrium versus frozen flow and ODK ) are noted to be a much more

  18. Low Thrust Trajectory Design for GSFC Missions Project

    National Aeronautics and Space Administration — The Evolutionary Mission Trajectory Generator (EMTG) is a global trajectory optimization tool. EMTG is intended for use in designing interplanetary missions which...

  19. Performance Testing of a Resistojet Thruster for Small Satellite Applications

    Lawrence, Timothy


    ... 1.4 mN compared to 140 mN) which is required to overcome drag at solar maximum. The wet mass of both systems is approximately equal although the propellant volume for the PPT is significantly lower since it is stored in solid form...

  20. Method of interplanetary trajectory optimization for the spacecraft with low thrust and swing-bys

    Konstantinov, M. S.; Thein, M.


    The method developed to avoid the complexity of solving the multipoint boundary value problem while optimizing interplanetary trajectories of the spacecraft with electric propulsion and a sequence of swing-bys is presented in the paper. This method is based on the use of the preliminary problem solutions for the impulsive trajectories. The preliminary problem analyzed at the first stage of the study is formulated so that the analysis and optimization of a particular flight path is considered as the unconstrained minimum in the space of the selectable parameters. The existing methods can effectively solve this problem and make it possible to identify rational flight paths (the sequence of swing-bys) to receive the initial approximation for the main characteristics of the flight path (dates, values of the hyperbolic excess velocity, etc.). These characteristics can be used to optimize the trajectory of the spacecraft with electric propulsion. The special feature of the work is the introduction of the second (intermediate) stage of the research. At this stage some characteristics of the analyzed flight path (e.g. dates of swing-bys) are fixed and the problem is formulated so that the trajectory of the spacecraft with electric propulsion is optimized on selected sites of the flight path. The end-to-end optimization is carried out at the third (final) stage of the research. The distinctive feature of this stage is the analysis of the full set of optimal conditions for the considered flight path. The analysis of the characteristics of the optimal flight trajectories to Jupiter with Earth, Venus and Mars swing-bys for the spacecraft with electric propulsion are presented. The paper shows that the spacecraft weighing more than 7150 kg can be delivered into the vicinity of Jupiter along the trajectory with two Earth swing-bys by use of the space transportation system based on the "Angara A5" rocket launcher, the chemical upper stage "KVTK" and the electric propulsion system with input electrical power of 100 kW.

  1. Power-limited low-thrust trajectory optimization with operation point detection

    Chi, Zhemin; Li, Haiyang; Jiang, Fanghua; Li, Junfeng


    The power-limited solar electric propulsion system is considered more practical in mission design. An accurate mathematical model of the propulsion system, based on experimental data of the power generation system, is used in this paper. An indirect method is used to deal with the time-optimal and fuel-optimal control problems, in which the solar electric propulsion system is described using a finite number of operation points, which are characterized by different pairs of thruster input power. In order to guarantee the integral accuracy for the discrete power-limited problem, a power operation detection technique is embedded in the fourth-order Runge-Kutta algorithm with fixed step. Moreover, the logarithmic homotopy method and normalization technique are employed to overcome the difficulties caused by using indirect methods. Three numerical simulations with actual propulsion systems are given to substantiate the feasibility and efficiency of the proposed method.

  2. Comparison of Low-Thrust Control Laws for Application in Planetocentric Space

    Falck, Robert D.; Sjauw, Waldy K.; Smith, David A.


    Recent interest at NASA for the application of solar electric propulsion for the transfer of significant payloads in cislunar space has led to the development of high-fidelity simulations of such missions. With such transfers involving transfer times on the order of months, simulation time can be significant. In the past, the examination of such missions typically began with the use of lower-fidelity trajectory optimization tools such as SEPSPOT to develop and tune guidance laws which delivered optimal or near- optimal trajectories, where optimal is generally defined as minimizing propellant expenditure or time of flight. The transfer of these solutions to a high-fidelity simulation is typically an iterative process whereby the initial solution may nearly, but not precisely, meet mission objectives. Further tuning of the guidance algorithm is typically necessary when accounting for high-fidelity perturbations such as those due to more detailed gravity models, secondary-body effects, solar radiation pressure, etc. While trajectory optimization is a useful method for determining optimal performance metrics, algorithms which deliver nearly optimal performance with minimal tuning are an attractive alternative.

  3. Multi-step optimization strategy for fuel-optimal orbital transfer of low-thrust spacecraft

    Rasotto, M.; Armellin, R.; Di Lizia, P.


    An effective method for the design of fuel-optimal transfers in two- and three-body dynamics is presented. The optimal control problem is formulated using calculus of variation and primer vector theory. This leads to a multi-point boundary value problem (MPBVP), characterized by complex inner constraints and a discontinuous thrust profile. The first issue is addressed by embedding the MPBVP in a parametric optimization problem, thus allowing a simplification of the set of transversality constraints. The second problem is solved by representing the discontinuous control function by a smooth function depending on a continuation parameter. The resulting trajectory optimization method can deal with different intermediate conditions, and no a priori knowledge of the control structure is required. Test cases in both the two- and three-body dynamics show the capability of the method in solving complex trajectory design problems.

  4. Automated Design of Propellant-Optimal, End-to-End, Low-Thrust Trajectories for Trojan Asteroid Tours

    Stuart, Jeffrey; Howell, Kathleen; Wilson, Roby


    The Sun-Jupiter Trojan asteroids are celestial bodies of great scientific interest as well as potential resources offering water and other mineral resources for longterm human exploration of the solar system. Previous investigations under this project have addressed the automated design of tours within the asteroid swarm. This investigation expands the current automation scheme by incorporating options for a complete trajectory design approach to the Trojan asteroids. Computational aspects of the design procedure are automated such that end-to-end trajectories are generated with a minimum of human interaction after key elements and constraints associated with a proposed mission concept are specified.

  5. Tuning Monotonic Basin Hopping: Improving the Efficiency of Stochastic Search as Applied to Low-Thrust Trajectory Optimization

    Englander, Jacob A.; Englander, Arnold C.


    Trajectory optimization methods using monotonic basin hopping (MBH) have become well developed during the past decade [1, 2, 3, 4, 5, 6]. An essential component of MBH is a controlled random search through the multi-dimensional space of possible solutions. Historically, the randomness has been generated by drawing random variable (RV)s from a uniform probability distribution. Here, we investigate the generating the randomness by drawing the RVs from Cauchy and Pareto distributions, chosen because of their characteristic long tails. We demonstrate that using Cauchy distributions (as first suggested by J. Englander [3, 6]) significantly improves monotonic basin hopping (MBH) performance, and that Pareto distributions provide even greater improvements. Improved performance is defined in terms of efficiency and robustness. Efficiency is finding better solutions in less time. Robustness is efficiency that is undiminished by (a) the boundary conditions and internal constraints of the optimization problem being solved, and (b) by variations in the parameters of the probability distribution. Robustness is important for achieving performance improvements that are not problem specific. In this work we show that the performance improvements are the result of how these long-tailed distributions enable MBH to search the solution space faster and more thoroughly. In developing this explanation, we use the concepts of sub-diffusive, normally-diffusive, and super-diffusive random walks (RWs) originally developed in the field of statistical physics.

  6. Green micro-resistojet research at Delft University of Technology: the new frontiers of Cubesat propulsion

    Cervone, A.; Zandbergen, B.T.C.; Cordeiro Guerrieri, D. C.; de Athayde Costa e Silva, Marsil A. C.; van Zeijl, H.W.


    The aerospace industry is recently expressing a growing interest in green, safe and non-toxic propellants for the propulsion systems of the new generation of space vehicles, which is especially true in the case of Cubesat micro-propulsion systems. Demanding requirements are associated to the future

  7. UV-Curable Hybrid Nanocomposite Coating to Protect Tether Polymer Materials, Phase I

    National Aeronautics and Space Administration — To address the NASA need for coatings to protect and strengthen tether materials for Momentum-exchange Electrodynamic Reboost (MXER) technology, Luminit, LLC,...

  8. Skylab viscous damper study


    The proposed magnetically anchored viscous fluid damper can maintain the Skylab in a gravity-gradient stabilized mode at the anticipated reboost altitudes. The parameters influencing damper performance (and thereby affecting the degree of risk) are: (1) amount of skylab pitch bias in the orbit plane which will result from aerodynamic trim conditions of the post-reboost configuration Skylab; (2) the lowest altitude to which the post-reboost Skylab will be allowed to decay prior to the next rendezvous; (3) maximum allowable weight and size of the proposed damper in order to match shuttle/TRS mission constraints; (4) the amount of magnetic materials expected to be in the vicinity of the damper.

  9. Shuttle performance enhancements using an OMS payload bay kit

    Templin, Kevin C.; Mallini, Charles J.


    The study focuses on the use of an orbital maneuvering system (OMS) payload bay kit (PBK) designed to utilize OMS tanks identical to those currently employed in the Orbiter OMS pods. Emphasis is placed on payload deployment capability and payload servicing/reboost capability augmentation from the point of view of payload mass, maximum deployment altitudes, and initial retrieval and final deployment altitudes. The deployment, servicing, and reboost requirements of the Hubble Space Telescope and Advanced X-ray and Astrophysics Facility are analyzed in order to show the benefits an OMS PBK can provide for these missions. It is shown that OMS PBKs can provide the required capability enhancement necessary to support deployment, reboost, and servicing of payloads requiring altitudes greater than 325 nautical miles.

  10. STS-49 Endeavour/Intelsat Briefing


    Lak Virdee of Intelsat, summarizes Intelsat's role in the STS-49 Endeavour mission. He discusses the reboost hardware, giving details on the capture arm and docker adapter assembly. He describes the rendezvous between Intelsat and the Endeavour Orbiter. Mr. Virdee then answers questions from the press.

  11. Propulsion element requirements using electrical power system unscheduled power

    Zimmermann, Frank; Hodge, Kathy


    The suitability of using the electrical energy from the Space Station's Electrical Power System (EPS) during the periods of peak solar insolation which is currently not specifically allocated (unscheduled power) to produce propulsion propellants, gaseous hydrogen, and oxygen by electrolyzing water is investigated. Reboost propellant requirements are emphasized, but the results are more generally relevant because the balance of recurring propellant requirements are an order of magnitude smaller and the nonrecurring requirements are not significant on an average basis.

  12. A users manual for a computer program which calculates time optical geocentric transfers using solar or nuclear electric and high thrust propulsion

    Sackett, L. L.; Edelbaum, T. N.; Malchow, H. L.


    This manual is a guide for using a computer program which calculates time optimal trajectories for high-and low-thrust geocentric transfers. Either SEP or NEP may be assumed and a one or two impulse, fixed total delta V, initial high thrust phase may be included. Also a single impulse of specified delta V may be included after the low thrust state. The low thrust phase utilizes equinoctial orbital elements to avoid the classical singularities and Kryloff-Boguliuboff averaging to help insure more rapid computation time. The program is written in FORTRAN 4 in double precision for use on an IBM 360 computer. The manual includes a description of the problem treated, input/output information, examples of runs, and source code listings.

  13. “You can get there from here”: Advanced low cost propulsion concepts for small satellites beyond LEO

    Baker, Adam M.; da Silva Curiel, Alex; Schaffner, Jake; Sweeting, Martin


    Small satellites have historically been forced to use low cost propulsion, or to do without in order to maintain low cost. Since 1999 an increasing number of SSTL's customers have demanded the capability to precisely position and subsequently manoeuvre their satellites, driven largely by the current attraction of small satellite constellations such as Disaster Monitoring (DMC), which require propulsion for launcher injection error correction, drag compensation, constellation phasing and proximity manoeuvring and rendezvous. SSTL has successfully flight qualified a simple, low cost propulsion system based on a low power (15-100 W) resistojet employing green propellants such as butane and xenon, and demonstrated key constellation manoeuvres. The system is capable of up to 60 m/s deltaV and will be described here. The SSTL low power resistojet is however limited by a low Isp ( ˜50s for Xenon in the present design, and ˜100s with nitrogen and butane) and a slow reaction time ( 10min warm-up required). An increasing desire to apply small satellite technology to high deltaV missions while retaining the low cost aspect demands new solutions. 'Industry standard' solutions based on cryogenic propulsion, or toxic, carcinogenic storable propellants such as hydrazine/nitrogen oxides combination are not favourable for small satellite missions developed within SSTL's low cost engineering environment. This paper describes a number of strawman missions with high deltaV and/or precision manoeuvring requirements and some low cost propulsion solutions which have been explored at the Surrey Space Centre to meet future needs: Deployment of a complex constellation of nano- or pico-satellites from a secondary launch to a new orbit. The S3TV concept has been developed to allow deployment up to 12 payloads from an 'off-the-shelf' thrust tube, using a restartable nitrous oxide hybrid engine, operating in a dual mode with resistojets for attitude control. Orbit transfer of an enhanced

  14. Preliminary trajectory design for a solar polar observatory using SEP and multiple gravity assists

    Corpaccioli, L.; Noomen, R.; De Smet, S.; Parker, J.S.; Herman, J.F.C.


    Satellite solar observatories have always been of central importance to heliophysics; while there have been numerous such missions, the solar poles have been extremely under-observed. This paper proposes to use low-thrust as well as multiple gravity assists to reach the enormous energies required

  15. Continuous Planetary Polar Observation from Hybrid Pole-Sitters at Venus, Earth, and Mars

    Heiligers, M.J.; van den Oever (student TUDelft), Tom; Ceriotti, M.; Mulligan, P.; McInnes, CR


    A pole-sitter is a satellite that is stationed along the polar axis of the Earth, or any other planet, to generate a continuous, hemispherical view of the planet’s polar regions. In order to maintain such a vantage point, a low-thrust propulsion system is required to counterbalance the gravitational

  16. Orbital Propagation of Momentum Exchange Tether Systems

    Westerhoff, John


    An advanced concept in in-space transportation currently being studied is the Momentum-Exchange/Electrodynamic Reboost Tether System (MXER). The system acts as a large momentum wheel, imparting a Av to a payload in low earth orbit (LEO) at the expense of its own orbital energy. After throwing a payload, the system reboosts itself using an electrodynamic tether to push against Earth's magnetic field and brings itself back up to an operational orbit to prepare for the next payload. The ability to reboost itself allows for continued reuse of the system without the expenditure of propellants. Considering the cost of lifting propellant from the ,ground to LEO to do the same Av boost at $10000 per pound, the system cuts the launch cost of the payload dramatically, and subsequently, the MXER system pays for itself after a small number of missions.1 One of the technical hurdles to be overcome with the MXER concept is the rendezvous maneuver. The rendezvous window for the capture of the payload is on the order of a few seconds, as opposed to traditional docking maneuvers, which can take as long ets necessary to complete a precise docking. The payload, therefore, must be able to match its orbit to meet up with the capture device on the end of the tether at a specific time and location in the future. In order to be able to determine that location, the MXER system must be numerically propagated forward in time to predict where the capture device will be at that instant. It should be kept in mind that the propagation computation must be done faster than real-time. This study focuses on the efforts to find and/or build the tools necessary to numerically propagate the motion of the MXER system as accurately as possible.

  17. Aerospace Fuels From Nonpetroleum Raw Materials

    Palaszewski, Bryan A.; Hepp, Aloysius F.; Kulis, Michael J.; Jaworske, Donald A.


    Recycling human metabolic and plastic wastes minimizes cost and increases efficiency by reducing the need to transport consumables and return trash, respectively, from orbit to support a space station crew. If the much larger costs of transporting consumables to the Moon and beyond are taken into account, developing waste recycling technologies becomes imperative and possibly mission enabling. Reduction of terrestrial waste streams while producing energy and/or valuable raw materials is an opportunity being realized by a new generation of visionary entrepreneurs; several relevant technologies are briefly compared, contrasted and assessed for space applications. A two-step approach to nonpetroleum raw materials utilization is presented; the first step involves production of supply or producer gas. This is akin to synthesis gas containing carbon oxides, hydrogen, and simple hydrocarbons. The second step involves production of fuel via the Sabatier process, a methanation reaction, or another gas-to-liquid technology, typically Fischer-Tropsch processing. Optimization to enhance the fraction of product stream relevant to transportation fuels via catalytic (process) development at NASA Glenn Research Center is described. Energy utilization is a concern for production of fuels whether for operation on the lunar or Martian surface, or beyond. The term green relates to not only mitigating excess carbon release but also to the efficiency of energy usage. For space, energy usage can be an essential concern. Another issue of great concern is minimizing impurities in the product stream(s), especially those that are potential health risks and/or could degrade operations through catalyst poisoning or equipment damage; technologies being developed to remove heteroatom impurities are discussed. Alternative technologies to utilize waste fluids, such as a propulsion option called the resistojet, are discussed. The resistojet is an electric propulsion technology with a powered

  18. Solar Electric Propulsion Triple-Satellite-Aided Capture With Mars Flyby

    Patrick, Sean

    Triple-Satellite-aided-capture sequences use gravity-assists at three of Jupiter's four massive Galilean moons to reduce the DeltaV required to enter into Jupiter orbit. A triple-satellite-aided capture at Callisto, Ganymede, and Io is proposed to capture a SEP spacecraft into Jupiter orbit from an interplanetary Earth-Jupiter trajectory that employs low-thrust maneuvers. The principal advantage of this method is that it combines the ISP efficiency of ion propulsion with nearly impulsive but propellant-free gravity assists. For this thesis, two main chapters are devoted to the exploration of low-thrust triple-flyby capture trajectories. Specifically, the design and optimization of these trajectories are explored heavily. The first chapter explores the design of two solar electric propulsion (SEP), low-thrust trajectories developed using the JPL's MALTO software. The two trajectories combined represent a full Earth to Jupiter capture split into a heliocentric Earth to Jupiter Sphere of Influence (SOI) trajectory and a Joviocentric capture trajectory. The Joviocentric trajectory makes use of gravity assist flybys of Callisto, Ganymede, and Io to capture into Jupiter orbit with a period of 106.3 days. Following this, in chapter two, three more SEP low-thrust trajectories were developed based upon those in chapter one. These trajectories, devised using the high-fidelity Mystic software, also developed by JPL, improve upon the original trajectories developed in chapter one. Here, the developed trajectories are each three separate, full Earth to Jupiter capture orbits. As in chapter one, a Mars gravity assist is used to augment the heliocentric trajectories. Gravity-assist flybys of Callisto, Ganymede, and Io or Europa are used to capture into Jupiter Orbit. With between 89.8 and 137.2-day periods, the orbits developed in chapters one and two are shorter than most Jupiter capture orbits achieved using low-thrust propulsion techniques. Finally, chapter 3 presents an

  19. Rapid space trajectory generation using a Fourier series shape-based approach

    Taheri, Ehsan

    With the insatiable curiosity of human beings to explore the universe and our solar system, it is essential to benefit from larger propulsion capabilities to execute efficient transfers and carry more scientific equipments. In the field of space trajectory optimization the fundamental advances in using low-thrust propulsion and exploiting the multi-body dynamics has played pivotal role in designing efficient space mission trajectories. The former provides larger cumulative momentum change in comparison with the conventional chemical propulsion whereas the latter results in almost ballistic trajectories with negligible amount of propellant. However, the problem of space trajectory design translates into an optimal control problem which is, in general, time-consuming and very difficult to solve. Therefore, the goal of the thesis is to address the above problem by developing a methodology to simplify and facilitate the process of finding initial low-thrust trajectories in both two-body and multi-body environments. This initial solution will not only provide mission designers with a better understanding of the problem and solution but also serves as a good initial guess for high-fidelity optimal control solvers and increases their convergence rate. Almost all of the high-fidelity solvers enjoy the existence of an initial guess that already satisfies the equations of motion and some of the most important constraints. Despite the nonlinear nature of the problem, it is sought to find a robust technique for a wide range of typical low-thrust transfers with reduced computational intensity. Another important aspect of our developed methodology is the representation of low-thrust trajectories by Fourier series with which the number of design variables reduces significantly. Emphasis is given on simplifying the equations of motion to the possible extent and avoid approximating the controls. These facts contribute to speeding up the solution finding procedure. Several example

  20. United States Control Module Guidance, Navigation, and Control Subsystem Design Concept

    Polites, M. E.; Bartlow, B. E.


    Should the Russian Space Agency (RSA) not participate in the International Space Station (ISS) program, then the United States (U.S.) National Aeronautics and Space Administration (NASA) may choose to execute the ISS mission. However, in order to do this, NASA must build two new space vehicles, which must perform the functions that the Russian vehicles and hardware were to perform. These functions include periodic ISS orbit reboost, initial ISS attitude control, and U.S. On-Orbit Segment (USOS) control Moment gyroscope (CMG) momentum desaturation. The two new NASA vehicles that must perform these functions are called the U.S. control module (USCM) and the U.S. resupply module. This paper presents a design concept for the USCM GN&C subsystem, which must play a major role in ISS orbit reboost and initial attitude control, plus USOS CMG momentum desaturation. The proposed concept is structured similar to the USOS GN&C subsystem, by design. It is very robust, in that it allows the USCM to assume a variety of vehicle attitudes and stay power-positive. It has a storage/safe mode that places the USCM in a gravity-gradient orientation and keeps it there for extended periods of time without consuming a great deal of propellant. Simulation results are presented and discussed that show the soundness of the design approach. An equipment list is included that gives detailed information on the baselined GN&C components.

  1. ISS Quasi-steady Accelerometric Data as a Tool for the Detection of External Disturbances During the Period 2009-2016

    Marín, M.; Dubert, D.; Simón, M. J.; Ollé, J.; Gavaldà, Jna.; Ruiz, X.


    The present work aims to investigate the degree of correlation existing between the information contained in the ISS reduced quasi-steady accelerometric data and different external mechanical disturbances (reboostings, dockings/undockings, berthings/deberthings and Extra Vehicular Activities), compiled for the period 2009 to 2016. The eight hour mean (Mean8h) and the eight hour root mean square (RMS8h) acceleration values, considered as reduced data, have been extracted from the quasi-steady records provided by NASA Principal Investigator Microgravity Services website. The advantage of applying the present strategy is to drastically reduce the amount of information to be processed all along these eight years. The Mean8h values have been used for the evaluation of trends as function of time while the RMS8h ones were used to define the level (weak, medium and strong) of the different kind of external mechanical disturbances considered. These criteria has been applied for approximately four hundred selected disturbances, compiled in the Appendix. Results indicate that reboosting is always detected as a strong disturbance, while dockings/undockings, as weak ones, having lower, though detectable level, depending on the type of spacecraft considered. Extra Vehicular Activities are undetectable by the use of this reduced quasi-steady approach. The inverse problem, in other words, knowing the value of the RMS8h one could try to predict the kind of disturbance responsible of it, is thus feasible except for berthing/deberthings and Extra Vehicular Activities.

  2. LOX/LH2 vane pump for auxiliary propulsion systems

    Hemminger, J. A.; Ulbricht, T. E.


    Positive displacement pumps offer potential efficiency advantages over centrifugal pumps for future low thrust space missions. Low flow rate applications, such as space station auxiliary propulsion or dedicated low thrust orbiter transfer vehicles, are typical of missions where low flow and high head rise challenge centrifugal pumps. The positive displacement vane pump for pumping of LOX and LH2 is investigated. This effort has included: (1) a testing program in which pump performance was investigated for differing pump clearances and for differing pump materials while pumping LN2, LOX, and LH2; and (2) an analysis effort, in which a comprehensive pump performance analysis computer code was developed and exercised. An overview of the theoretical framework of the performance analysis computer code is presented, along with a summary of analysis results. Experimental results are presented for pump operating in liquid nitrogen. Included are data on the effects on pump performance of pump clearance, speed, and pressure rise. Pump suction performance is also presented.

  3. On spacecraft maneuvers control subject to propellant engine modes.

    Mazinan, A H


    The paper attempts to address a new control approach to spacecraft maneuvers based upon the modes of propellant engine. A realization of control strategy is now presented in engine on mode (high thrusts as well as further low thrusts), which is related to small angle maneuvers and engine off mode (specified low thrusts), which is also related to large angle maneuvers. There is currently a coarse-fine tuning in engine on mode. It is shown that the process of handling the angular velocities are finalized via rate feedback system in engine modes, where the angular rotations are controlled through quaternion based control (QBCL)strategy in engine off mode and these ones are also controlled through an optimum PID (OPIDH) strategy in engine on mode. Copyright © 2015 ISA. Published by Elsevier Ltd. All rights reserved.

  4. Rapid Preliminary Design of Interplanetary Trajectories Using the Evolutionary Mission Trajectory Generator

    Englander, Jacob


    Preliminary design of interplanetary missions is a highly complex process. The mission designer must choose discrete parameters such as the number of flybys, the bodies at which those flybys are performed, and in some cases the final destination. In addition, a time-history of control variables must be chosen that defines the trajectory. There are often many thousands, if not millions, of possible trajectories to be evaluated. This can be a very expensive process in terms of the number of human analyst hours required. An automated approach is therefore very desirable. This work presents such an approach by posing the mission design problem as a hybrid optimal control problem. The method is demonstrated on notional high-thrust chemical and low-thrust electric propulsion missions. In the low-thrust case, the hybrid optimal control problem is augmented to include systems design optimization.

  5. GOCE reentry predictions for the Italian civil protection authorities

    Pardini, Carmen; Anselmo, Luciano


    The ESA's GOCE satellite was launched on 17 March 2009. After mapping the geopotential with unrivalled accuracy and detail for four years from an extremely low circular polar orbit, on 21 October 2013 the low thrust ion propulsion motor used to contrast the atmospheric drag was automatically shut down when the pressure in the xenon propellant tank dropped below a critical threshold. Then the satellite entered in "fine-pointing mode" (FPM), a phase of orbital altitude decay with active fine at...

  6. Report of the terrestrial bodies science working group. Volume 2. Mercury

    Albee, A.L.; Coroniti, F.V.; Malin, M.C.; Sonett, C.P.


    The objectives and rationale for scientific investigation of Mercury are explored. Knowledge already obtained by astronomical observations and Mariner 10 spacecraft is reviewed and measurements required for the principal scientific goals are described. The use of low thrust propulsion systems is recommended so that maximum scientific return may be achieved and the reconnaissance phase of Mercury exploration may be completed in a single mission. Accelerated development is recommended on solar electric propulsion, solar sails, passive and active cooling mechanisms, and single rough landers

  7. Trajectory design for a rendezvous mission to Earth's Trojan asteroid 2010 TK7

    Lei, Hanlun; Xu, Bo; Zhang, Lei


    In this paper a rendezvous mission to the Earth's Trojan asteroid 2010 TK7 is proposed, and preliminary transfer trajectories are designed. Due to the high inclination (∼ 20.9°) of the target asteroid relative to the ecliptic plane, direct transfers usually require large amounts of fuel consumption, which is beyond the capacity of current technology. As gravity assist technique could effectively change the inclination of spacecraft's trajectory, it is adopted to reduce the launch energy and rendezvous velocity maneuver. In practical computation, impulsive and low-thrust, gravity-assisted trajectories are considered. Among all the trajectories computed, the low-thrust gravity-assisted trajectory with Venus-Earth-Venus (V-E-V) swingby sequence performs the best in terms of propellant mass. For a spacecraft with initial mass of 800 kg , propellant mass of the best trajectory is 36.74 kg . Numerical results indicate that both the impulsive and low-thrust, gravity-assisted trajectories corresponding to V-E-V sequence could satisfy mission constraints, and can be applied to practical rendezvous mission.

  8. The French photovoltaic between light and shade. Self-consumption, a future way for the photovoltaic sector?

    Mary, Olivier; Petitot, Pauline


    This article outlines that France, after having been a leader in the photovoltaic sector, has lost ground during the 2000's. Companies and particularly hardware manufacturers are suffering in front of a harsh competitive environment. However, France still possesses some assets, notably in research and development, and professionals are waiting for some public support to re-boost a sector which is strongly growing everywhere in the world. To illustrate this situation, figures indicate the rate between added value and production for the various concerned activities, the level of added value for these different activities, the distribution of jobs among them, and the distribution of direct, indirect and induced jobs. A second article quotes interventions of a colloquium which addressed the issue of self-consumption of photovoltaic solar energy. It outlines that the development of self-consumption could be an opportunity for the photovoltaic sector if a favourable evolution of the regulatory and tariff framework is introduced

  9. Requirements for a near-earth space tug vehicle

    Gunn, Charles R.


    The requirement for a small but powerful space tug, which will be capable of autonomous orbital rendezvous, docking and translating cargos between near-earth orbits by the end of this decade to support the growing national and international space infrastructure focused near the Space Station Freedom, is described. An aggregate of missions drives the need for a space tug including reboosting decaying satellites back to their operational altitudes, retrieving failed or exhausted satellites to Shuttle or SSF for on-orbit refueling or repair, and transporting a satellite servicer system with an FTS to ailing satellites for supervised in-place repair. It is shown that the development and operation of a space tug to perform such numerous missions is more cost effective than separate module and satellite systems to perform the same tasks.

  10. Space Environmental Effects on Coated Tether Materials

    Gittemeier, Keith A.; Hawk, Clark W.; Finckenor, Miria M.; Watts, Ed


    The University of Alabama in Huntsville s Propulsion Research Center has teamed with NASA's Marshall Space Flight Center (MSFC) to research the effects of atomic oxygen (AO) bombardment on coated tether materials. Tethers Unlimited Inc. has provided several candidate tether materials with various coatings for AO exposure in MSFC s Atomic Oxygen Beam Facility. Additional samples were exposed to ultraviolet (UV) radiation at MSFC. AO erodes most organic materials, and ultraviolet radiation embrittles polymers. This test series was performed to determine the effect of AO and UV on the mechanical integrity of tether materials that were treated with AO-protective coatings, such as polyhedral oligomeric silsesquioxane (POSS) or metallization. Both TUI's Multi-Application Survivable Tether (MAST) Experiment and Marshall Space Flight Center s Momentum Exchange Electrodynamic Reboost (MXER) programs will benefit from this research by helping to determine tether materials and coatings that give the longest life with the lowest mass penalty.

  11. Evolution of International Space Station GN&C System Across ISS Assembly Stages

    Lee, Roscoe; Frank, K. D. (Technical Monitor)


    The Guidance Navigation and Control (GN&C) system for the International Space Station is initially implemented by the Functional Cargo Block (FGB) which was built by the Khrunichev Space Center under direct contract to Boeing. This element (Stage 1A/R) was launched on 20 November 1998 and is currently operating on-orbit. The components and capabilities of the FGB Motion Control System (MCS) are described. The next ISS element, which has GN&C functionality will be the Service Module (SM) built by Rocket Space Corporation-Energia. This module is scheduled for launch (Stage 1R) in early 2000. Following activation of the SM GN&C system, the FGB MCS is deactivated and no longer used. The components and capabilities of the SM GN&C system are described. When a Progress vehicle is attached to the ISS it can be used for reboost operations, based on commands provided by the Mission Control Center-Moscow. When a data connection is implemented between the SM and the Progress, the SM can command the Progress thrusters for attitude control and reboosts. On Stage 5A, the U.S. GN&C system will become activated when the U.S. Laboratory is de loyed and installed (launch schedule is currently TBD). The U.S. GN&C system provides non-propulsive control capabilities to support micro-gravity operations and minimize the use of propellant for attitude control, and an independent capability for determining the ISS state vector, attitude, attitude rate. and time.. The components and capabilities of the U.S. GN&C system are described and the interactions between the U.S. and Russian Segment GN&C systems are also described.

  12. Use of High-Power Brayton Nuclear Electric Propulsion (NEP) for a 2033 Mars Round-Trip Mission

    McGuire, Melissa L.; Martini, Michael C.; Packard, Thomas W.; Weglian, John E.; Gilland, James H.


    The Revolutionary Aerospace Systems Concepts (RASC) team, led by the NASA Langley Research Center, is tasked with exploring revolutionary new approaches to enabling NASA to achieve its strategic goals and objectives in future missions. This paper provides the details from the 2004-2005 RASC study of a point-design that uses a high-power nuclear electric propulsion (NEP) based space transportation architecture to support a manned mission to Mars. The study assumes a high-temperature liquid-metal cooled fission reactor with a Brayton power conversion system to generate the electrical power required by magnetoplasmadynamic (MPD) thrusters. The architecture includes a cargo vehicle with an NEP system providing 5 MW of electrical power and a crewed vehicle with an NEP system with two reactors providing a combined total of 10 MW of electrical power. Both vehicles use a low-thrust, high-efficiency (5000 sec specific impulse) MPD system to conduct a spiral-out of the Earth gravity well, a low-thrust heliocentric trajectory, and a spiral-in at Mars with arrival late in 2033. The cargo vehicle carries two moon landers to Mars and arrives shortly before the crewed vehicle. The crewed vehicle and cargo vehicle rendezvous in Mars orbit and, over the course of the 60-day stay, the crew conducts nine-day excursions to Phobos and Deimos with the landers. The crewed vehicle then spirals out of Martian orbit and returns via a low-thrust trajectory to conduct an Earth flyby. The crew separates from the vehicle prior to Earth flyby and aerobrakes for a direct-entry landing

  13. A Crewed Mission to Apophis Using a Hybrid Bimodal Nuclear Thermal Electric Propulsion (BNTEP) System

    Mccurdy, David R.; Borowski, Stanley K.; Burke, Laura M.; Packard, Thomas W.


    A BNTEP system is a dual propellant, hybrid propulsion concept that utilizes Bimodal Nuclear Thermal Rocket (BNTR) propulsion during high thrust operations, providing 10's of kilo-Newtons of thrust per engine at a high specific impulse (Isp) of 900 s, and an Electric Propulsion (EP) system during low thrust operations at even higher Isp of around 3000 s. Electrical power for the EP system is provided by the BNTR engines in combination with a Brayton Power Conversion (BPC) closed loop system, which can provide electrical power on the order of 100's of kWe. High thrust BNTR operation uses liquid hydrogen (LH2) as reactor coolant propellant expelled out a nozzle, while low thrust EP uses high pressure xenon expelled by an electric grid. By utilizing an optimized combination of low and high thrust propulsion, significant mass savings over a conventional NTR vehicle can be realized. Low thrust mission events, such as midcourse corrections (MCC), tank settling burns, some reaction control system (RCS) burns, and even a small portion at the end of the departure burn can be performed with EP. Crewed and robotic deep space missions to a near Earth asteroid (NEA) are best suited for this hybrid propulsion approach. For these mission scenarios, the Earth return V is typically small enough that EP alone is sufficient. A crewed mission to the NEA Apophis in the year 2028 with an expendable BNTEP transfer vehicle is presented. Assembly operations, launch element masses, and other key characteristics of the vehicle are described. A comparison with a conventional NTR vehicle performing the same mission is also provided. Finally, reusability of the BNTEP transfer vehicle is explored.

  14. Asteroid retrieval missions enabled by invariant manifold dynamics

    Sánchez, Joan Pau; García Yárnoz, Daniel


    Near Earth Asteroids are attractive targets for new space missions; firstly, because of their scientific importance, but also because of their impact threat and prospective resources. The asteroid retrieval mission concept has thus arisen as a synergistic approach to tackle these three facets of interest in one single mission. This paper reviews the methodology used by the authors (2013) in a previous search for objects that could be transported from accessible heliocentric orbits into the Earth's neighbourhood at affordable costs (or Easily Retrievable Objects, a.k.a. EROs). This methodology consisted of a heuristic pruning and an impulsive manoeuvre trajectory optimisation. Low thrust propulsion on the other hand clearly enables the transportation of much larger objects due to its higher specific impulse. Hence, in this paper, low thrust retrieval transfers are sought using impulsive trajectories as first guesses to solve the optimal control problem. GPOPS-II is used to transcribe the continuous-time optimal control problem to a nonlinear programming problem (NLP). The latter is solved by IPOPT, an open source software package for large-scale NLPs. Finally, a natural continuation procedure that increases the asteroid mass allows to find out the largest objects that could be retrieved from a given asteroid orbit. If this retrievable mass is larger than the actual mass of the asteroid, the asteroid retrieval mission for this particular object is said to be feasible. The paper concludes with an updated list of 17 EROs, as of April 2016, with their maximum retrievable masses by means of low thrust propulsion. This ranges from 2000 tons for the easiest object to be retrieved to 300 tons for the least accessible of them.

  15. A globally nonsingular quaternion-based formulation for all-electric satellite trajectory optimization

    Libraro, Paola

    The general electric propulsion orbit-raising maneuver of a spacecraft must contend with four main limiting factors: the longer time of flight, multiple eclipses prohibiting continuous thrusting, long exposure to radiation from the Van Allen belt and high power requirement of the electric engines. In order to optimize a low-thrust transfer with respect to these challenges, the choice of coordinates and corresponding equations of motion used to describe the kinematical and dynamical behavior of the satellite is of critical importance. This choice can potentially affect the numerical optimization process as well as limit the set of mission scenarios that can be investigated. To increase the ability to determine the feasible set of mission scenarios able to address the challenges of an all-electric orbit-raising, a set of equations free of any singularities is required to consider a completely arbitrary injection orbit. For this purpose a new quaternion-based formulation of a spacecraft translational dynamics that is globally nonsingular has been developed. The minimum-time low-thrust problem has been solved using the new set of equations of motion inside a direct optimization scheme in order to investigate optimal low-thrust trajectories over the full range of injection orbit inclinations between 0 and 90 degrees with particular focus on high-inclinations. The numerical results consider a specific mission scenario in order to analyze three key aspects of the problem: the effect of the initial guess on the shape and duration of the transfer, the effect of Earth oblateness on transfer time and the role played by, radiation damage and power degradation in all-electric minimum-time transfers. Finally trade-offs between mass and cost savings are introduced through a test case.

  16. Radioisotope electric propulsion of sciencecraft to the outer solar system and near-interstellar space

    Noble, R.J.


    Recent results are presented in the study of radioisotope electric propulsion as a near-term technology for sending small robotic sciencecraft to the outer Solar System and near-interstellar space. Radioisotope electric propulsion (REP) systems are low-thrust, ion propulsion units based on radioisotope electric generators and ion thrusters. Powerplant specific masses are expected to be in the range of 100 to 200 kg/kW of thrust power. Planetary rendezvous missions to Pluto, fast missions to the heliopause (100 AU) with the capability to decelerate an orbiter for an extended science program and prestellar missions to the first gravitational lens focus of the Sun (550 AU) are investigated

  17. Interplanetary Trajectory Design for the Asteroid Robotic Redirect Mission Alternate Approach Trade Study

    Merrill, Raymond Gabriel; Qu, Min; Vavrina, Matthew A.; Englander, Jacob A.; Jones, Christopher A.


    This paper presents mission performance analysis methods and results for the Asteroid Robotic Redirect Mission (ARRM) option to capture a free standing boulder on the surface of a 100 m or larger NEA. It details the optimization and design of heliocentric low-thrust trajectories to asteroid targets for the ARRM solar electric propulsion spacecraft. Extensive searches were conducted to determine asteroid targets with large pick-up mass potential and potential observation opportunities. Interplanetary trajectory approximations were developed in method based tools for Itokawa, Bennu, 1999 JU3, and 2008 EV5 and were validated by end-to-end integrated trajectories.

  18. Trajectory Design for a Cislunar Cubesat Leveraging Dynamical Systems Techniques: The Lunar Icecube Mission

    Bosanac, Natasha; Cox, Andrew; Howell, Kathleen C.; Folta, David


    Lunar IceCube is a 6U CubeSat that is designed to detect and observe lunar volatiles from a highly inclined orbit. This spacecraft, equipped with a low-thrust engine, will be deployed from the upcoming Exploration Mission-1 vehicle in late 2018. However, significant uncertainty in the deployment conditions for secondary payloads impacts both the availability and geometry of transfers that deliver the spacecraft to the lunar vicinity. A framework that leverages dynamical systems techniques is applied to a recently updated set of deployment conditions and spacecraft parameter values for the Lunar IceCube mission, demonstrating the capability for rapid trajectory design.

  19. DESTINY+ Trajectory Design to (3200) Phaethon

    Sarli, Bruno Victorino; Horikawa, Makoto; Yam, Chit Hong; Kawakatsu, Yasuhiro; Yamamoto, Takayuki


    This work explores the target selection and trajectory design of the mission candidate for ISAS/JAXA's small science satellite series, DESTINY PLUS or DESTINY+. This mission combines unique aspects of the latest satellite technology and exploration of transition bodies to fill a technical and scientific gap in the Japanese space science program. The spacecraft is targeted to study the comet-asteroid transition body (3200) Phaethon through a combination of low-thrust propulsion and Earth Gravity Assist. The trajectory design concept is presented in details together with the launch window and flyby date analysis. Alternative targets for a possible mission extension scenario are also explored.

  20. Effects of rocket engines on laser during lunar landing

    Wan, Xiong, E-mail: [Key Laboratory of Space Active Opto-Electronics Technology, Shanghai Institute of Technical Physics, Chinese Academy of Science, Shanghai 200083 (China); Key Laboratory of Nondestructive Test (Ministry of Education), Nanchang Hangkong University, Nanchang 330063 (China); Shu, Rong; Huang, Genghua [Key Laboratory of Space Active Opto-Electronics Technology, Shanghai Institute of Technical Physics, Chinese Academy of Science, Shanghai 200083 (China)


    In the Chinese moon exploration project “ChangE-3”, the laser telemeter and lidar are important equipments on the lunar landing vehicle. A low-thrust vernier rocket engine works during the soft landing, whose plume may influence on the laser equipments. An experiment has first been accomplished to evaluate the influence of the plume on the propagation characteristics of infrared laser under the vacuum condition. Combination with our theoretical analysis has given an appropriate assessment of the plume's effects on the infrared laser hence providing a valuable basis for the design of lunar landing systems.

  1. Industrial numerical analysis

    McKee, S.; Elliott, C.M.


    The applications of mathematics to industrial problems involves the formulation of problems which are amenable to mathematical investigation, mathematical modelling, the solution of the mathematical problem and the inter-pretation of the results. There are 12 chapters describing industrial problems where mathematics and numerical analysis can be applied. These range from the numerical assessment of the flatness of engineering surfaces and plates, the design of chain links, control problems in tidal power generation and low thrust satellite trajectory optimization to mathematical models in welding. One chapter, on the ageing of stainless steels, is indexed separately. (UK)

  2. Effects of rocket engines on laser during lunar landing

    Wan, Xiong; Shu, Rong; Huang, Genghua


    In the Chinese moon exploration project “ChangE-3”, the laser telemeter and lidar are important equipments on the lunar landing vehicle. A low-thrust vernier rocket engine works during the soft landing, whose plume may influence on the laser equipments. An experiment has first been accomplished to evaluate the influence of the plume on the propagation characteristics of infrared laser under the vacuum condition. Combination with our theoretical analysis has given an appropriate assessment of the plume's effects on the infrared laser hence providing a valuable basis for the design of lunar landing systems

  3. Are there heavy quarks of mass 23 GeV

    Cornet, F.; Hagiwara, K.; Zeppenfeld, D.; Glover, E.W.N.; Martin, A.D.


    An excess of events with an isolated muon and low thrust observed by the MARK-J collaboration at the highest PETRA energy √s=46.7 GeV, is found to be consistent with the near threshold production of heavy quarks of charge -1/3. A natural candidate is a fourth generation 'down' quark or, possibly, a member of a 27 representation of E 6 . We investigate signatures of such heavy quark pair production at the CERN panti p collider and conclude that the present data have a chance to confirm the signal. (orig.)

  4. Reference Design for a Simple, Durable and Refuelable Interplanetary Spacecraft

    McConnell, B. S.; Tolley, A. M.

    This article describes a reference design for interplanetary vessels, composed mostly of water, that utilize simplified RF engines for low thrust, long duration propulsion, and hydrogen peroxide for short duration, high thrust burns. The electrothermal engines are designed to heat a wide range of liquid materials, possibly also milled solids or surface dusts. The system emphasizes simple components and processes based on older technologies, many well known since the 1960s, that are understandable, can process a variety of materials, and are easily serviced in flight. The goal is to radically simplify systems and their inter-dependencies, to a point where a reasonably skilled person can learn to operate these vessels, not unlike a sailboat, and to eliminate many design and testing bottlenecks in their construction. The use of water, or hydrogen peroxide generated in situ from that water, is multiply advantageous because it can be used for structure, consumption, irrigation, radiation and debris shielding, and thermal regulation, and thus greatly reduce dead weight by creating an almost fully consumable ship. This also enables the ship to utilize a wide range of in situ materials, and eventually obtain reaction mass from lower gravity sites. The ability to switch between low thrust, constant power and high thrust, short duration maneuvers will enable these ships to travel freely and reach many interesting destinations throughout the solar system. One can think of them as “spacecoaches”, not unlike the prairie schooners of the Old West, which were rugged, serviceable by tradesmen, and easily maintained.

  5. Space-to-Space Power Beaming Enabling High Performance Rapid Geocentric Orbit Transfer

    Dankanich, John W.; Vassallo, Corinne; Tadge, Megan


    The use of electric propulsion is more prevalent than ever, with industry pursuing all electric orbit transfers. Electric propulsion provides high mass utilization through efficient propellant transfer. However, the transfer times become detrimental as the delta V transitions from near-impulsive to low-thrust. Increasing power and therefore thrust has diminishing returns as the increasing mass of the power system limits the potential acceleration of the spacecraft. By using space-to-space power beaming, the power system can be decoupled from the spacecraft and allow significantly higher spacecraft alpha (W/kg) and therefore enable significantly higher accelerations while maintaining high performance. This project assesses the efficacy of space-to-space power beaming to enable rapid orbit transfer while maintaining high mass utilization. Concept assessment requires integrated techniques for low-thrust orbit transfer steering laws, efficient large-scale rectenna systems, and satellite constellation configuration optimization. This project includes the development of an integrated tool with implementation of IPOPT, Q-Law, and power-beaming models. The results highlight the viability of the concept, limits and paths to infusion, and comparison to state-of-the-art capabilities. The results indicate the viability of power beaming for what may be the only approach for achieving the desired transit times with high specific impulse.

  6. Water Vapour Propulsion Powered by a High-Power Laser-Diode

    Minami, Y.; Uchida, S.

    Most of the laser propulsion schemes now being proposed and developed assume neither power supplies nor on-board laser devices and therefore are bound to remote laser stations like a kite via a laser beam “string”. This is a fatal disadvantage for a space vehicle that flies freely though it is often said that no need of installing an energy source is an advantage of a laser propulsion scheme. The possibility of an independent laser propulsion space vehicle that carries a laser source and a power supply on board is discussed. This is mainly due to the latest development of high power laser diode (LD) technology. Both high specific impulse-low thrust mode and high thrust-low specific impulse mode can be selected by controlling the laser output by using vapour or water as a propellant. This mode change can be performed by switching between a high power continuous wave (cw), LD engine for high thrust with a low specific impulse mode and high power LD pumping Q-switched Nd:YAG laser engine for low thrust with the high specific impulse mode. This paper describes an Orbital Transfer Vehicle equipped with the above-mentioned laser engine system and fuel cell that flies to the Moon from a space platform or space hotel in Earth orbit, with cargo shipment from lunar orbit to the surface of the Moon, including the possibility of a sightseeing trip.

  7. Strategies for the sustained human exploration of Mars

    Landau, Damon Frederick

    A variety of mission scenarios are compared in this thesis to assess the strengths and weaknesses of options for Mars exploration. The mission design space is modeled along two dimensions: trajectory architectures and propulsion system technologies. Direct, semi-direct, stop-over, semi-cycler, and cycler architectures are examined, and electric propulsion, nuclear thermal rockets, methane and oxygen production on Mars, Mars water excavation, aerocapture, and reusable propulsion systems are included in the technology assessment. The mission sensitivity to crew size, vehicle masses, and crew travel time is also examined. The primary figure of merit for a mission scenario is the injected mass to low-Earth orbit (IMLEO), though technology readiness levels (TRL) are also included. Several elements in the architecture dimension are explored in more detail. The Earth-Mars semi-cycler architecture is introduced and five families of Earth-Mars semi-cycler trajectories are presented along with optimized itineraries. Optimized cycler trajectories are also presented. In addition to Earth-Mars semi-cycler and cycler trajectories, conjunction-class, free-return, Mars-Earth semi-cycler, and low-thrust trajectories are calculated. Design parameters for optimal DeltaV trajectories are provided over a range of flight times (from 120 to 270 days) and launch years (between 2009 and 2022). Unlike impulsive transfers, the mass-optimal low-thrust trajectory depends strongly on the thrust and specific impulse of the propulsion system. A low-thrust version of the rocket equation is provided where the initial mass or thrust may be minimized by varying the initial acceleration and specific impulse. Planet-centered operations are also examined. A method to rotate a parking orbit about the line of apsides to achieve the proper orientation at departure is discussed, thus coupling the effects of parking-orbit orientation with the interplanetary trajectories. Also, a guidance algorithm for

  8. PEG Enhancement for EM1 and EM2+ Missions

    Von der Porten, Paul; Ahmad, Naeem; Hawkins, Matt


    NASA is currently building the Space Launch System (SLS) Block-1 launch vehicle for the Exploration Mission 1 (EM-1) test flight. The next evolution of SLS, the Block-1B Exploration Mission 2 (EM-2), is currently being designed. The Block-1 and Block-1B vehicles will use the Powered Explicit Guidance (PEG) algorithm. Due to the relatively low thrust-to-weight ratio of the Exploration Upper Stage (EUS), certain enhancements to the Block-1 PEG algorithm are needed to perform Block-1B missions. In order to accommodate mission design for EM-2 and beyond, PEG has been significantly improved since its use on the Space Shuttle program. The current version of PEG has the ability to switch to different targets during Core Stage (CS) or EUS flight, and can automatically reconfigure for a single Engine Out (EO) scenario, loss of communication with the Launch Abort System (LAS), and Inertial Navigation System (INS) failure. The Thrust Factor (TF) algorithm uses measured state information in addition to a priori parameters, providing PEG with an improved estimate of propulsion information. This provides robustness against unknown or undetected engine failures. A loft parameter input allows LAS jettison while maximizing payload mass. The current PEG algorithm is now able to handle various classes of missions with burn arcs much longer than were seen in the shuttle program. These missions include targeting a circular LEO orbit with a low-thrust, long-burn-duration upper stage, targeting a highly eccentric Trans-Lunar Injection (TLI) orbit, targeting a disposal orbit using the low-thrust Reaction Control System (RCS), and targeting a hyperbolic orbit. This paper will describe the design and implementation of the TF algorithm, the strategy to handle EO in various flight regimes, algorithms to cover off-nominal conditions, and other enhancements to the Block-1 PEG algorithm. This paper illustrates challenges posed by the Block-1B vehicle, and results show that the improved PEG

  9. The GNC Measurement System for the Automated Transfer Vehicle

    Roux, Y.; da Cunha, P.

    The Automated Transfer Vehicle (ATV) is a European Space Agency (ESA) funded spacecraft developed by EADS Space Transportation as prime contractor for the space segment together with major European industrial partners, in the frame of the International Space Station (ISS). Its mission objective is threefold : to supply the station with fret and propellant, to reboost ISS to a higher orbit and to dispose of waste from the station. The ATV first flight, called Jules Verne and planned on 2005, will be the first European Vehicle to perform an orbital rendezvous. The GNC Measurement System (GMS) is the ATV on board function in charge of the measurement data collection and preconditioning for the navigation, guidance and control (GNC) algorithms. The GMS is made up of hardware which are the navigation sensors (with a certain level of hardware redundancy for each of them), and of an on-board software that manages, monitors and performs consistency checks to detect and isolate potential sensor failures. The GMS relies on six kinds of navigation sensors, used during various phases of the mission : the gyrometers assembly (GYRA), the accelerometers assembly (ACCA), the star trackers (STR), the GPS receivers, the telegoniometers (TGM) and the videometers (VDM), the last two being used for the final rendezvous phase. The GMS function is developed by EADS Space Transportation together with other industrial partners: EADS Astrium, EADS Sodern, Laben and Dasa Jena Optronik.

  10. Natural Hazards of the Space Environment

    Evans, Steven W.; Kross, Dennis A. (Technical Monitor)


    Spacecraft in Low Earth Orbit (LEO) are subject to numerous environmental hazards. Here I'll briefly discuss three environment factors that pose acute threats to the survival of spacecraft systems and crew: atmospheric drag, impacts by meteoroids and orbital debris, and ionizing radiation. Atmospheric drag continuously opposes the orbital motion of a satellite, causing the orbit to decay. This decay will lead to reentry if not countered by reboost maneuvers. Orbital debris is a by-product of man's activities in space, and consists of objects ranging in size from miniscule paint chips to spent rocket stages and dead satellites. Ionizing radiation experienced in LEO has several components: geomagnetically trapped protons and electrons (Van Allen belts); energetic solar particles; galactic cosmic rays; and albedo neutrons. These particles can have several types of prompt harmful effects on equipment and crew, from single-event upsets, latchup, and burnout of electronics, to lethal doses to crew.All three types of prompt threat show some dependence on the solar activity cycle. Atmospheric drag mitigation and large debris avoidance require propulsive maneuvers. M/OD and ionizing radiation require some form of shielding for crew and sensitive equipment. Limiting exposure time is a mitigation technique for ionizing radiation and meteor streams.

  11. Real-time on-line space research laboratory environment monitoring with off-line trend and prediction analysis

    Jules, Kenol; Lin, Paul P.


    With the International Space Station currently operational, a significant amount of acceleration data is being down-linked, processed and analyzed daily on the ground on a continuous basis for the space station reduced gravity environment characterization, the vehicle design requirements verification and science data collection. To help understand the impact of the unique spacecraft environment on the science data, an artificial intelligence monitoring system was developed, which detects in near real time any change in the reduced gravity environment susceptible to affect the on-going experiments. Using a dynamic graphical display, the monitoring system allows science teams, at any time and any location, to see the active vibration disturbances, such as pumps, fans, compressor, crew exercise, re-boost and extra-vehicular activities that might impact the reduced gravity environment the experiments are exposed to. The monitoring system can detect both known and unknown vibratory disturbance activities. It can also perform trend analysis and prediction by analyzing past data over many increments (an increment usually lasts 6 months) collected onboard the station for selected disturbances. This feature can be used to monitor the health of onboard mechanical systems to detect and prevent potential systems failures. The monitoring system has two operating modes: online and offline. Both near real-time on-line vibratory disturbance detection and off-line detection and trend analysis are discussed in this paper.

  12. Integrating environment protection, a new challenge: strategy of the International Union of Radioecology

    Brechignac, F.; Alexakhin, R.; Godoy, J.M.; Oughton, D.; Sheppard, S.; Strand, P.


    Born in the fifties together with the emergence of the nuclear technologies, radioecology is a scientific discipline that primarily addresses environmental issues relevant to radioprotection. With a current membership of nearly 600 worldwide, the International Union of Radioecology was founded in the seventies as a non-governmental knowing society dedicated to the development and the promotion of this discipline. The scientific directions taken in Radioecology have been drastically influenced in the past by the Chernobyl accident, which forced a focus on environmental transfers through the environment to feed human radioprotection needs. Currently, a profound evolution is underway towards more ecological effects research and studies, under the driving pressure of the raise of society concern on environmental issues and the concomitant re-boost of nuclear industry to face global warming and the future energetic demands. The I.U.R. plays a central role within this evolution which is described here in more details along a description of its four major tools of action: dedicated task groups; workshops, seminars and conferences; training courses; web site tool for information and communication. Finally, together with the recent election of a new Board of Council to manage the Union, the main lines of the new strategic plan for the coming years are given. (author)

  13. Integrating environment protection, a new challenge: strategy of the International Union of Radioecology; Protection de l'environnement, un nouveau chantier: strategie de l'Union internationale de radioecologie

    Brechignac, F. [Institut de Radioprotection et de Serete Nucleaire (IRSN), Dir. Scientifique, Centre de Cadarache, 13 - Saint Paul lez Durance (France); Alexakhin, R. [Russian Institute of Agricultural Radiology and Agroecology (RIARAE), Obninsk, Kaluga Region (Russian Federation); Godoy, J.M. [Instituto de Radioprotecao, Comissao National de Energia Nuclear, Rio de Janeiro, (Brazil); Oughton, D. [Norvegian Univ. of Life Sciences, Dpt. of Plant Environmental Sciences, As (Norway); Sheppard, S. [ECO Matters Incorporation., Pinawa, MB (Canada); Strand, P. [Norwegian Radiation Protection Authority, Osteraas (Norway)


    Born in the fifties together with the emergence of the nuclear technologies, radioecology is a scientific discipline that primarily addresses environmental issues relevant to radioprotection. With a current membership of nearly 600 worldwide, the International Union of Radioecology was founded in the seventies as a non-governmental knowing society dedicated to the development and the promotion of this discipline. The scientific directions taken in Radioecology have been drastically influenced in the past by the Chernobyl accident, which forced a focus on environmental transfers through the environment to feed human radioprotection needs. Currently, a profound evolution is underway towards more ecological effects research and studies, under the driving pressure of the raise of society concern on environmental issues and the concomitant re-boost of nuclear industry to face global warming and the future energetic demands. The I.U.R. plays a central role within this evolution which is described here in more details along a description of its four major tools of action: dedicated task groups; workshops, seminars and conferences; training courses; web site tool for information and communication. Finally, together with the recent election of a new Board of Council to manage the Union, the main lines of the new strategic plan for the coming years are given. (author)

  14. Space tug applications. Final report


    This article is the final report of the conceptual design efforts for a 'space tug'. It includes preliminary efforts, mission analysis, configuration analysis, impact analysis, and conclusions. Of the several concepts evaluated, the nuclear bimodal tug was one of the top candidates, with the two options being the NEBA-1 and NEBA-3 systems. Several potential tug benefits were identified during the mission analysis. The tug enables delivery of large (>3,500 kg) payloads to the outer planets and it increases the GSO delivery capability by 20% relative to current systems. By providing end of life disposal, the tug can be used to extend the life of existing space assets. It can also be used to reboost satellites which were not delivered to their final orbit by the launch system. A specific mission model is the key to validating the tug concept. Once a mission model can be established, mission analysis can be used to determine more precise propellant quantities and burn times. In addition, the specific payloads can be evaluated for mass and volume capability with the launch systems. Results of the economic analysis will be dependent on the total years of operations and the number of missions in the mission model. The mission applications evaluated during this phase drove the need for large propellant quantities and thus did not allow the payloads to step down to smaller and less expensive launch systems

  15. Impact risk assessment for the ATV using ESABASE/DEBRIS

    Beltrami Karlezi, P.; Drolshagen, G.; Lambert, M.


    The European Space Agency ESA participates in the International Space Station with various programs, one of them being the Automated Transfer Vehicle (ATV). The ATV is an unmanned servicing and logistics vehicle launched on Ariane 5 and designed to fulfil different roles like cargo transport, re-supply of fuel and consumables and orbit re-boost of the International Space Station (ISS). For this reason it is important that the risks imposed on these modules by meteoroids and orbital debris are calculated accurately. Following such calculations the Meteoroid and Orbital Debris Protection System (M/ODPS) can be optimised. This paper presents the results of the risk assessment of meteoroids and space debris for the ATV spacecraft attached to the ISS using different shield configurations. The results are presented as the probability of no penetration (PNP) for each component and each configuration. They are compared to a target PNP requirement of 0.999 for 135 days and the weight penalty produced by the extra shielding is given.

  16. Human Spaceflight Recent Conjunctions of Interest

    Browns, Ansley C.


    I. During each nine-hour shift (or upon request), the Orbital Safety Analyst (OSA) at JSpOC updates the entire tracked catalog with the latest tracking data from the SSN and screens this catalog against NASA s assets. a) For ISS operations, a 72-hour advance screening is performed. b) For Shuttle orbit operations, a 36-hour advance screening is performed. c) If a vehicle is performing a maneuver during the screening period, OSA uses post-reboost-trajectory data supplied by Mission Control Center-Houston (MCC-H) for screening. II. An automated process is used to detect any conjunctions within 10 x 40 x 40 km box (centered on the vehicle) using Special Perturbation (SP) processing: a) Box dimensions are Radial x Downtrack x Crosstrack. b) Any object found within this box has the tracking tasking level increased to improve (hopefully) its uncertainty in its current and predicted orbital trajectory. c) OSA informs NASA if any object is found inside a 2 x 25 x 25 km box then creates and sends an Orbital Conjunction Message (OCM) to NASA which contains detailed information about the conjunction. d) For Shuttle, the box size used for screening and reporting is altered for special operations (day of rendezvous, launch screening, etc.).

  17. Dual-purpose self-deliverable lunar surface PV electrical power system

    Arnold, Jack H.; Harris, David W.; Cross, Eldon R.; Flood, Dennis J.


    A safe haven and work supported PV power systems on the lunar surface will likely be required by NASA in support of the manned outpost scheduled for the post-2000 lunar/Mars exploration and colonization initiative. Initial system modeling and computer analysis shows that the concept is workable and contains no major high risk technology issues which cannot be resolved in the circa 2000 to 2025 timeframe. A specific selection of the best suited type of electric thruster has not been done; the initial modeling was done using an ion thruster, but Rocketdyne must also evaluate arc and resisto-jets before a final design can be formulated. As a general observation, it appears that such a system can deliver itself to the Moon using many system elements that must be transported as dead payload mass in more conventional delivery modes. It further appears that a larger power system providing a much higher safe haven power level is feasible if this delivery system is implemented, perhaps even sufficient to permit resource prospecting and/or lab experimentation. The concept permits growth and can be expanded to include cargo transport such as habitat and working modules. In short, the combined payload could be manned soon after landing and checkout. NASA has expended substantial resources in the development of electric propulsion concepts and hardware that can be applied to a lunar transport system such as described herein. In short, the paper may represent a viable mission on which previous investments play an invaluable role. A more comprehensive technical paper which embodies second generation analysis and system size will be prepared for near-term presentation.

  18. Pole-shape optimization of permanent-magnet linear synchronous motor for reduction of thrust ripple

    Tavana, Nariman Roshandel, E-mail: [Department of Electrical Engineering, Iran University of Science and Technology, Narmak, Tehran 16846-13114 (Iran, Islamic Republic of); Shoulaie, Abbas, E-mail: [Department of Electrical Engineering, Iran University of Science and Technology, Narmak, Tehran 16846-13114 (Iran, Islamic Republic of)


    In this paper, we have used magnet arc shaping technique in order to improve the performance of permanent-magnet linear synchronous motor (PMLSM). At first, a detailed analytical modeling based on Maxwell equations is presented for the analysis and design of PMLSM with the arc-shaped magnetic poles (ASMPs). Then the accuracy of presented method is verified by finite-element method. Very close agreement between the analytical and finite-element results shows the effectiveness of the proposed method. Finally, a magnet shape design is carried out based on the analytical method to enhance the motor developed thrust. Pertinent evaluations on the optimal design performance demonstrate that shape optimization leads to a design with extra low thrust ripple.

  19. Astrodynamics Conference, Williamsburg, VA, August 18-20, 1986, Technical Papers



    Various papers on satellite and planetary mission dynamics are presented. Individual topics discussed include: orbital motion in the frequency domain, the Galileo orbital tour for the 1986 launch opportunity, circulating transportation orbits between earth and Mars, optimum burn scheduling for low-thrust orbital transfers, comparison of nonlinear filters for orbit determination, common period four-satellite continuous global coverage, motion of a drag-free experiment on the Space Station, and engineering improvements, sequencing and orbital determination for Voyager 2 Uranus encounter. Also considered are: approach strategy for rendezvous with Comet Tempel 2, maneuver strategies for Magellan Mission, Jupiter Polar Orbiter Mission concepts, modelling and simulation of Spacecraft Solar Array deployment, sensitivity analysis for lunar trajectories, maximum likelihood estimation for orbit determination, adaptive estimation for an accelerating spacecraft, Slinger in-orbit booster facility, simulation of satellite breakups, missile launch windows of avoiding collisions with satellites, linear-quadratic stationkeeping for the STS Orbiter

  20. Hybrid Differential Dynamic Programming with Stochastic Search

    Aziz, Jonathan; Parker, Jeffrey; Englander, Jacob


    Differential dynamic programming (DDP) has been demonstrated as a viable approach to low-thrust trajectory optimization, namely with the recent success of NASAs Dawn mission. The Dawn trajectory was designed with the DDP-based Static Dynamic Optimal Control algorithm used in the Mystic software. Another recently developed method, Hybrid Differential Dynamic Programming (HDDP) is a variant of the standard DDP formulation that leverages both first-order and second-order state transition matrices in addition to nonlinear programming (NLP) techniques. Areas of improvement over standard DDP include constraint handling, convergence properties, continuous dynamics, and multi-phase capability. DDP is a gradient based method and will converge to a solution nearby an initial guess. In this study, monotonic basin hopping (MBH) is employed as a stochastic search method to overcome this limitation, by augmenting the HDDP algorithm for a wider search of the solution space.

  1. LO2/LH2 propulsion for outer planet orbiter spacecraft

    Garrison, P. W.; Sigurdson, K. B.


    Galileo class orbiter missions (750-1500 kg) to the outer planets require a large postinjection delta-V for improved propulsion performance. The present investigation shows that a pump-fed low thrust LO2/LH2 propulsion system can provide a significantly larger net on-orbit mass for a given delta-V than a state-of-the-art earth storable, N2O4/monomethylhydrazine pressure-fed propulsion system. A description is given of a conceptual design for a LO2/LH2 pump-fed propulsion system developed for a Galileo class mission to the outer planets. Attention is given to spacecraft configuration, details regarding the propulsion system, the thermal control of the cryogenic propellants, and aspects of mission performance.

  2. Design Analysis of Taper Width Variations in Magnetless Linear Machine for Traction Applications

    Saadha Aminath


    Full Text Available Linear motors are being used in a different application with a huge popularity in the use of transport industry. With the invention of maglev trains and other high-speed trains, linear motors are being used for the translation and braking applications for these systems. However, a huge drawback of the linear motor design is the cogging force, low thrust values, and voltage ripples. This paper aims to study the force analysis with change in taper/teeth width of the motor stator and mover to understand the best teeth ratio to obtain a high flux density and a high thrust. The analysis is conducted through JMAG software and it is found that the optimum teeth ratio for both the stator and mover gives an increase of 94.4% increases compared to the 0.5mm stator and mover width.

  3. Operations analysis (study 2.1). Program SEPSIM (solar electric propulsion stage simulation). [in FORTRAN: space tug

    Lang, T. J.


    Program SEPSIM is a FORTRAN program which performs deployment, servicing, and retrieval missions to synchronous equatorial orbit using a space tug with a continuous low thrust upper stage known as a solar electric propulsion stage (SEPS). The SEPS ferries payloads back and forth between an intermediate orbit and synchronous orbit, and performs the necessary servicing maneuvers in synchronous orbit. The tug carries payloads between the orbiter and the intermediate orbit, deploys fully fueled SEPS vehicles, and retrieves exhausted SEPS vehicles when, and if, required. The program is presently contained in subroutine form in the Logistical On-orbit VEhicle Servicing (LOVES) Program, but can also be run independently with the addition of a simple driver program.

  4. Multiple Turbine Wakes

    Machefaux, Ewan; Mann, Jakob

    and to obtain an estimate of the wake expansion in a fixed frame of reference. A comparison of selected datasets from the campaign showed good far wake agreements of mean wake expansion with Actuator Line CFD computations and simpler engineering models. An empirical relationship, relating maximum wake induction...... for modeling the resulting double wake deficit is only relevant at high turbine thrust coefficients. For high wind speed and low thrust coefficient, linear summation should be primarily used. The first iteration of a new engineering model capable of modeling the overlapped wake deficit is formulated and its...... measurement and simulation is seen in both the fixed and the meandering frame of reference. A benchmark of several wake accumulation models is performed as a basis for the subsequent development of an engineering model for wake interaction.Finally, the validated numerical CFD model is used as part...

  5. Application of Solar-Electric Propulsion to Robotic and Human Missions in Near-Earth Space

    Woodcock, Gordon R.; Dankanich, John


    Interest in applications of solar electric propulsion (SEP) is increasing. Application of SEP technology is favored when: (1) the mission is compatible with low-thrust propulsion, (2) the mission needs high total delta V such that chemical propulsion is disadvantaged; and (3) performance enhancement is needed. If all such opportunities for future missions are considered, many uses of SEP are likely. Representative missions are surveyed and several SEP applications selected for analysis, including orbit raising, lunar science, lunar exploration, lunar exploitation, planetary science, and planetary exploration. These missions span SEP power range from 10s of kWe to several MWe. Modes of use and benefits are described, and potential SEP evolution is discussed.

  6. Application of Solar-Electric Propulsion to Robotic Missions in Near-Earth Space

    Woodcock, Gordon R.; Dankanich, John


    Interest in applications of solar electric propulsion (SEP) is increasing. Application of SEP technology is favored when: (1) the mission is compatible with low-thrust propulsion, (2) the mission needs high total delta V such that chemical propulsion is disadvantaged; and (3) performance enhancement is needed. If all such opportunities for future missions are considered, many uses of SEP are likely. Representative missions are surveyed and several SEP applications selected for analysis, including orbit raising, lunar science and robotic exploration, and planetary science. These missions span SEP power range from 10 kWe to about 100 kWe. A SEP design compatible with small inexpensive launch vehicles, and capable of lunar science missions, is presented. Modes of use and benefits are described, and potential SEP evolution is discussed.

  7. Mission Analysis for the Don Quijote Phase-A Study

    Cano, Juan L.; Sanchez, Mariano; Cornara, Stefania; Carnelli, Ian


    The Don Quijote Phase-A study is a definition study funded by ESA and devoted to the analysis of the possibilities to deflect a Near Earth Object (NEO) in the range of 300-800 m diameter. DEIMOS Space S.L. and EADS Astrium have teamed up within this study to form one of the three consortia that have analyzed these aspects for ESA. Target asteroids for the mission are 1989 ML, 2002 AT4 and Apophis. This paper presents the mission analysis activities within the consortium providing: low-thrust interplanetary rendezvous Orbiter trajectories to the target asteroids, ballistic interplanetary trajectories for the Impactor, Orbiter arrival description at the asteroids, Orbiter stable orbits characterization at the asteroid, deflection determination by means of a Radio Science Experiment (RSE) as well as the mission timelines and overall mission scenarios.

  8. Solar Power System Options for the Radiation and Technology Demonstration Spacecraft

    Kerslake, Thomas W.; Haraburda, Francis M.; Riehl, John P.


    The Radiation and Technology Demonstration (RTD) Mission has the primary objective of demonstrating high-power (10 kilowatts) electric thruster technologies in Earth orbit. This paper discusses the conceptual design of the RTD spacecraft photovoltaic (PV) power system and mission performance analyses. These power system studies assessed multiple options for PV arrays, battery technologies and bus voltage levels. To quantify performance attributes of these power system options, a dedicated Fortran code was developed to predict power system performance and estimate system mass. The low-thrust mission trajectory was analyzed and important Earth orbital environments were modeled. Baseline power system design options are recommended on the basis of performance, mass and risk/complexity. Important findings from parametric studies are discussed and the resulting impacts to the spacecraft design and cost.

  9. 10 KWe dual-mode space nuclear power system for military and scientific applications

    Malloy, J.; Westerman, K.; Rochow, R.; Scoles, S.


    This paper discusses a 10 KWe dual-mode space power system concept which has been identified and is based on INEL's Small Externally-fueled Heat Pipe Thermionic Reactor (SEHPTR) concept. This power system will enhance user capabilities by providing reliable electric power and by providing two propulsion systems; electric power for an arc-jet electric propulsion system and direct thrust by heating hydrogen propellant inside the reactor. The low thrust electric thrusters allow efficient station keeping and long-term maneuvering. This paper will focus on the nuclear power system design, including: the reactor with its UO 2 fuel in tungsten clad, 36 thermionic heat pipe modules (THPMs) which produce electricity within the reactor and remove waste heat, radiation shielding, waste heat radiators, and reactivity control systems. The use of non-vented fuel elements for short lifetime missions (under five years) will be described

  10. The Near Earth Object (NEO) Scout Spacecraft: A Low-cost Approach to In-situ Characterization of the NEO Population

    Woeppel, Eric A.; Balsamo, James M.; Fischer, Karl J.; East, Matthew J.; Styborski, Jeremy A.; Roche, Christopher A.; Ott, Mackenzie D.; Scorza, Matthew J.; Doherty, Christopher D.; Trovato, Andrew J.; hide


    This paper describes a microsatellite spacecraft with supporting mission profile and architecture, designed to enable preliminary in-situ characterization of a significant number of Near Earth Objects (NEOs) at reasonably low cost. The spacecraft will be referred to as the NEO-Scout. NEO-Scout spacecraft are to be placed in Geosynchronous Equatorial Orbit (GEO), cis-lunar space, or on earth escape trajectories as secondary payloads on launch vehicles headed for GEO or beyond, and will begin their mission after deployment from the launcher. A distinguishing key feature of the NEO-Scout system is to design the spacecraft and mission timeline so as to enable rendezvous with and landing on the target NEO during NEO close approach (the Earth-Moon system using low-thrust/high-impulse propulsion systems. Mission durations are on the order 100 to 400 days. Mission feasibility and preliminary design analysis are presented, along with detailed trajectory calculations.

  11. Trajectory Design for the Lunar Polar Hydrogen Mapper Mission

    Genova, Anthony L.; Dunham, David W.


    The presented trajectory was designed for the Lunar Polar Hydrogen Mapper (LunaH-Map) 6U CubeSat, which was awarded a ride on NASAs Space Launch System (SLS) with Exploration Mission 1 (EM-1) via NASAs 2015 SIMPLEX proposal call. After deployment from EM-1s upper stage (which is planned to enter heliocentric space via a lunar flyby), the LunaH-Map CubeSat will alter its trajectory via its low-thrust ion engine to target a lunar flyby that yields a Sun-Earth-Moon weak stability boundary transfer to set up a ballistic lunar capture. Finally, the orbit energy is lowered to reach the required quasi-frozen science orbit with periselene above the lunar south pole.

  12. Recent advances in celestial and space mechanics

    Chyba, Monique


    This book presents recent advances in space and celestial mechanics, with a focus on the N-body problem and astrodynamics, and explores the development and application of computational techniques in both areas. It highlights the design of space transfers with various modes of propulsion, like solar sailing and low-thrust transfers between libration point orbits, as well as a broad range of targets and applications, like rendezvous with near Earth objects. Additionally, it includes contributions on the non-integrability properties of the collinear three- and four-body problem, and on general conditions for the existence of stable, minimum energy configurations in the full N-body problem. A valuable resource for physicists and mathematicians with research interests in celestial mechanics, astrodynamics and optimal control as applied to space transfers, as well as for professionals and companies in the industry.

  13. Aeroelastic behavior of composite rotor blades with swept tips

    Yuan, Kuo-An; Friedmann, Peretz P.; Venkatesan, Comandur


    This paper presents an analytical study of the aeroelastic behavior of composite rotor blades with straight and swept tips. The blade is modeled by beam type finite elements. A single finite element is used to model the swept tip. The nonlinear equations of motion for the finite element model are derived using Hamilton's principle and based on a moderate deflection theory and accounts for: arbitrary cross-sectional shape, pretwist, generally anisotropic material behavior, transverse shears and out-of-plane warping. Numerical results illustrating the effects of tip sweep, anhedral and composite ply orientation on blade aeroelastic behavior are presented. It is shown that composite ply orientation has a substantial effect on blade stability. At low thrust conditions, certain ply orientations can cause instability in the lag mode. The flap-torsion coupling associated with tip sweep can also induce aeroelastic instability in the blade. This instability can be removed by appropriate ply orientation in the composite construction.

  14. Pole-shape optimization of permanent-magnet linear synchronous motor for reduction of thrust ripple

    Tavana, Nariman Roshandel; Shoulaie, Abbas


    In this paper, we have used magnet arc shaping technique in order to improve the performance of permanent-magnet linear synchronous motor (PMLSM). At first, a detailed analytical modeling based on Maxwell equations is presented for the analysis and design of PMLSM with the arc-shaped magnetic poles (ASMPs). Then the accuracy of presented method is verified by finite-element method. Very close agreement between the analytical and finite-element results shows the effectiveness of the proposed method. Finally, a magnet shape design is carried out based on the analytical method to enhance the motor developed thrust. Pertinent evaluations on the optimal design performance demonstrate that shape optimization leads to a design with extra low thrust ripple.

  15. Outer Planet Missions with Electric Propulsion Systems—Part I

    Carlos Renato Huaura Solórzano


    Full Text Available For interplanetary missions, efficient electric propulsion systems can be used to increase the mass delivered to the destination. Outer planet exploration has experienced new interest with the launch of the Cassini and New Horizons Missions. At the present, new technologies are studied for better use of electric propulsion systems in missions to the outer planets. This paper presents low-thrust trajectories using the method of the transporting trajectory to Uranus, Neptune, and Pluto. They use nuclear and radio isotopic electric propulsion. These direct transfers have continuous electric propulsion of low power along the entire trajectory. The main goal of the paper is to optimize the transfers, that is, to provide maximum mass to be delivered to the outer planets.

  16. Analysis of Electric Propulsion System for Exploration of Saturn

    Carlos Renato Huaura Solórzano


    Full Text Available Exploration of the outer planets has experienced new interest with the launch of the Cassini and the New Horizons Missions. At the present time, new technologies are under study for the better use of electric propulsion system in deep space missions. In the present paper, the method of the transporting trajectory is used to study this problem. This approximated method for the flight optimization with power-limited low thrust is based on the linearization of the motion of a spacecraft near a keplerian orbit that is close to the transfer trajectory. With the goal of maximizing the mass to be delivered in Saturn, several transfers were studied using nuclear, radioisotopic and solar electric propulsion systems.

  17. Space engineering modeling and optimization with case studies

    Pintér, János


    This book presents a selection of advanced case studies that cover a substantial range of issues and real-world challenges and applications in space engineering. Vital mathematical modeling, optimization methodologies and numerical solution aspects of each application case study are presented in detail, with discussions of a range of advanced model development and solution techniques and tools. Space engineering challenges are discussed in the following contexts: •Advanced Space Vehicle Design •Computation of Optimal Low Thrust Transfers •Indirect Optimization of Spacecraft Trajectories •Resource-Constrained Scheduling, •Packing Problems in Space •Design of Complex Interplanetary Trajectories •Satellite Constellation Image Acquisition •Re-entry Test Vehicle Configuration Selection •Collision Risk Assessment on Perturbed Orbits •Optimal Robust Design of Hybrid Rocket Engines •Nonlinear Regression Analysis in Space Engineering< •Regression-Based Sensitivity Analysis and Robust Design ...

  18. Energy management - The delayed flap approach

    Bull, J. S.


    Flight test evaluation of a Delayed Flap approach procedure intended to provide reductions in noise and fuel consumption is underway using the NASA CV-990 test aircraft. Approach is initiated at a high airspeed (240 kt) and in a drag configuration that allows for low thrust. The aircraft is flown along the conventional ILS glide slope. A Fast/Slow message display signals the pilot when to extend approach flaps, landing gear, and land flaps. Implementation of the procedure in commercial service may require the addition of a DME navigation aid co-located with the ILS glide slope transmitter. The Delayed Flap approach saves 250 lb of fuel over the Reduced Flap approach, with a 95 EPNdB noise contour only 43% as large.

  19. Comparison of three Stark problem solution techniques for the bounded case

    Hatten, Noble; Russell, Ryan P.


    Three methods of obtaining solutions to the Stark problem—one developed by Lantoine and Russell using Jacobi elliptic and related functions, one developed by Biscani and Izzo using Weierstrass elliptic and related functions, and one developed by Pellegrini, Russell, and Vittaldev using and Taylor series extended to the Stark problem—are compared qualitatively and quantitatively for the bounded motion case. For consistency with existing available code for the series solution, Fortran routines of the Lantoine method and Biscani method are newly implemented and made available. For these implementations, the Lantoine formulation is found to be more efficient than the Biscani formulation in the propagation of a single trajectory segment. However, for applications for which acceptable accuracy may be achieved by orders up to 16, the Pellegrini series solution is shown to be more efficient than either analytical method. The three methods are also compared in the propagation of sequentially connected trajectory segments in a low-thrust orbital transfer maneuver. Separate tests are conducted for discretizations between 8 and 96 segments per orbit. For the series solution, the interaction between order and step size leads to computation times that are nearly invariable to discretization for a given truncation error tolerance over the tested range of discretizations. This finding makes the series solution particularly attractive for mission design applications where problems may require both coarse and fine discretizations. Example applications include the modeling of low-thrust propulsion and time-varying perturbations—problems for which the efficient propagation of relatively short Stark segments is paramount because the disturbing acceleration generally varies continuously.

  20. A Historical Systems Study of Liquid Rocket Engine Throttling Capabilities

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


    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.

  1. Design of cycler trajectories and analysis of solar influences on radioactive decay rates during space missions

    Rogers, Blake A.

    This thesis investigates the design of interplanetary missions for the continual habitation of Mars via Earth-Mars cyclers and for the detection of variations in nuclear decay rates due to solar influences. Several cycler concepts have been proposed to provide safe and comfortable quarters for astronauts traveling between the Earth and Mars. However, no literature has appeared to show how these massive vehicles might be placed into their cycler trajectories. Trajectories are designed that use either Vinfinity leveraging or low thrust to establish cycler vehicles in their desired orbits. In the cycler trajectory cases considered, the use of Vinfinity leveraging or low thrust substantially reduces the total propellant needed to achieve the cycler orbit compared to direct orbit insertion. In the case of the classic Aldrin cycler, the propellant savings due to Vinfinity leveraging can be as large as a 24 metric ton reduction for a cycler vehicle with a dry mass of 75 metric tons, and an additional 111 metric ton reduction by instead using low thrust. The two-synodic period cyclers considered benefit less from Vinfinity leveraging, but have a smaller total propellant mass due to their lower approach velocities at Earth and Mars. It turns out that, for low-thrust establishment, the propellant required is approximately the same for each of the cycler trajectories. The Aldrin cycler has been proposed as a transportation system for human missions between Earth and Mars. However, the hyperbolic excess velocity values at the planetary encounters for these orbits are infeasibly large, especially at Mars. In a new version of the Aldrin cycler, low thrust is used in the interplanetary trajectories to reduce the encounter velocities. Reducing the encounter velocities at both planets reduces the propellant needed by the taxis (astronauts use these taxis to transfer between the planetary surfaces and the cycler vehicle) to perform hyperbolic rendezvous. While the propellant

  2. ``DMS-R, the Brain of the ISS'', 10 Years of Continuous Successful Operation in Space

    Wolff, Bernd; Scheffers, Peter


    Space industries on both sides of the Atlantic were faced with a new situation of collaboration in the beginning of the 1990s.In 1995, industrial cooperation between ASTRIUM ST, Bremen and RSC-E, Moscow started aiming the outfitting of the Russian Service Module ZVEZDA for the ISS with computers. The requested equipments had to provide not only redundancy but fault tolerance and high availability. The design and development of two fault tolerant computers, (FTCs) responsible for the telemetry (Telemetry Computer: TC) and the central control (CC), as well as the man machine interface CPC were contracted to ASTRIUM ST, Bremen. The computer system is responsible e.g. for the life support system and the ISS re-boost control.In July 2000, the integration of the Russian Service Module ZVEZDA with Russian ZARYA FGB and American Node 1 bears witness for transatlantic and European cooperation.The Russian Service module ZVEZDA provides several basic functions as Avionics Control, the Environmental Control and Life Support (ECLS) in the ISS and control of the docked Automatic Transfer Vehicle (ATV) which includes re-boost of ISS. If these elementary functions fail or do not work reliable the effects for the ISS will be catastrophic with respect to Safety (manned space) and ISS mission.For that reason the responsible computer system Data Management System - Russia (DMS-R) is also called "The brain of the ISS".The Russian Service module ZVEZDA, including DMS-R, was launched on 12th of July, 2000. DMS-R was operational also during launch and docking.The talk provide information about the definition, design and development of DMS-R, the integration of DMS-R in the Russian Service module and the maintenance of the system in space. Besides the technical aspects are also the German - Russian cooperation an important subject of this speech. An outlook finalises the talk providing further development activities and application of fault tolerant systems.The importance of the DMS

  3. Mission Analysis for LEO Microwave Power-Beaming Station in Orbital Launch of Microwave Lightcraft

    Myrabo, L. N.; Dickenson, T.


    A detailed mission analysis study has been performed for a 1 km diameter, rechargeable satellite solar power station (SPS) designed to boost 20m diameter, 2400 kg Micr,oWave Lightcraft (MWLC) into low earth orbit (LEO) Positioned in a 476 km daily-repeating oi.bit, the 35 GHz microwave power station is configured like a spinning, thin-film bicycle wheel covered by 30% efficient sola cells on one side and billions of solid state microwave transmitter elements on the other, At the rim of this wheel are two superconducting magnets that can stor,e 2000 G.J of energy from the 320 MW, solar array over a period of several orbits. In preparation for launch, the entire station rotates to coarsely point at the Lightcraft, and then phases up using fine-pointing information sent from a beacon on-board the Lightcraft. Upon demand, the station transmits a 10 gigawatt microwave beam to lift the MWLC from the earth surface into LEO in a flight of several minutes duration. The mission analysis study was comprised of two parts: a) Power station assessment; and b) Analysis of MWLC dynamics during the ascent to orbit including the power-beaming relationships. The power station portion addressed eight critical issues: 1) Drag force vs. station orbital altitude; 2) Solar pressure force on the station; 3) Station orbital lifetime; 4) Feasibility of geo-magnetic re-boost; 5) Beta angle (i..e., sola1 alignment) and power station effective area relationship; 6) Power station percent time in sun vs, mission elapsed time; 7) Station beta angle vs.. charge time; 8) Stresses in station structures.. The launch dynamics portion examined four issues: 1) Ascent mission/trajecto1y profile; 2) MWLC/power-station mission geometry; 3) MWLC thrust angle vs. time; 4) Power station pitch rate during power beaming. Results indicate that approximately 0 58 N of drag force acts upon the station when rotated edge-on to project the minimum frontal area of 5000 sq m. An ion engine or perhaps an electrodynamic

  4. A goat poxvirus-vectored peste-des-petits-ruminants vaccine induces long-lasting neutralization antibody to high levels in goats and sheep.

    Chen, Weiye; Hu, Sen; Qu, Linmao; Hu, Qianqian; Zhang, Qian; Zhi, Haibing; Huang, Kehe; Bu, Zhigao


    Recombinant capripoxvirus (CPV) is a promising candidate differentiating infected from vaccinated animals (DIVA) vaccine against peste-des-petits-ruminants (PPR). In order for recombinant CPV to be successfully used in the field, there should exist dependable indicators for quality control of vaccine products, surveillance and vaccination evaluation. Viral neutralization antibody (VNA) is correlated to protection against PPR and is a technically feasible indicator for this purpose. The immunogenicity of this vectored vaccine in goats and sheep, however, has not been fully evaluated. In this study, we generated two recombinant CPV viruses, rCPV-PPRVH and rCPV-PPRVF, that express PPR virus (PPRV) glycoproteins H and F, respectively. Vaccination studies with different dosages of recombinant viruses showed that rCPV-PPRVH was a more potent inducer of PPRV VNA than rCPV-PPRVF. One dose of rCPV-PPRVH was enough to seroconvert 80% of immunized sheep. A second dose induced significantly higher PPRV VNA titers. There was no significant difference in PPRV VNA responses between goats and sheep. Subcutaneous inoculation also induced a significant PPRV VNA response. PPRV VNA could be detected for over 6 months in more than 80% of vaccinated goats and sheep. Boost vaccination at 6-month intervals induced significant re-boost efficacy of PPRV VNA in goats and sheep. More over, two doses of rCPV-PPRVH could completely overcome the interference caused by pre-existing immunity to the CPV vaccine backbone in animals. Vaccination with rCPV-PPRVH also protected goats from virulent CPV challenge. Our results demonstrate that VNA can serve as a dependent indicator for effective vaccination and immune protection of animals in the field. The recombinant CPV vaccine used in our studies could be a practical and useful candidate DIVA vaccine in countries where PPR newly emerges or where stamp-out plans are yet to be implemented. Copyright 2010 Elsevier Ltd. All rights reserved.

  5. Flight times to the heliopause using a combination of solar and radioisotope electric propulsion

    Ohndorf, Andreas [DLR Deutsches Zentrum fuer Luft- und Raumfahrt e.V., Wessling (Germany); Dachwald, Bernd [FH Univ. of Applied Sciences, Aachen (Germany); Seboldt, Wolfgang [DLR Deutsches Zentrum fuer Luft- und Raumfahrt e.V., Koeln (Germany); Loeb, Horst W.; Schartner, Karl-Heinz [Giessen Univ. (Germany)


    We investigate the interplanetary flight of a low-thrust space probe to the heliopause, located at a distance of about 200 AU from the Sun. Our goal was to reach this distance within the 25 years postulated by ESA for such a mission (which is less ambitious than the 15-year goal set by NASA). Contrary to solar sail concepts and combinations of ballistic and electrically propelled flight legs, we have investigated whether the set flight time limit could also be kept with a combination of solar-electric propulsion and a second, RTG-powered upper stage. The used ion engine type was the RIT-22 for the first stage and the RIT-10 for the second stage. Trajectory optimization was carried out with the low-thrust optimization program InTrance, which implements the method of Evolutionary Neurocontrol, using Artificial Neural Networks for spacecraft steering and Evolutionary Algorithms to optimize the Neural Networks' parameter set. Based on a parameter space study, in which the number of thrust units, the unit's specific impulse, and the relative size of the solar power generator were varied, we have chosen one configuration as reference. The transfer time of this reference configuration was 29.6 years and the fastest one, which is technically more challenging, still required 28.3 years. As all flight times of this parameter study were longer than 25 years, we further shortened the transfer time by applying a launcher-provided hyperbolic excess energy up to 49 km{sup 2}/s{sup 2}. The resulting minimal flight time for the reference configuration was then 27.8 years. The following, more precise optimization to a launch with the European Ariane 5 ECA rocket reduced the transfer time to 27.5 years. This is the fastest mission design of our study that is flexible enough to allow a launch every year. The inclusion of a fly-by at Jupiter finally resulted in a flight time of 23.8 years, which is below the set transfer-time limit. However, compared to the 27.5-year transfer

  6. A One-year, Short-Stay Crewed Mars Mission Using Bimodal Nuclear Thermal Electric Propulsion (BNTEP) - A Preliminary Assessment

    Burke, Laura A.; Borowski, Stanley K.; McCurdy, David R.; Packard, Thomas W.


    A crewed mission to Mars poses a signi cant challenge in dealing with the physiolog- ical issues that arise with the crew being exposed to a near zero-gravity environment as well as signi cant solar and galactic radiation for such a long duration. While long sur- face stay missions exceeding 500 days are the ultimate goal for human Mars exploration, short round trip, short surface stay missions could be an important intermediate step that would allow NASA to demonstrate technology as well as study the physiological e ects on the crew. However, for a 1-year round trip mission, the outbound and inbound hy- perbolic velocity at Earth and Mars can be very large resulting in a signi cant propellant requirement for a high thrust system like Nuclear Thermal Propulsion (NTP). Similarly, a low thrust Nuclear Electric Propulsion (NEP) system requires high electrical power lev- els (10 megawatts electric (MWe) or more), plus advanced power conversion technology to achieve the lower speci c mass values needed for such a mission. A Bimodal Nuclear Thermal Electric Propulsion (BNTEP) system is examined here that uses three high thrust Bimodal Nuclear Thermal Rocket (BNTR) engines allowing short departure and capture maneuvers. The engines also generate electrical power that drives a low thrust Electric Propulsion (EP) system used for ecient interplanetary transit. This combined system can help reduce the total launch mass, system and operational requirements that would otherwise be required for equivalent NEP or Solar Electric Propulsion (SEP) mission. The BNTEP system is a hybrid propulsion concept where the BNTR reactors operate in two separate modes. During high-thrust mode operation, each BNTR provides 10's of kilo- Newtons of thrust at reasonably high speci c impulse (Isp) of 900 seconds for impulsive trans-planetary injection and orbital insertion maneuvers. When in power generation / EP mode, the BNTR reactors are coupled to a Brayton power conversion system allowing each

  7. Mars Earth Return Vehicle (MERV) Propulsion Options

    Oleson, Steven R.; McGuire, Melissa L.; Burke, Laura; Fincannon, James; Warner, Joe; Williams, Glenn; Parkey, Thomas; Colozza, Tony; Fittje, Jim; Martini, Mike; hide


    The COMPASS Team was tasked with the design of a Mars Sample Return Vehicle. The current Mars sample return mission is a joint National Aeronautics and Space Administration (NASA) and European Space Agency (ESA) mission, with ESA contributing the launch vehicle for the Mars Sample Return Vehicle. The COMPASS Team ran a series of design trades for this Mars sample return vehicle. Four design options were investigated: Chemical Return /solar electric propulsion (SEP) stage outbound, all-SEP, all chemical and chemical with aerobraking. The all-SEP and Chemical with aerobraking were deemed the best choices for comparison. SEP can eliminate both the Earth flyby and the aerobraking maneuver (both considered high risk by the Mars Sample Return Project) required by the chemical propulsion option but also require long low thrust spiral times. However this is offset somewhat by the chemical/aerobrake missions use of an Earth flyby and aerobraking which also take many months. Cost and risk analyses are used to further differentiate the all-SEP and Chemical/Aerobrake options.

  8. A comparison of propulsion systems for potential space mission applications

    Harvego, E.A.; Sulmeisters, T.K.


    A derivative of the NERVA nuclear rocket engine was compared with a chemical propulsion system and a nuclear electric propulsion system to assess the relative capabilities of the different propulsion system options for three potential space missions. The missions considered were (1) orbital transfer from low earth orbit (LEO) to geosynchronous earth orbit (GEO), (2) LEO to a lunar base, and (3) LEO to Mars. The results of this comparison indicate that the direct-thrust NERVA-derivative nuclear rocket engine has the best performance characteristics for the missions considered. The combined high thrust and high specific impulse achievable with a direct-thrust nuclear stage permits short operating times (transfer times) comparable to chemical propulsion systems, but with considerably less required propellant. While nuclear-electric propulsion systems are more fuel efficient than either direct-nuclear or chemical propulsion, they are not stand-alone systems, since their relatively low thrust levels require the use of high-thrust ferry or lander stages in high gravity applications such as surface-to-orbit propulsion. The extremely long transfer times and inefficient trajectories associated with electric propulsion systems were also found to be a significant drawback

  9. Radioisotope electric propulsion for robotic science missions to near-interstellar space

    Noble, R.J.


    The use of radioisotope electric propulsion for sending small robotic probes on fast science missions several hundred astronomical units (AU) from the Sun is investigated. Such missions would address a large variety of solar, interstellar, galactic and cosmological science themes from unique vantage points at 100 to 600 AU, including parallax distance measurements for the entire Milky Way Galaxy, sampling of the interstellar medium and imaging of cosmological objects at the gravitational lens foci of the Sun (≥ 550 AU). Radioisotope electric propulsion (REP) systems are low-thrust, ion propulsion units based on multi-hundred watt, radioisotope electric generators and ion thrusters. In a previous work, the flight times for rendezvous missions to the outer planets (< 30 AU) using REP were found to be less than fifteen years. However fast prestellar missions to several hundred AU are not possible unless the probe's energy can be substantially increased in the inner Solar System so as to boost the final hyperbolic excess velocity. In this paper an economical hybrid propulsion scheme combining chemical propulsion and gravity assist in the inner Solar System and radioisotope electric propulsion in the outer Solar System is studied which enables fast prestellar missions. Total hyperbolic excess velocities of 15 AU/year and flight times to 550 AU of about 40 years are possible using REP technology that may be available in the next decade

  10. Effects of the Eccentricity of a Perturbing Third Body on the Orbital Correction Maneuvers of a Spacecraft

    R. C. Domingos


    Full Text Available The fuel consumption required by the orbital maneuvers when correcting perturbations on the orbit of a spacecraft due to a perturbing body was estimated. The main goals are the measurement of the influence of the eccentricity of the perturbing body on the fuel consumption required by the station keeping maneuvers and the validation of the averaged methods when applied to the problem of predicting orbital maneuvers. To study the evolution of the orbits, the restricted elliptic three-body problem and the single- and double-averaged models are used. Maneuvers are made by using impulsive and low thrust maneuvers. The results indicated that the averaged models are good to make predictions for the orbital maneuvers when the spacecraft is in a high inclined orbit. The eccentricity of the perturbing body plays an important role in increasing the effects of the perturbation and the fuel consumption required for the station keeping maneuvers. It is shown that the use of more frequent maneuvers decreases the annual cost of the station keeping to correct the orbit of a spacecraft. An example of an eccentric planetary system of importance to apply the present study is the dwarf planet Haumea and its moons, one of them in an eccentric orbit.

  11. Asteroid Redirection Mission Evaluation Using Multiple Landers

    Bazzocchi, Michael C. F.; Emami, M. Reza


    In this paper, a low-thrust tugboat redirection method is assessed using multiple spacecraft for a target range of small near-Earth asteroids. The benefits of a landed configuration of tugboat spacecraft in formation are examined for the redirection of a near-Earth asteroid. The tugboat method uses a gimballed thruster with a highly collimated ion beam to generate a thrust on the asteroid. The target asteroid range focuses on near-Earth asteroids smaller than 150 m in diameter, and carbonaceous (C-type) asteroids, due to the volatiles available for in-situ utilization. The assessment focuses primarily on the three key parameters, i.e., the asteroid mass redirected, the timeframe for redirection, and the overall system cost. An evaluation methodology for each parameter is discussed in detail, and the parameters are employed to determine the expected return and feasibility of the redirection mission. The number of spacecraft employed is optimized along with the electrical power needed for each spacecraft to ensure the highest possible return on investment. A discussion of the optimization results and the benefits of spacecraft formation for the tugboat method are presented.

  12. Gravitational disturbances generated by the Sun, Phobos and Deimos in orbital maneuvers around Mars with automatic correction of the semi-major axis

    Rocco, E M


    The objective of this work is to analyze orbital maneuvers of a spacecraft orbiting Mars, considering disturbance effects due to the gravitational attraction of the Sun, Phobos and Deimos, beyond the disturbances due to the gravitational potential of Mars. To simulate the trajectory, constructive aspects of the propulsion system were considered. Initially ideal thrusters, capable of applying infinite magnitude of the thrust, were used. Thus, impulsive optimal maneuvers were obtained by scanning the solutions of the Lambert's problem in order to select the maneuver of minimum fuel consumption. Due to the impossibility of applying an impulse, the orbital maneuver must be distributed in a propulsive arc around the position of the impulse given by the solution of the Lambert's problem. However the effect of the propulsive arc is not exactly equivalent to the application of an impulse due to the errors in magnitude and direction of applied thrust. Therefore, the influence of the thrusters’ capacity in the trajectory was evaluated for a more realistic model instead of the ideal case represented by the impulsive approach. Beyond the evaluation of the deviation in the orbital path, was considered an automatic correction of the semi-major axis using continuous low thrust controlled in closed loop to minimize the error in the trajectory after the application of the main thrust. (paper)

  13. Genetic Algorithm Optimizes Q-LAW Control Parameters

    Lee, Seungwon; von Allmen, Paul; Petropoulos, Anastassios; Terrile, Richard


    A document discusses a multi-objective, genetic algorithm designed to optimize Lyapunov feedback control law (Q-law) parameters in order to efficiently find Pareto-optimal solutions for low-thrust trajectories for electronic propulsion systems. These would be propellant-optimal solutions for a given flight time, or flight time optimal solutions for a given propellant requirement. The approximate solutions are used as good initial solutions for high-fidelity optimization tools. When the good initial solutions are used, the high-fidelity optimization tools quickly converge to a locally optimal solution near the initial solution. Q-law control parameters are represented as real-valued genes in the genetic algorithm. The performances of the Q-law control parameters are evaluated in the multi-objective space (flight time vs. propellant mass) and sorted by the non-dominated sorting method that assigns a better fitness value to the solutions that are dominated by a fewer number of other solutions. With the ranking result, the genetic algorithm encourages the solutions with higher fitness values to participate in the reproduction process, improving the solutions in the evolution process. The population of solutions converges to the Pareto front that is permitted within the Q-law control parameter space.

  14. TAU: a design for a thousand astronomical unit voyage

    Eubanks, D.; Alvis, J.; Bechler, E.; Lyon, W. III; McFarlane, D.; Palmrose, D.; Schmitz, P.


    The Jet Propulsion Lab. (JPL) has proposed a deep-space probe to travel to a distance of one thousand astronomical units -25 times further from the Sun than Pluto. In order to achieve this goal within the lifetime of the investigators, the mission time is set at a maximum of 50 yr. The JPL proposal postulates a design in which the probe is under powered thrust for the first 10 yr of the mission and coasts for the next 40 yr. A continuous high specific impulse, Isp (the ratio of thrust to propellant mass flow rate), low thrust propulsion system (either magnetoplasmadynamic (MPD) or ion thrusters) is required in order to achieve this goal. This in turn necessitates electrical power in the megawatt range. The only power source that is practical for this situation is a nuclear reactor. It was a this point that the Nuclear Engineering Dept. at Texas A and M Univ. began its ongoing work, looking into several areas of the proposal in which a more detailed description was needed. These areas of interest were power, propulsion, heavy lift launch capabilities, and trajectory analysis. In addition to all of the boundaries previously outlined, the technology level is assumed to be that of 1995, 8 yr from now

  15. Multi-Axis Thrust Measurements of the EO-1 Pulsed Plasma Thruster

    Arrington, Lynn A.; Haag, Thomas W.


    Pulsed plasma thrusters are low thrust propulsive devices which have a high specific impulse at low power. A pulsed plasma thruster is currently scheduled to fly as an experiment on NASA's Earth Observing-1 satellite mission. The pulsed plasma thruster will be used to replace one of the reaction wheels. As part of the qualification testing of the thruster it is necessary to determine the nominal thrust as a function of charge energy. These data will be used to determine control algorithms. Testing was first completed on a breadboard pulsed plasma thruster to determine nominal or primary axis thrust and associated propellant mass consumption as a function of energy and then later to determine if any significant off-axis thrust component existed. On conclusion that there was a significant off-axis thrust component with the bread-board in the direction of the anode electrode, the test matrix was expanded on the flight hardware to include thrust measurements along all three orthogonal axes. Similar off-axis components were found with the flight unit.

  16. Comparison study of exhaust plume impingement effects of small mono- and bipropellant thrusters using parallelized DSMC method.

    Kyun Ho Lee

    Full Text Available A space propulsion system is important for the normal mission operations of a spacecraft by adjusting its attitude and maneuver. Generally, a mono- and a bipropellant thruster have been mainly used for low thrust liquid rocket engines. But as the plume gas expelled from these small thrusters diffuses freely in a vacuum space along all directions, unwanted effects due to the plume collision onto the spacecraft surfaces can dramatically cause a deterioration of the function and performance of a spacecraft. Thus, aim of the present study is to investigate and compare the major differences of the plume gas impingement effects quantitatively between the small mono- and bipropellant thrusters using the computational fluid dynamics (CFD. For an efficiency of the numerical calculations, the whole calculation domain is divided into two different flow regimes depending on the flow characteristics, and then Navier-Stokes equations and parallelized Direct Simulation Monte Carlo (DSMC method are adopted for each flow regime. From the present analysis, thermal and mass influences of the plume gas impingements on the spacecraft were analyzed for the mono- and the bipropellant thrusters. As a result, it is concluded that a careful understanding on the plume impingement effects depending on the chemical characteristics of different propellants are necessary for the efficient design of the spacecraft.

  17. Solar Electric Propulsion Technologies Being Designed for Orbit Transfer Vehicle Applications

    Sarver-Verhey, Timothy R.; Hoffman, David J.; Kerslake, Thomas W.; Oleson, Steven R.; Falck, Robert D.


    There is increasing interest in employing Solar Electric Propulsion (SEP) for new missions requiring transfer from low Earth orbit to the Earth-Moon Lagrange point, L1. Mission architecture plans place the Gateway Habitat at L1 in the 2011 to 2016 timeframe. The Gateway Habitat is envisioned to be used for Lunar exploration, space telescopes, and planetary mission staging. In these scenarios, an SEP stage, or "tug," is used to transport payloads to L1--such as the habitat module, lunar excursion and return vehicles, and chemical propellant for return crew trips. SEP tugs are attractive because they are able to efficiently transport large (less than 10,000 kg) payloads while minimizing propellant requirements. To meet the needs of these missions, a preliminary conceptual design for a general-purpose SEP tug was developed that incorporates several of the advanced space power and in-space propulsion technologies (such as high-power gridded ion and Hall thrusters, high-performance thin-film photovoltaics, lithium-ion batteries, and advanced high-voltage power processing) being developed at the NASA Glenn Research Center. A spreadsheet-based vehicle system model was developed for component sizing and is currently being used for mission planning. This model incorporates a low-thrust orbit transfer algorithm to make preliminary determinations of transfer times and propellant requirements. Results from this combined tug mass estimation and orbit transfer model will be used in a higher fidelity trajectory model to refine the analysis.

  18. Advanced concept for a crewed mission to the martian moons

    Conte, Davide; Di Carlo, Marilena; Budzyń, Dorota; Burgoyne, Hayden; Fries, Dan; Grulich, Maria; Heizmann, Sören; Jethani, Henna; Lapôtre, Mathieu; Roos, Tobias; Castillo, Encarnación Serrano; Schermann, Marcel; Vieceli, Rhiannon; Wilson, Lee; Wynard, Christopher


    This paper presents the conceptual design of the IMaGInE (Innovative Mars Global International Exploration) Mission. The mission's objectives are to deliver a crew of four astronauts to the surface of Deimos and perform a robotic exploration mission to Phobos. Over the course of the 343 day mission during the years 2031 and 2032, the crew will perform surface excursions, technology demonstrations, In Situ Resource Utilization (ISRU) of the Martian moons, as well as site reconnaissance for future human exploration of Mars. This mission design makes use of an innovative hybrid propulsion concept (chemical and electric) to deliver a relatively low-mass reusable crewed spacecraft (approximately 100 mt) to cis-martian space. The crew makes use of torpor which minimizes launch payload mass. Green technologies are proposed as a stepping stone towards minimum environmental impact space access. The usage of beamed energy to power a grid of decentralized science stations is introduced, allowing for large scale characterization of the Martian environment. The low-thrust outbound and inbound trajectories are computed through the use of a direct method and a multiple shooting algorithm that considers various thrust and coast sequences to arrive at the final body with zero relative velocity. It is shown that the entire mission is rooted within the current NASA technology roadmap, ongoing scientific investments and feasible with an extrapolated NASA Budget. The presented mission won the 2016 Revolutionary Aerospace Systems Concepts - Academic Linkage (RASC-AL) competition.

  19. Electrostatic Plasma Accelerator (EPA)

    Brophy, John R.; Aston, Graeme


    The application of electric propulsion to communications satellites, however, has been limited to the use of hydrazine thrusters with electric heaters for thrust and specific impulse augmentation. These electrothermal thrusters operate at specific impulse levels of approximately 300 s with heater powers of about 500 W. Low power arcjets (1-3 kW) are currently being investigated as a way to increase specific impulse levels to approximately 500 s. Ion propulsion systems can easily produce specific impulses of 3000 s or greater, but have yet to be applied to communications satellites. The reasons most often given for not using ion propulsion systems are their high level of overall complexity, low thrust with long burn times, and the difficulty of integrating the propulsion system into existing commercial spacecraft busses. The Electrostatic Plasma Accelerator (EPA) is a thruster concept which promises specific impulse levels between low power arcjets and those of the ion engine while retaining the relative simplicity of the arcjet. The EPA thruster produces thrust through the electrostatic acceleration of a moderately dense plasma. No accelerating electrodes are used and the specific impulse is a direct function of the applied discharge voltage and the propellant atomic mass.

  20. Mars mission performance enhancement with hybrid nuclear propulsion

    Dagle, J. E. [Pacific Northwest Lab., Richland, WA (United States); Noffsinger, K. E. [Pacific Northwest Lab., Richland, WA (United States); Segna, D. R. [USDOE Richland Operations Office, WA (United States)


    Nuclear electric propulsion (NEP), compared with chemical and nuclear thermal propulsion (NTP), can effectively deliver the same mass to Mars using much less propellant, consequently requiring less mass delivered to Earth orbit. The lower thrust of NEP requires a spiral trajectory near planetary bodies, which significantly increases the travel time. Although the total travel time is long, the portion of the flight time spent during interplanetary transfer is shorter, because the vehicle is thrusting for much longer periods of time. This has led to the supposition that NEP, although very attractive for cargo missions, is not suitable for piloted missions to Mars. However, with the application of a hybrid application of a hybrid approach to propulsion, the benefits of NEP can be utilized while drastically reducing the overall travel time required. Development of a dual-mode system, which utilizes high-thrust NTP to propel the spacecraft from the planetary gravitational influence and low-thrust NEP to accelerate in interplanetary space, eliminates the spiral trajectory and results in a much faster transit time than could be obtained by either NEP or NTP alone. This results in a mission profile with a lower initial mass in low Earth orbit. In addition, the propulsion system would have the capability to provide electrical power for mission applications.

  1. The effect of radio-frequency self bias on ion acceleration in expanding argon plasmas in helicon sources

    Wiebold, Matthew D.

    Time-averaged plasma potential differences up to ˜ 165 V over several hundred Debye lengths are observed in low pressure (pn floating potential for argon (Vp ≈ 5kTe/e). In the capacitive mode, the ion acceleration is not well described by an ambipolar relation. The accelerated population decay is consistent with that predicted by charge-exchange collisions. Grounding the upstream endplate increases the self-bias voltage compared to a floating endplate. In the inductive and helicon modes, the ion acceleration more closely follows an ambipolar relation, a result of decreased capacitive coupling due to the decreased RF skin depth. The scaling of the potential gradient with the argon flow rate, magnetic field and RF power are investigated, with the highest potential gradients observed for the lowest flow rates in the capacitive mode. The magnitude of the self-bias voltage agrees well with that predicted for RF sheaths. Use of the self-bias effect in a plasma thruster is explored, possibly for a low thrust, high specific impulse mode in a multi-mode helicon thruster. This work could also explain similar potential gradients in expanding helicon plasmas that are ascribed to double layer formation in the literature.

  2. Aeroelastic modeling of composite rotor blades with straight and swept tips

    Yuan, Kuo-An; Friedmann, Peretz P.; Venkatesan, Comandur


    This paper presents an analytical study of the aeroelastic behavior of composite rotor blades with straight and swept tips. The blade is modeled by beam type finite elements. A single finite element is used to model the swept tip. The nonlinear equations of motion for the FEM are derived using Hamilton's principle and based on a moderate deflection theory and accounts for: arbitrary cross-sectional shape, pretwist, generally anisotropic material behavior, transverse shears and out-of-plane warping. Numerical results illustrating the effects of tip sweep, anhedral and composite ply orientation on blade aeroelastic behavior are presented. It is shown that composite ply orientation has a substantial effect on blade stability. At low thrust conditions, certain ply orientations can cause instability in the lag mode. The flap-torsion coupling associated with tip sweep can also induce aeroelastic instability in the blade. This instability can be removed by appropriate ply orientation in the composite construction. These results illustrate the inherent potential for aeroelastic tailoring present in composite rotor blades with swept tips, which still remains to be exploited in the design process.

  3. Measured Boundary Layer Transition and Rotor Hover Performance at Model Scale

    Overmeyer, Austin D.; Martin, Preston B.


    An experiment involving a Mach-scaled, 11:08 f t: diameter rotor was performed in hover during the summer of 2016 at NASA Langley Research Center. The experiment investigated the hover performance as a function of the laminar to turbulent transition state of the boundary layer, including both natural and fixed transition cases. The boundary layer transition locations were measured on both the upper and lower aerodynamic surfaces simultaneously. The measurements were enabled by recent advances in infrared sensor sensitivity and stability. The infrared thermography measurement technique was enhanced by a paintable blade surface heater, as well as a new high-sensitivity long wave infrared camera. The measured transition locations showed extensive amounts, x=c>0:90, of laminar flow on the lower surface at moderate to high thrust (CT=s > 0:068) for the full blade radius. The upper surface showed large amounts, x=c > 0:50, of laminar flow at the blade tip for low thrust (CT=s boundary layer transition models in CFD and rotor design tools. The data is expected to be used as part of the AIAA Rotorcraft SimulationWorking Group

  4. Affordable Development and Demonstration of a Small NTR engine and Stage: A Preliminary NASA, DOE, and Industry Assessment

    Borowski, S. K.; Sefcik, R. J.; Fittje, J. E.; McCurdy, D. R.; Qualls, A. L.; Schnitzler, B. G; Werner, J.; Weitzberg, A.; Joyner, C. R.


    In FY'11, Nuclear Thermal Propulsion (NTP) was identified as a key propulsion option under the Advanced In-Space Propulsion (AISP) component of NASA's Exploration Technology Development and Demonstration (ETDD) program A strategy was outlined by GRC and NASA HQ that included 2 key elements -"Foundational Technology Development" followed by specific "Technology Demonstration" projects. The "Technology Demonstration "element proposed ground technology demonstration (GTD) testing in the early 2020's, followed by a flight technology demonstration (FTD) mission by approx. 2025. In order to reduce development costs, the demonstration projects would focus on developing a small, low thrust (approx. 7.5 -16.5 klb(f)) engine that utilizes a "common" fuel element design scalable to the higher thrust (approx. 25 klb(f)) engines used in NASA's Mars DRA 5.0 study(NASA-SP-2009-566). Besides reducing development costs and allowing utilization of existing, flight proven engine hard-ware (e.g., hydrogen pumps and nozzles), small, lower thrust ground and flight demonstration engines can validate the technology and offer improved capability -increased payloads and decreased transit times -valued for robotic science missions identified in NASA's Decadal Study.

  5. Closed Loop Guidance Trade Study for Space Launch System Block-1B Vehicle

    Von der Porten, Paul; Ahmad, Naeem; Hawkins, Matt


    NASA is currently building the Space Launch System (SLS) Block-1 launch vehicle for the Exploration Mission 1 (EM-1) test flight. The design of the next evolution of SLS, Block-1B, is well underway. The Block-1B vehicle is more capable overall than Block-1; however, the relatively low thrust-to-weight ratio of the Exploration Upper Stage (EUS) presents a challenge to the Powered Explicit Guidance (PEG) algorithm used by Block-1. To handle the long burn durations (on the order of 1000 seconds) of EUS missions, two algorithms were examined. An alternative algorithm, OPGUID, was introduced, while modifications were made to PEG. A trade study was conducted to select the guidance algorithm for future SLS vehicles. The chosen algorithm needs to support a wide variety of mission operations: ascent burns to LEO, apogee raise burns, trans-lunar injection burns, hyperbolic Earth departure burns, and contingency disposal burns using the Reaction Control System (RCS). Additionally, the algorithm must be able to respond to a single engine failure scenario. Each algorithm was scored based on pre-selected criteria, including insertion accuracy, algorithmic complexity and robustness, extensibility for potential future missions, and flight heritage. Monte Carlo analysis was used to select the final algorithm. This paper covers the design criteria, approach, and results of this trade study, showing impacts and considerations when adapting launch vehicle guidance algorithms to a broader breadth of in-space operations.

  6. Powered Explicit Guidance Modifications and Enhancements for Space Launch System Block-1 and Block-1B Vehicles

    Von der Porten, Paul; Ahmad, Naeem; Hawkins, Matt; Fill, Thomas


    NASA is currently building the Space Launch System (SLS) Block-1 launch vehicle for the Exploration Mission 1 (EM-1) test flight. NASA is also currently designing the next evolution of SLS, the Block-1B. The Block-1 and Block-1B vehicles will use the Powered Explicit Guidance (PEG) algorithm (of Space Shuttle heritage) for closed loop guidance. To accommodate vehicle capabilities and design for future evolutions of SLS, modifications were made to PEG for Block-1 to handle multi-phase burns, provide PEG updated propulsion information, and react to a core stage engine out. In addition, due to the relatively low thrust-to-weight ratio of the Exploration Upper Stage (EUS) and EUS carrying out Lunar Vicinity and Earth Escape missions, certain enhancements to the Block-1 PEG algorithm are needed to perform Block-1B missions to account for long burn arcs and target translunar and hyperbolic orbits. This paper describes the design and implementation of modifications to the Block-1 PEG algorithm as compared to Space Shuttle. Furthermore, this paper illustrates challenges posed by the Block-1B vehicle and the required PEG enhancements. These improvements make PEG capable for use on the SLS Block-1B vehicle as part of the Guidance, Navigation, and Control (GN&C) System.

  7. A High Power Solar Electric Propulsion - Chemical Mission for Human Exploration of Mars

    Burke, Laura M.; Martini, Michael C.; Oleson, Steven R.


    Recently Solar Electric Propulsion (SEP) as a main propulsion system has been investigated as an option to support manned space missions to near-Earth destinations for the NASA Gateway spacecraft. High efficiency SEP systems are able to reduce the amount of propellant long duration chemical missions require, ultimately reducing the required mass delivered to Low Earth Orbit (LEO) by a launch vehicle. However, for long duration interplanetary Mars missions, using SEP as the sole propulsion source alone may not be feasible due to the long trip times to reach and insert into the destination orbit. By combining an SEP propulsion system with a chemical propulsion system the mission is able to utilize the high-efficiency SEP for sustained vehicle acceleration and deceleration in heliocentric space and the chemical system for orbit insertion maneuvers and trans-earth injection, eliminating the need for long duration spirals. By capturing chemically instead of with low-thrust SEP, Mars stay time increases by nearly 200 days. Additionally, the size the of chemical propulsion system can be significantly reduced from that of a standard Mars mission because the SEP system greatly decreases the Mars arrival and departure hyperbolic excess velocities (V(sub infinity)).

  8. Study and Developement of Compact Permanent Magnet Hall Thrusters for Future Brazillian Space Missions

    Ferreira, Jose Leonardo; Martins, Alexandre; Cerda, Rodrigo


    The Plasma Physics Laboratory of UnB has been developing a Permanent Magnet Hall Thruster (PHALL) for the UNIESPAÇO program, part of the Space Activities Program conducted by AEB- The Brazillian Space Agency since 2004. Electric propulsion is now a very successful method for primary and secondary propulsion systems. It is essential for several existing geostationary satellite station keeping systems and for deep space long duration solar system missions, where the thrusting system can be designed to be used on orbit transfer maneuvering and/or for satellite attitude control in long term space missions. Applications of compact versions of Permanent Magnet Hall Thrusters on future brazillian space missions are needed and foreseen for the coming years beginning with the use of small divergent cusp field (DCFH) Hall Thrusters type on CUBESATS ( 5-10 kg , 1W-5 W power consumption) and on Micro satellites ( 50- 100 kg, 10W-100W). Brazillian (AEB) and German (DLR) space agencies and research institutions are developing a new rocket dedicated to small satellite launching. The VLM- Microsatellite Launch Vehicle. The development of PHALL compact versions can also be important for the recently proposed SBG system, a future brazillian geostationary satellite system that is already been developed by an international consortium of brazillian and foreign space industries. The exploration of small bodies in the Solar System with spacecraft has been done by several countries with increasing frequency in these past twenty five years. Since their historical beginning on the sixties, most of the Solar System missions were based on gravity assisted trajectories very much depended on planet orbit positioning relative to the Sun and the Earth. The consequence was always the narrowing of the mission launch window. Today, the need for Solar System icy bodies in situ exploration requires less dependence on gravity assisted maneuvering and new high precision low thrust navigation methods

  9. Comparison of technologies for deorbiting spacecraft from low-earth-orbit at end of mission

    Sánchez-Arriaga, G.; Sanmartín, J. R.; Lorenzini, E. C.


    An analytical comparison of four technologies for deorbiting spacecraft from Low-Earth-Orbit at end of mission is presented. Basic formulas based on simple physical models of key figures of merit for each device are found. Active devices - rockets and electrical thrusters - and passive technologies - drag augmentation devices and electrodynamic tethers - are considered. A basic figure of merit is the deorbit device-to-spacecraft mass ratio, which is, in general, a function of environmental variables, technology development parameters and deorbit time. For typical state-of-the-art values, equal deorbit time, middle inclination and initial altitude of 850 km, the analysis indicates that tethers are about one and two orders of magnitude lighter than active technologies and drag augmentation devices, respectively; a tether needs a few percent mass-ratio for a deorbit time of a couple of weeks. For high inclination, the performance drop of the tether system is moderate: mass ratio and deorbit time increase by factors of 2 and 4, respectively. Besides collision risk with other spacecraft and system mass considerations, such as main driving factors for deorbit space technologies, the analysis addresses other important constraints, like deorbit time, system scalability, manoeuver capability, reliability, simplicity, attitude control requirement, and re-entry and multi-mission capability (deorbit and re-boost) issues. The requirements and constraints are used to make a critical assessment of the four technologies as functions of spacecraft mass and initial orbit (altitude and inclination). Emphasis is placed on electrodynamic tethers, including the latest advances attained in the FP7/Space project BETs. The superiority of tape tethers as compared to round and multi-line tethers in terms of deorbit mission performance is highlighted, as well as the importance of an optimal geometry selection, i.e. tape length, width, and thickness, as function of spacecraft mass and initial

  10. Forecasting the Impact of an 1859-calibre Superstorm on Satellite Resources

    Odenwald, Sten; Green, James; Taylor, William


    We have assembled a database of operational satellites in orbit as of 2004, and have developed a series of simple models to assess the economic impacts to this resource caused by various scenarios of superstorm events possible during the next sunspot cycle between 2010 and 2014. Despite the apparent robustness of our satellite assets against the kinds of storms we have encountered during the satellite era, our models suggest a potential economic loss exceeding $10(exp 11) for satellite replacement and lost profitability caused by a once a century single storm similar to the 1859 superstorm. From a combination of power system and attitude control system (the most vulnerable) failures, we estimate that 80 satellites (LEO, MEO, GEO) may be disabled as a consequence of a superstorm event. Additional consequences may include the failure of many of the GPS, GLONASS and Galileo satellite systems in MEO. Approximately 98 LEO satellites that normally would not have re-entered for many decades, may prematurely de-orbit in ca 2021 as a result of the temporarily increased atmospheric drag caused by the superstorm event occurring in 2012. The $10(exp 11) International Space Station may lose at least 15 kilometers of altitude, placing it in critical need for re-boosting by an amount that is potentially outside the range of typical Space Shuttle operations during the previous solar maximum in ca 2000, and at a time when NASA plans to decommission the Space Shuttle. Several LEO satellites will unexpectedly be placed on orbits that enter the ISS zone of avoidance, requiring some action by ground personnel and ISS astronauts to avoid close encounters. Radiation effects on astronauts have also been considered and could include a range of possibilities from acute radiation sickness for astronauts inside spacecraft, to near-lethal doses during EVAs. The specifics depends very sensitively on the spectral hardness of the accompanying SPE event. Currently, the ability to forecast extreme

  11. Non-Toxic Orbiter Maneuvering System (OMS) and Reaction Control System

    Hurlbert, Eric A.; Nicholson, Leonard S. (Technical Monitor)


    NASA is pursuing the technology and advanced development of a non-toxic (NT) orbital maneuvering system (OMS) and reaction control system (RCS) for shuttle upgrades, RLV, and reusable first stages. The primary objectives of the shuttle upgrades program are improved safety, improved reliability, reduced operations time and cost, improved performance or capabilities, and commonality with future space exploration needs. Non-Toxic OMS/RCS offers advantages in each of these categories. A non-toxic OMS/RCS eliminates the ground hazards and the flight safety hazards of the toxic and corrosive propellants. The cost savings for ground operations are over $24M per year for 7 flights, and the savings increase with increasing flight rate up to $44M per year. The OMS/RCS serial processing time is reduced from 65 days to 13 days. The payload capability can be increased up to 5100 Ibms. The non-toxic OMS/RCS also provides improved space station reboost capability up to 20 nautical miles over the current toxic system of 14 nautical miles. A NT OMS/RCS represents a clear advancement in the SOA over MMH/NTO. Liquid oxygen and ethanol are clean burning, high-density propellants that provide a high degree of commonality with other spacecraft subsystems including life support, power, and thermal control, and with future human exploration and development of space missions. The simple and reliable pressure-fed design uses sub-cooled liquid oxygen at 250 to 350 psia, which allows a propellant to remain cryogenic for longer periods of time. The key technologies are thermal insulation and conditioning techniques are used to maintain the sub-cooling. Phase I successfully defined the system architecture, designed an integrated OMS/RCS propellant tank, analyzed the feed system, built and tested the 870 lbf RCS thrusters, and tested the 6000 lbf OMS engine. Phase 11 is currently being planned for the development and test of full-scale prototype of the system in 1999 and 2000

  12. Design and "As Flown" Radiation Environments for Materials in Low Earth Orbit

    Minow, Joseph; McWilliams, Brett; Altstatt, Richard; Koontz, Steven


    the ISS flight trajectory including variations in altitude due to decay of the vehicle orbit and periodic reboosts to higher altitudes. In addition, an estimate of the AE-8 model to predict low Earth orbit electron flux (because the radiation dose for thin materials is dominated by the electron component of the radiation environment) is presented based on comparisons of the AE-8 model to measurements of electron integral flux at approximately 850 km from the Medium Energy Proton and Electron Detector on board the NOAA Polar Operational Environmental Satellite.

  13. Positive serum specific IgE has a short half-life in patients with penicillin allergy and reversal does not always indicate tolerance.

    Hjortlund, Janni; Mortz, Charlotte Gotthard; Stage, Tore Bjerregaard; Skov, Per Stahl; Dahl, Ronald; Bindslev-Jensen, Carsten


    The positive and negative predictive values of specific IgE to penicillins are not well established for penicillin hypersensitivity. One reason may be that serum IgE levels to penicillin diminish over time. The objective in this study was to investigate variations in serum half-life (T½) for specific IgE to penicillins (s-IgE) and to evaluate the outcome of penicillin challenges in patients with previous but not present specific IgE to penicillins. Two subgroups were investigated. All included patients had a history of penicillin allergy with reported symptoms such as urticaria/angioedema or unclassified cutaneous rash. T½ of specific IgE to penicillins was calculated based on sera from 29 patients with repeated measurements of s-IgE. Twenty-two patients with a previous positive s-IgE was followed and challenged with penicillin when IgE had become negative. The T½ for s-IgE varied between the 26 patients with decreasing s-IgE from 1.6 months to 76.4 months and 52% had a T½ of less than a year. The three patients with stable and increasing IgE-values showed T½ approaching infinity A total of 29 challenges with β-lactams were performed. Four different patterns were seen when evaluating the clinical reaction to challenge (positive/negative) and post-challenge boost of s-IgE (yes/no). Eight (36.4%) had negative challenge and negative post-challenge s-IgE, eight (36.4%) negative challenge, but positive post-challenge s-IgE levels. 3 (13.6%) had positive challenge and positive post-challenge s-IgE whereas 3 (13.6%) were challenge positive, but had negative post-challenge s-IgE. Specific IgE to penicillins declines over time stressing the importance of a close time relation between diagnostic work-up and clinical reaction. Reversal of previously positive s-IgE may still be associated with positive penicillin challenges and/or re-boostering of s-IgE to positivity.

  14. Preliminary Thermohydraulic Analysis of a New Moderated Reactor Utilizing an LEU-Fuel for Space Nuclear Thermal Propulsion

    Nam, Seung Hyun; Choi, Jae Young; Venneria, Paolo F.; Jeong, Yong Hoon; Chang, Soon Heung


    The Korea Advanced NUclear Thermal Engine Rocket utilizing an LEU fuel (KANUTER-LEU) is a non-proliferative and comparably efficient NTR engine with relatively low thrust levels of 40 - 50 kN for in-space transportation. The small modular engine can expand mission versatility, when flexibly used in a clustered engine arrangement, so that it can perform various scale missions from low-thrust robotic science missions to high-thrust manned missions. In addition, the clustered engine system can enhance engine redundancy and ensuing crew safety as well as the thrust. The propulsion system is an energy conversion system to transform the thermal energy of the reactor into the kinetic energy of the propellant to produce the powers for thrust, propellant feeding and electricity. It is mainly made up of a propellant Feeding System (PFS) comprising a Turbo-Pump Assembly (TPA), a Regenerative Nozzle Assembly (RNA), etc. For this core design study, an expander cycle is assumed to be the propulsion system. The EGS converts the thermal energy of the EHTGR in the idle operation (only 350 kW th power) to electric power during the electric power mode. This paper presents a preliminary thermohydraulic design analysis to explore the design space for the new reactor and to estimate the referential engine performance. The new non-proliferative NTR engine concept, KANUTER-LEU, is under designing to surmount the nuclear proliferation obstacles on allR and Dactivities and eventual commercialization for future generations. To efficiently implement a heavy LEU fuel for the NTR engine, its reactor design innovatively possesses the key characteristics of the high U density fuel with high heating and H 2 corrosion resistances, the thermal neutron spectrum core and also minimizing non-fission neutron loss, and the compact reactor design with protectively cooling capability. To investigate feasible design space for the moderated EHTGR-LEU and resultant engine performance, the preliminary design

  15. Preliminary Thermohydraulic Analysis of a New Moderated Reactor Utilizing an LEU-Fuel for Space Nuclear Thermal Propulsion

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


    The Korea Advanced NUclear Thermal Engine Rocket utilizing an LEU fuel (KANUTER-LEU) is a non-proliferative and comparably efficient NTR engine with relatively low thrust levels of 40 - 50 kN for in-space transportation. The small modular engine can expand mission versatility, when flexibly used in a clustered engine arrangement, so that it can perform various scale missions from low-thrust robotic science missions to high-thrust manned missions. In addition, the clustered engine system can enhance engine redundancy and ensuing crew safety as well as the thrust. The propulsion system is an energy conversion system to transform the thermal energy of the reactor into the kinetic energy of the propellant to produce the powers for thrust, propellant feeding and electricity. It is mainly made up of a propellant Feeding System (PFS) comprising a Turbo-Pump Assembly (TPA), a Regenerative Nozzle Assembly (RNA), etc. For this core design study, an expander cycle is assumed to be the propulsion system. The EGS converts the thermal energy of the EHTGR in the idle operation (only 350 kW{sub th} power) to electric power during the electric power mode. This paper presents a preliminary thermohydraulic design analysis to explore the design space for the new reactor and to estimate the referential engine performance. The new non-proliferative NTR engine concept, KANUTER-LEU, is under designing to surmount the nuclear proliferation obstacles on allR and Dactivities and eventual commercialization for future generations. To efficiently implement a heavy LEU fuel for the NTR engine, its reactor design innovatively possesses the key characteristics of the high U density fuel with high heating and H{sub 2} corrosion resistances, the thermal neutron spectrum core and also minimizing non-fission neutron loss, and the compact reactor design with protectively cooling capability. To investigate feasible design space for the moderated EHTGR-LEU and resultant engine performance, the

  16. Mars Sample Return Using Solar Sail Propulsion

    Johnson, Les; Macdonald, Malcolm; Mcinnes, Colin; Percy, Tom


    Many Mars Sample Return (MSR) architecture studies have been conducted over the years. A key element of them is the Earth Return Stage (ERS) whose objective is to obtain the sample from the Mars Ascent Vehicle (MAV) and return it safely to the surface of the Earth. ERS designs predominantly use chemical propulsion [1], incurring a significant launch mass penalty due to the low specific impulse of such systems coupled with the launch mass sensitivity to returned mass. It is proposed to use solar sail propulsion for the ERS, providing a high (effective) specific impulse propulsion system in the final stage of the multi-stage system. By doing so to the launch mass of the orbiter mission can be significantly reduced and hence potentially decreasing mission cost. Further, solar sailing offers a unique set of non-Keplerian low thrust trajectories that may enable modifications to the current approach to designing the Earth Entry Vehicle by potentially reducing the Earth arrival velocity. This modification will further decrease the mass of the orbiter system. Solar sail propulsion uses sunlight to propel vehicles through space by reflecting solar photons from a large, mirror-like surface made of a lightweight, reflective material. The continuous photonic pressure provides propellantless thrust to conduct orbital maneuvering and plane changes more efficiently than conventional chemical propulsion. Because the Sun supplies the necessary propulsive energy, solar sails require no onboard propellant, thus reducing system mass. This technology is currently at TRL 7/8 as demonstrated by the 2010 flight of the Japanese Aerospace Exploration Agency, JAXA, IKAROS mission. [2

  17. Control Relevant Modeling and Design of Scramjet-Powered Hypersonic Vehicles

    Dickeson, Jeffrey James

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

  18. Development of solid-gas equilibrium propulsion system for small spacecraft

    Chujo, Toshihiro; Mori, Osamu; Kubo, Yuki


    A phase equilibrium propulsion system is a kind of cold-gas jet in which the phase equilibrium state of the fuel is maintained in a tank and its vapor is ejected when a valve is opened. One such example is a gas-liquid equilibrium propulsion system that uses liquefied gas as fuel. This system was mounted on the IKAROS solar sail and has been demonstrated in orbit. The system has a higher storage efficiency and a lighter configuration than a high-pressure cold-gas jet because the vapor pressure is lower, and is suitable for small spacecraft. However, the system requires a gas-liquid separation device in order to avoid leakage of the liquid, which makes the system complex. As another example of a phase equilibrium propulsion system, we introduce a solid-gas equilibrium propulsion system, which uses a sublimable substance as fuel and ejects its vapor. This system has an even lower vapor pressure and does not require such a separation device, instead requiring only a filter to keep the solid inside the tank. Moreover, the system is much simpler and lighter, making it more suitable for small spacecraft, especially CubeSat-class spacecraft, and the low thrust of the system allows spacecraft motion to be controlled precisely. In addition, the thrust level can be controlled by controlling the temperature of the fuel, which changes the vapor pressure. The present paper introduces the concept of the proposed system, and describes ejection experiments and its evaluation. The basic function of the proposed system is demonstrated in order to verify its usefulness.

  19. Using Gravity Assists in the Earth-moon System as a Gateway to the Solar System

    McElrath, Tim; Lantoine, Gregory; Landau, Damon; Grebow, Dan; Strange, Nathan; Wilson, Roby; Sims, Jon


    For spacecraft departing the Earth - Moon system, lunar flybys can significantly increase the hype rbolic escape energy (C3, in km 2 /sec 2 ) for a modest increase in flight time. Within 2 months, lunar flybys can produce a C3 of 2. Over 4 - 6 months, lunar flybys alone can increase the C3 to 4.5, or they can provide for additional periapsis burns to increase the C3 from 2 -3 to 10 or more, suitable for planetary missions. A lunar flyby departure can be followed by additional ? -V (such as that efficiently provided by a low thrust system, eg. Solar Electric Propulsion (SEP)) to raise the Earth - relative velocity (at a ratio of more than 2:1) before a subsequent Earth flyby, which redirects that velocity to a more di stant target, all within not much more than a year. This paper describes the applicability of lunar flybys for different flight times and propulsi on systems, and illustrates this with instances of past usage and future possibilities. Examples discussed i nclude ISEE - 3, Nozomi, STEREO, 2018 Mars studies (which showed an 8% payload increase), and missions to Near Earth Objects (NEOs). In addition, the options for the achieving the initial lunar flyby are systematically discussed, with a view towards their p ractical use with in a compact launch period. In particular, we show that launches to geosynchronous transfer orbit (GTO) as a secondary payload provide a feasible means of obtaining a lunar flyby for an acceptable cost, even for SEP systems that cannot ea sily deliver large ? - Vs at periapsis. Taken together, these results comprise a myriad of options for increasing the mission performance, by the efficient use of lunar flybys within an acceptable extension of the flight time.

  20. Escape manoeuvres in the spiny dogfish (Squalus acanthias).

    Domenici, Paolo; Standen, Emily M; Levine, Robert P


    The locomotor performance of dogfish during escape responses was observed by means of high-speed video. Dogfish show C-type escape responses that are comparable with those shown previously in teleosts. Dogfish show high variability of turning rates of the anterior part of the body (head to centre of mass), i.e. with peak values from 434 to 1023 deg. s(-1). We suggest that this variability may be due to the presence of two types of escape manoeuvres, i.e. responses with high and low turning rates, as previously found in a teleost species. Fast responses (i.e. with high maximum turning rates, ranging between 766 and 1023 deg. s(-1)) showed significantly higher locomotor performance than slow responses (i.e. with low maximum turning rates, ranging between 434 and 593 deg. s(-1)) in terms of distance covered, speed and acceleration, although no differences were found in the turning radius of the centre of mass during the escape manoeuvres. The existence of two types of escape responses would have implications in terms of both neural control and muscular activation patterns. When compared with literature data for the locomotor performance of bony fishes, dogfish showed relatively low speed and acceleration, comparable turning rates and a turning radius that is in the low part of the range when compared with teleosts, indicating relatively high manoeuvrability. The locomotor performance observed in dogfish is consistent with their morphological characteristics: (1) low locomotor performance associated with low thrust developed by their relatively small posterior depth of section and (2) relatively high manoeuvrability associated with their high flexibility.

  1. Fiber-Optic Continuous Liquid Sensor for Cryogenic Propellant Gauging

    Xu. Wei


    An innovative fiber-optic sensor has been developed for low-thrust-level settled mass gauging with measurement uncertainty optical fiber to measure liquid level and liquid distribution of cryogenic propellants. Every point of the sensing fiber is a point sensor that not only distinguishes liquid and vapor, but also measures temperature. This sensor is able to determine the physical location of each point sensor with 1-mm spatial resolution. Acting as a continuous array of numerous liquid/vapor point sensors, the truly distributed optical sensing fiber can be installed in a propellant tank in the same manner as silicon diode point sensor stripes using only a single feedthrough to connect to an optical signal interrogation unit outside the tank. Either water or liquid nitrogen levels can be measured within 1-mm spatial resolution up to a distance of 70 meters from the optical interrogation unit. This liquid-level sensing technique was also compared to the pressure gauge measurement technique in water and liquid nitrogen contained in a vertical copper pipe with a reasonable degree of accuracy. It has been demonstrated that the sensor can measure liquid levels in multiple containers containing water or liquid nitrogen with one signal interrogation unit. The liquid levels measured by the multiple fiber sensors were consistent with those virtually measured by a ruler. The sensing performance of various optical fibers has been measured, and has demonstrated that they can survive after immersion at cryogenic temperatures. The fiber strength in liquid nitrogen has also been measured. Multiple water level tests were also conducted under various actual and theoretical vibration conditions, and demonstrated that the signal-to-noise ratio under these vibration conditions, insofar as it affects measurement accuracy, is manageable and robust enough for a wide variety of spacecraft applications. A simple solution has been developed to absorb optical energy at the termination of

  2. Computational Fluid Dynamics Analysis Method Developed for Rocket-Based Combined Cycle Engine Inlet


    Renewed interest in hypersonic propulsion systems has led to research programs investigating combined cycle engines that are designed to operate efficiently across the flight regime. The Rocket-Based Combined Cycle Engine is a propulsion system under development at the NASA Lewis Research Center. This engine integrates a high specific impulse, low thrust-to-weight, airbreathing engine with a low-impulse, high thrust-to-weight rocket. From takeoff to Mach 2.5, the engine operates as an air-augmented rocket. At Mach 2.5, the engine becomes a dual-mode ramjet; and beyond Mach 8, the rocket is turned back on. One Rocket-Based Combined Cycle Engine variation known as the "Strut-Jet" concept is being investigated jointly by NASA Lewis, the U.S. Air Force, Gencorp Aerojet, General Applied Science Labs (GASL), and Lockheed Martin Corporation. Work thus far has included wind tunnel experiments and computational fluid dynamics (CFD) investigations with the NPARC code. The CFD method was initiated by modeling the geometry of the Strut-Jet with the GRIDGEN structured grid generator. Grids representing a subscale inlet model and the full-scale demonstrator geometry were constructed. These grids modeled one-half of the symmetric inlet flow path, including the precompression plate, diverter, center duct, side duct, and combustor. After the grid generation, full Navier-Stokes flow simulations were conducted with the NPARC Navier-Stokes code. The Chien low-Reynolds-number k-e turbulence model was employed to simulate the high-speed turbulent flow. Finally, the CFD solutions were postprocessed with a Fortran code. This code provided wall static pressure distributions, pitot pressure distributions, mass flow rates, and internal drag. These results were compared with experimental data from a subscale inlet test for code validation; then they were used to help evaluate the demonstrator engine net thrust.

  3. Enhanced Gravity Tractor Technique for Planetary Defense

    Mazanek, Daniel D.; Reeves, David M.; Hopkins, Joshua B.; Wade, Darren W.; Tantardini, Marco; Shen, Haijun


    Given sufficient warning time, Earth-impacting asteroids and comets can be deflected with a variety of different "slow push/pull" techniques. The gravity tractor is one technique that uses the gravitational attraction of a rendezvous spacecraft to the impactor and a low-thrust, high-efficiency propulsion system to provide a gradual velocity change and alter its trajectory. An innovation to this technique, known as the Enhanced Gravity Tractor (EGT), uses mass collected in-situ to augment the mass of the spacecraft, thereby greatly increasing the gravitational force between the objects. The collected material can be a single boulder, multiple boulders, regolith or a combination of different sources. The collected mass would likely range from tens to hundreds of metric tons depending on the size of the impactor and warning time available. Depending on the propulsion system's capability and the mass collected, the EGT approach can reduce the deflection times by a factor of 10 to 50 or more, thus reducing the deflection times of several decades to years or less and overcoming the main criticism of the traditional gravity tractor approach. Additionally, multiple spacecraft can orbit the target in formation to provide the necessary velocity change and further reduce the time needed by the EGT technique to divert hazardous asteroids and comets. The robotic segment of NASA's Asteroid Redirect Mission (ARM) will collect a multi-ton boulder from the surface of a large Near-Earth Asteroid (NEA) and will provide the first ever demonstration of the EGT technique and validate one method of collecting in-situ mass on an asteroid of hazardous size.

  4. An Overview of the CNES Propulsion Program for Spacecraft

    Cadiou, A.; Darnon, F.; Gibek, I.; Jolivet, L.; Pillet, N.


    This paper presents an overview of the CNES spacecraft propulsion activities. The main existing and future projects corresponding to low earth orbit and geostationary platforms are described. These projects cover various types of propulsion subsystems: monopropellant, bipropellant and electric. Monopropellant is mainly used for low earth orbit applications such as earth observation (SPOT/Helios, PLEIADES) or scientific applications (minisatellite PROTEUS line and micro satellites MYRIADE line). Bipropellant is used for geostationary telecommunications satellites (@BUS). The field of application of electric propulsion is the station keeping of geostationary telecommunication satellites (@BUS), main propulsion for specific probes (SMART 1) and fine attitude control for dedicated micro satellites (MICROSCOPE). The preparation of the future and the associated Research and Technology program are also described in the paper. The future developments are mainly dedicated to the performance improvements of electric propulsion which leads to the development of thrusters with higher thrust and higher specific impulse than those existing today, the evaluation of the different low thrust technologies for formation flying applications, the development of new systems to pressurize the propellants (volatile liquid, micro pump), the research on green propellants and different actions concerning components such as over wrapped pressure vessels, valves, micro propulsion. A constant effort is also put on plume effect in chemical and electrical propulsion area (improvement of tools and test activities) in the continuity of the previous work. These different R &T activities are described in detail after a presentation of the different projects and of their propulsion subsystems. The scientific activity supporting the development of Hall thrusters is going on in the frame of the GDR (Groupement de Recherche) CNRS / Universities / CNES / SNECMA on Plasma Propulsion.

  5. 'Bimodal' Nuclear Thermal Rocket (BNTR) propulsion for an artificial gravity HOPE mission to Callisto

    Borowski, Stanley K.; McGuire, Melissa L.; Mason, Lee M.; Gilland, James H.; Packard, Thomas W.


    This paper summarizes the results of a year long, multi-center NASA study which examined the viability of nuclear fission propulsion systems for Human Outer Planet Exploration (HOPE). The HOPE mission assumes a crew of six is sent to Callisto. Jupiter's outermost large moon, to establish a surface base and propellant production facility. The Asgard asteroid formation, a region potentially rich in water-ice, is selected as the landing site. High thrust BNTR propulsion is used to transport the crew from the Earth-Moon L1 staging node to Callisto then back to Earth in less than 5 years. Cargo and LH2 'return' propellant for the piloted Callisto transfer vehicle (PCTV) is pre-deployed at the moon (before the crew's departure) using low thrust, high power, nuclear electric propulsion (NEP) cargo and tanker vehicles powered by hydrogen magnetoplasmadynamic (MPD) thrusters. The PCTV is powered by three 25 klbf BNTR engines which also produce 50 kWe of power for crew life support and spacecraft operational needs. To counter the debilitating effects of long duration space flight (∼855 days out and ∼836 days back) under '0-gE' conditions, the PCTV generates an artificial gravity environment of '1-gE' via rotation of the vehicle about its center-of-mass at a rate of ∼4 rpm. After ∼123 days at Callisto, the 'refueled' PCTV leaves orbit for the trip home. Direct capsule re-entry of the crew at mission end is assumed. Dynamic Brayton power conversion and high temperature uranium dioxide (UO2) in tungsten metal ''cermet'' fuel is used in both the BNTR and NEP vehicles to maximize hardware commonality. Technology performance levels and vehicle characteristics are presented, and requirements for PCTV reusability are also discussed

  6. Radioisotope electric propulsion of sciencecraft to the outer Solar System and near-interstellar space

    Noble, R.J.


    Radioisotopes have been used successfully for more than 25 years to supply the heat for thermoelectric generators on various deep-space probes. Radioisotope electric propulsion (REP) systems have been proposed as low-thrust ion propulsion units based on radioisotope electric generators and ion thrusters. The perceived liability of radioisotope electric generators for ion propulsion is their high mass. Conventional radioisotope thermoelectric generators have a specific mass of about 200 kg/kW of electric power. Many development efforts have been undertaken with the aim of reducing the specific mass of radioisotope electric systems. Recent performance estimates suggest that specific masses of 50 kg/kW may be achievable with thermophotovoltaic and alkali metal thermal-to-electric conversion generators. Powerplants constructed from these near-term radioisotope electric generators and long-life ion thrusters will likely have specific masses in the range of 100 to 200 kg/kW of thrust power if development continues over the next decade. In earlier studies, it was concluded that flight times within the Solar System are indeed insensitive to reductions in the powerplant specific mass, and that a timely scientific program of robotic planetary rendezvous and near-interstellar space missions is enabled by primary electric propulsion once the powerplant specific mass is in the range of 100 to 200 kg/kW. Flight times can be substantially reduced by using hybrid propulsion schemes that combine chemical propulsion, gravity assist, and electric propulsion. Hybrid schemes are further explored in this article to illustrate how the performance of REP is enhanced for Pluto rendezvous, heliopause orbiter, and gravitational lens missions

  7. Characterizing Longitude-Dependent Orbital Debris Congestion in the Geosynchronous Orbit Regime

    Anderson, Paul V.

    The geosynchronous orbit (GEO) is a unique commodity of the satellite industry that is becoming increasingly contaminated with orbital debris, but is heavily populated with high-value assets from the civil, commercial, and defense sectors. The GEO arena is home to hundreds of communications, data transmission, and intelligence satellites collectively insured for an estimated 18.3 billion USD. As the lack of natural cleansing mechanisms at the GEO altitude renders the lifetimes of GEO debris essentially infinite, conjunction and risk assessment must be performed to safeguard operational assets from debris collisions. In this thesis, longitude-dependent debris congestion is characterized by predicting the number of near-miss events per day for every longitude slot at GEO, using custom debris propagation tools and a torus intersection metric. Near-miss events with the present-day debris population are assigned risk levels based on GEO-relative position and speed, and this risk information is used to prioritize the population for debris removal target selection. Long-term projections of debris growth under nominal launch traffic, mitigation practices, and fragmentation events are also discussed, and latitudinal synchronization of the GEO debris population is explained via node variations arising from luni-solar gravity. In addition to characterizing localized debris congestion in the GEO ring, this thesis further investigates the conjunction risk to operational satellites or debris removal systems applying low-thrust propulsion to raise orbit altitude at end-of-life to a super-synchronous disposal orbit. Conjunction risks as a function of thrust level, miss distance, longitude, and semi-major axis are evaluated, and a guidance method for evading conjuncting debris with continuous thrust by means of a thrust heading change via single-shooting is developed.

  8. Affordable Development and Demonstration of a Small NTR Engine and Stage: A Preliminary NASA, DOE, and Industry Assessment

    Borowski, Stanley K.; Sefcik, Robert J.; Fittje, James E.; McCurdy, David R.; Qualls, Arthur L.; Schnitzler, Bruce G.; Werner, James E.; Weitzberg, Abraham; Joyner, Claude R.


    The Nuclear Thermal Rocket (NTR) represents the next evolutionary step in cryogenic liquid rocket engines. Deriving its energy from fission of uranium-235 atoms contained within fuel elements that comprise the engine's reactor core, the NTR can generate high thrust at a specific impulse of approx. 900 seconds or more - twice that of today's best chemical rockets. In FY'11, as part of the AISP project, NASA proposed a Nuclear Thermal Propulsion (NTP) effort that envisioned two key activities - "Foundational Technology Development" followed by system-level "Technology Demonstrations". Five near-term NTP activities identified for Foundational Technology Development became the basis for the NCPS project started in FY'12 and funded by NASA's AES program. During Phase 1 (FY'12-14), the NCPS project was focused on (1) Recapturing fuel processing techniques and fabricating partial length "heritage" fuel elements for the two candidate fuel forms identified by NASA and the DOE - NERVA graphite "composite" and the uranium dioxide (UO2) in tungsten "cermet". The Phase 1 effort also included: (2) Engine Conceptual Design; (3) Mission Analysis and Requirements Definition; (4) Identification of Affordable Options for Ground Testing; and (5) Formulation of an Affordable and Sustainable NTP Development Strategy. During FY'14, a preliminary plan for DDT&E was outlined by GRC, the DOE and industry for NASA HQ that involved significant system-level demonstration projects that included GTD tests at the NNSS, followed by a FTD mission. To reduce development costs, the GTD and FTD tests use a small, low thrust (approx. 7.5 or 16.5 klbf) engine. Both engines use graphite composite fuel and a "common" fuel element design that is scalable to higher thrust (approx. 25 klbf) engines by increasing the number of elements in a larger diameter core that can produce greater thermal power output. To keep the FTD mission cost down, a simple "1-burn" lunar flyby mission was considered along with

  9. Embedded model control GNC for the Next Generation Gravity Mission

    Colangelo, Luigi; Massotti, Luca; Canuto, Enrico; Novara, Carlo


    role of a wide-band inner loop, and orbit/formation and attitude/pointing controls are the narrow band outer loops. The relevant state equations were converted to discrete time providing the embedded model, a fundamental part of the control unit. The state predictor, control law and reference generator were built on and interfaced to the embedded model. Simulated results, via a high-fidelity simulator, prove the concept validity and show that the control performances are in agreement with the defined mission requirements. Indeed, the presented control strategy is shown to be capable of keeping the attitude and formation variables stable within the required boundaries, all over the 10-year mission, through a low-thrust authority in the order of a few milli-Newton.

  10. Full-Span Tiltrotor Aeroacoustic Model (TRAM) Overview and 40- by 80-Foot Wind Tunnel Test. [conducted in the 40- by 80-Foot Wind Tunnel at Ames Research Center

    McCluer, Megan S.; Johnson, Jeffrey L.; Rutkowski, Michael (Technical Monitor)


    a variety of tunnel speeds. Wake geometry images were acquired using LLS photographs and suggest dual tip vortex formation at low thrust conditions. The full paper will include comparisons to isolated-rotor TRAM data acquired at the Duits-Nederlandse Windtunnel (DNW) in 1998. The FS TRAM has been established as a valuable national asset for tiltrotor research. Data reduction and analysis of the 40- by 80-Foot Wind Tunnel test results are underway. Follow-on testing of the FS TRAM is currently being planned for the NASA Ames 80- by 120-Foot Wind Tunnel in late 2001.

  11. Evaluation of vertical profiles to design continuous descent approach procedure

    Pradeep, Priyank

    The current research focuses on predictability, variability and operational feasibility aspect of Continuous Descent Approach (CDA), which is among the key concepts of the Next Generation Air Transportation System (NextGen). The idle-thrust CDA is a fuel economical, noise and emission abatement procedure, but requires increased separation to accommodate for variability and uncertainties in vertical and speed profiles of arriving aircraft. Although a considerable amount of researches have been devoted to the estimation of potential benefits of the CDA, only few have attempted to explain the predictability, variability and operational feasibility aspect of CDA. The analytical equations derived using flight dynamics and Base of Aircraft and Data (BADA) Total Energy Model (TEM) in this research gives insight into dependency of vertical profile of CDA on various factors like wind speed and gradient, weight, aircraft type and configuration, thrust settings, atmospheric factors (deviation from ISA (DISA), pressure and density of the air) and descent speed profile. Application of the derived equations to idle-thrust CDA gives an insight into sensitivity of its vertical profile to multiple factors. This suggests fixed geometric flight path angle (FPA) CDA has higher degree of predictability and lesser variability at the cost of non-idle and low thrust engine settings. However, with optimized design this impact can be overall minimized. The CDA simulations were performed using Future ATM Concept Evaluation Tool (FACET) based on radar-track and aircraft type data (BADA) of the real air-traffic to some of the busiest airports in the USA (ATL, SFO and New York Metroplex (JFK, EWR and LGA)). The statistical analysis of the vertical profiles of CDA shows 1) mean geometric FPAs derived from various simulated vertical profiles are consistently shallower than 3° glideslope angle and 2) high level of variability in vertical profiles of idle-thrust CDA even in absence of

  12. Syn- and Post-Accretionary Structures in the Neoproterozoic Central Allaqi-Heiani Suture Zone, Southeastern Egypt

    Abdeen, M. M.; Abdelghaffar, A. A.


    The Allaqi-Heiani suture (AHS) is the western part of the main Allaqi-Heiani-Gerf-Onib-Sol Hamed-Yanbu suture and represents one of the Neoproterozoic, arc-arc sutures in the Arabian-Nubian Shield (ANS). It separates the ca. 750 Ma South Eastern Desert terrane in the north from the ca. 830-720 Ma Gabgaba terrane in the south. The AHS is a deformed belt of ophiolitic rocks, syn-tectonic granitoids and metasediments. The central AHS zone is divided into three structural domains. The western domain (Ι) is characterized by NNE low thrusts and SSW-vergent folds. The central domain (ΙΙ) includes upright tight to isoclinal NNW-SSE oriented folds and transpressional faults. The eastern domain (ΙΙΙ) shows NNW-SSE oriented open folds. Structural analysis indicates that the area has a poly-phase deformation history involving at least two events. Event D1 was an N-S to NNE-SSW regional shortening generating the SSW-verging folds and the NNE dipping thrusts. Event D2 was an ENE-WSW shortening producing NNW-SSE oriented folds in the central and eastern parts of the study area and reactivating older thrusts with oblique-slip reverse fault movement. The tectonic evolution of the area involves two episodes of collision: an early collision between the South Eastern Desert terrane and the Gabgaba terrane along the AHS after the consumption of a basin floored by oceanic crust above a north-dipping subduction zone; and a later collision between East- and West-Gondwanas at ca. 750-650 Ma, leading to the closure of the Mozambique Ocean. This collision deformed the AHS along N-S trending shortening zones and produced NW-SE and NE-SW oriented sinistral and dextral transpressional faults, respectively. The early collision episode is related to the terrane accretion during the early Pan-African orogen, while the later phase is related to a late Pan-African or Najd orogen.

  13. Development and Qualification of ATV Propulsion Assemblies

    Riehle, M.; Jost, R.


    In the frame of the development and operation of the International Space Station ISS, the European Space Agency ESA is not only contributing experiments and a laboratory module but also logistics capacity. This purpose of supplying the ISS shall be covered by an unmanned, Automated Transfer Vehicle (ATV) that will be launched for the first time in 2004 by Ariane 5. The development of the ATV is in close conjunction to the future Ariane 5 launch capacity of about 20 tons injected into low earth orbit. Thus this unmanned transporter will be a quite large space craft that is subjected to fulfil several mission objectives apart of only delivering cargo such as multiple automatic docking/de-docking, re-boost services and re-fuelling. For those reasons and due to its dimensions the propulsion sub-system is one of the most sophisticated in the field of space propulsion. Even safety issues of manned space flight have to be applied since the pressurised cargo section is part of the ISS when docked to the manned modules. This leads to by far the largest but also the most sophisticated propulsion system ever built in Europe. Astrium as one of the major partners of this european project is responsible for this major system that will be described in the paper. Focusing on the major core assemblies such as multi thruster platforms, pressure control system incl. safety and redundancy mechanisms as well as tanks and other components that completes a propulsion system. System Design and Qualification Starting from the basic criteria the paper will present the major performance requirements such as pressures, thrust levels and other parameters that led to the selection of major components of the system such as thrusters, valves, tanks, etc. Some of the component could be selected from off the shelve, whereas other core components such as the 200N Attitude Control and Braking Thrusters or Propellant Tanks had to be newly developed. The stepwise approach of development and careful

  14. A mars communication constellation for human exploration and network science

    Castellini, Francesco; Simonetto, Andrea; Martini, Roberto; Lavagna, Michèle


    This paper analyses the possibility of exploiting a small spacecrafts constellation around Mars to ensure a complete and continuous coverage of the planet, for the purpose of supporting future human and robotic operations and taking advantage of optical transmission techniques. The study foresees such a communications mission to be implemented at least after 2020 and a high data-rate requirement is imposed for the return of huge scientific data from massive robotic exploration or to allow video transmissions from a possible human outpost. In addition, the set-up of a communication constellation around Mars would give the opportunity of exploiting this multi-platform infrastructure to perform network science, that would largely increase our knowledge of the planet. The paper covers all technical aspects of a feasibility study performed for the primary communications mission. Results are presented for the system trade-offs, including communication architecture, constellation configuration and transfer strategy, and the mission analysis optimization, performed through the application of a multi-objective genetic algorithm to two models of increasing difficulty for the low-thrust trajectory definition. The resulting communication architecture is quite complex and includes six 530 kg spacecrafts on two different orbital planes, plus one redundant unit per plane, that ensure complete coverage of the planet’s surface; communications between the satellites and Earth are achieved through optical links, that allow lower mass and power consumption with respect to traditional radio-frequency technology, while inter-satellite links and spacecrafts-to-Mars connections are ensured by radio transmissions. The resulting data-rates for Earth-Mars uplink and downlink, satellite-to-satellite and satellite-to-surface are respectively 13.7 Mbps, 10.2 Mbps, 4.8 Mbps and 4.3 Mbps, in worst-case. Two electric propulsion modules are foreseen, to be placed on a C3˜0 escape orbit with two

  15. Status and Perspectives of Electric Propulsion in Italy

    Svelto, F.; Marcuccio, S.; Matticari, G.


    Electric Propulsion (EP) is recognized as one of today's enabling technologies for scientific and commercial missions. In consideration of EP's major strategic impact on the near and long term scenarios, an EP development programme has been established within the Italian Space Agency (ASI), aimed at the development of a variety of propulsion capabilities covering different fields of application. This paper presents an overview of Electric Propulsion (EP) activities underway in Italy and outlines the planned development lines, both in research institutions and in industry. Italian EP activities are essentially concentrated in Pisa, at Centrospazio and Alta, and in Florence, at LABEN - Proel Tecnologie Division (LABEN/Proel). Centrospazio/Alta and LABEN/Proel have established a collaboration program for joint advanced developments in the EP field. Established in 1989, Centrospazio is a private research center closely related to the Department of Aerospace Engineering of Pisa University. Along the years, Centrospazio lines of development have included arcjets, magneto- plasma-dynamic thrusters, FEEP and Hall thrusters, as well as computational plasma dynamics and low-thrust mission studies. Alta, a small enterprise, was founded in 1999 to exploit in an industrial setting the results of research previously carried out at Centrospazio. Alta's activities include the development of micronewton and millinewton FEEP thrusters, and testing of high power Hall and ion thrusters in specialised facilities. A full micronewton FEEP propulsion system is being developed for the Microscope spacecraft, a scientific mission by CNES aimed at verification of the Equivalence Principle. FEEP will also fly on ASI's HypSEO, a technological demonstrator for Earth Observation, and is being considered for ESA's GOCE (geodesy) and SMART-2 (formation flying), as well as for the intended scientific spacecraft GG by ASI. The ASI-funded STEPS facility will be placed on an external site on the

  16. Solar-Thermal Engine Testing

    Tucker, Stephen; Salvail, Pat; Haynes, Davy (Technical Monitor)


    A solar-thermal engine serves as a high-temperature solar-radiation absorber, heat exchanger, and rocket nozzle. collecting concentrated solar radiation into an absorber cavity and transferring this energy to a propellant as heat. Propellant gas can be heated to temperatures approaching 4,500 F and expanded in a rocket nozzle, creating low thrust with a high specific impulse (I(sub sp)). The Shooting Star Experiment (SSE) solar-thermal engine is made of 100 percent chemical vapor deposited (CVD) rhenium. The engine 'module' consists of an engine assembly, propellant feedline, engine support structure, thermal insulation, and instrumentation. Engine thermal performance tests consist of a series of high-temperature thermal cycles intended to characterize the propulsive performance of the engines and the thermal effectiveness of the engine support structure and insulation system. A silicone-carbide electrical resistance heater, placed inside the inner shell, substitutes for solar radiation and heats the engine. Although the preferred propellant is hydrogen, the propellant used in these tests is gaseous nitrogen. Because rhenium oxidizes at elevated temperatures, the tests are performed in a vacuum chamber. Test data will include transient and steady state temperatures on selected engine surfaces, propellant pressures and flow rates, and engine thrust levels. The engine propellant-feed system is designed to Supply GN2 to the engine at a constant inlet pressure of 60 psia, producing a near-constant thrust of 1.0 lb. Gaseous hydrogen will be used in subsequent tests. The propellant flow rate decreases with increasing propellant temperature, while maintaining constant thrust, increasing engine I(sub sp). In conjunction with analytical models of the heat exchanger, the temperature data will provide insight into the effectiveness of the insulation system, the structural support system, and the overall engine performance. These tests also provide experience on operational

  17. The Near Earth Object Scout Spacecraft: A Low Cost Approach to in-situ Characterization of the NEO Population

    Koontz, Steven L.; Condon, Gerald; Graham, Lee; Bevilacqua, Ricardo


    In this paper we describe a micro/nano satellite spacecraft and a supporting mission profile and architecture designed to enable preliminary in-situ characterization of a significant number of Near Earth Objects (NEOs) at reasonable cost. The spacecraft will be referred to as the NEO Scout. NEO Scout spacecraft are to be placed in GTO, GEO, or cis-lunar space as secondary payloads on launch vehicles headed for GTO or beyond and will begin their mission after deployment from the launcher. A distinguishing key feature of the NEO scout system is to design the mission timeline and spacecraft to rendezvous with and land on the target NEOs during close approach to the Earth-Moon system using low-thrust/high- impulse propulsion systems. Mission feasibility and preliminary design analysis are presented along with detailed trajectory calculations. The use of micro/nano satellites in low-cost interplanetary exploration is attracting increasing attention and is the subject of several annual workshops and published design studies (1-4). The NEO population consists of those asteroids and short period comets orbiting the Sun with a perihelion of 1.3 astronomical units or less (5-8). As of July 30, 2013 10065 Near-Earth objects have been discovered. The spin rate, mass, density, surface physical (especially mechanical) properties, composition, and mineralogy of the vast majority of these objects are highly uncertain and the limited available telescopic remote sensing data imply a very diverse population (5-8). In-situ measurements by robotic spacecraft are urgently needed to provide the characterization data needed to support hardware and mission design for more ambitious human and robotic NEO operations. Large numbers of NEOs move into close proximity with the Earth-Moon system every year (9). The JPL Near-Earth Object Human Space Flight Accessible Targets Study (NHATS) (10) has produced detailed mission profile and delta V requirements for various NEO missions ranging from 30

  18. Editorial on Future Jet Technologies

    Gal-Or, Benjamin


    competions and Awards. [Rule-13]. We also provide 26 references [17-43] to a different, unclassified technology that enhances TV-induced flight safety for passenger jets, turboprops and helicopters. It is based on patented stowed-away/emergency-deployed TV-kits added to fixed-configuration, subsonic exhaust nozzles of low thrust-to-weight ratio vehicles. Expected benefits include anti-terror recoveries from emergencies, like forced landing on unprepared runways or highways, or recoveries from all airframe-hydraulics-outs, asymmetric ice on wings, landing gear catastrophes, and recoveries from pilot errors and bad-whether incidents [Rule 9(7)]. Other TV technologies involve preventing catastrophes in speed and patrol boats, racing and regular cars/SUVs, buses and trucks. [Rule 9(8)] and faster helicopters [Rule 9(10)].

  19. Impact landing ends SMART-1 mission to the Moon


    science at the same time,” Racca concluded. “Operating SMART-1 has been an extremely complex but rewarding task,” said Octavio Camino-Ramos, ESA SMART-1 Spacecraft Operations Manager. “The long spiralling trajectory around Earth to test solar electric propulsion (a low-thrust approach), the long exposure to radiation, the strong perturbations of the gravity fields of the Earth-Moon system and then the reaching of a lunar orbit optimised for the scientific investigations, have allowed us to gain valuable expertise in navigation techniques for low-thrust propulsion and innovative operations concepts: telemetry distribution and alerting through the internet, and a high degree of ground operations automation - a remarkable benchmark for the future,” he explained. “For ESA’s Science Programme, SMART-1 represents a great success and a very good return on investment, both from the technological and the scientific point of view,” said Professor Southwood, ESA’s Director of Science. “It seems that right now everyone in the world is planning on going to the Moon. Future scientific missions will greatly benefit from the technological and operational experience gained thanks to this small spacecraft, while the set of scientific data gathered by SMART-1 is already helping to update our current picture of the Moon.” Note to editors More images and further updates on the SMART-1 mission end can be found SMART-1, (Small Mission for Advanced Research and Technology) is the first European mission to the Moon. It was launched on 27 September 2003 on board an Ariane 5 rocket, from the CSG, Europe’s Spaceport in Kourou, French Guiana and reached its destination in November 2004 after following a long spiralling trajectory around Earth. In this phase, the spacecraft successfully tested for the first time in space the series of advanced technologies it carried on board. The technology demonstration part of the mission was declared successfully

  20. Short primary linear drive designed for synchronous and induction operation mode with on-board energy storage

    Fernandes Neto, Tobias Rafael


    guide way (induction rail or stationary magnets), and the energy and information should be transmitted contactless to the active vehicle. Regarding the features of the material handling application, the short or long primary topology can be used. Short primary linear drives on passive track are advantageous in material handling applications, where high precision, moderate dynamic, very long track and closed paths are required. Nevertheless, depending on the requirements of the section, the costs can be reduced considerably by using a simple induction rail at the long transporting sections, instead of permanent magnets on the track. Therefore, in this thesis a combined operation of permanent magnet linear synchronous motor (PMLSM) and linear induction motor (LIM) is applied to operate the short primary as vehicle, avoiding adjustment or releasing of the material during the drive cycle. In summary, the passive track will consist of two section types: a high thrust force section (processing station) with PMLSM and a low thrust force section with LIM (transporting section). To the author's knowledge, using two operation modes (PMLSM / LIM) in the same drive is a new approach. A theoretical and experimental study was conducted to assess the feasibility of employing the short primary linear motor for a flexible manufacturing system, in which a contactless energy transmission provides the basic power and an ultracapacitor (UC) storage system provides the peak power. The system uses a bidirectional DC-DC converter between the ultracapacitor bank and the DC-link, to make sure that the ultracapacitor can store the braking energy and supply the peak power demanded by the active vehicle. A control strategy has been developed for controlling the ultracapacitor to deliver the peak of power, to charge, to protect against overvoltage and to recover the energy generated when the vehicle is braking. A control strategy for the transition between the two operation modes (PMLSM / LIM

  1. CAD-Based Modeling of Advanced Rotary Wing Structures for Integrated 3-D Aeromechanics Analysis

    Staruk, William

    validation were carried out systematically, covering formulation, model accuracy, and accuracy of the physics of the problem and the many complex coupled aeromechanical phenomena that characterize the behavior of a tiltrotor in the conversion corridor. Compatibility of the new structural analysis models with X3D is demonstrated using analytical test cases, including 90° twisted beams and thick composite plates, and a notional bearingless rotor. Prediction of deformations and stresses in composite beams and plates is validated and verified against experimental measurements, theory, and state-of-the-art beam models. The second goal was met through integrated analysis of the Tilt Rotor Aeroacoustic Model (TRAM) proprotor using X3D coupled to Helios--the US Army's next generation CFD framework featuring a high fidelity Reynolds-average Navier-Stokes (RANS) structured/unstructured overset solver--as well as low order aerodynamic models. Although development of CFD was not part of this work, coupling X3D with Helios was, including establishing consistent interface definitions for blade deformations (for CFD mesh motion), aerodynamic interfaces (for loads transfer), and rotor control angles (for trim). It is expected that this method and solver will henceforth be an integral part of the Helios framework, providing an equal fidelity of representation for fluids and structures in the development of future advanced rotor systems. Structural dynamics analysis of the TRAM model show accurate prediction of the lower natural frequencies, demonstrating the ability to model advanced rotors from first principles using 3-D structural dynamics, and a study of how joint properties affect these frequencies reveals how X3D can be used as a detailed design tool. The CFD/CSD analysis reveals accurate prediction of rotor performance and airloads in edgewise flight when compared to wind tunnel test data. Structural blade loads trends are well predicted at low thrust, but a 3/rev component of flap

  2. Hubble Space Telescope: The Telescope, the Observations & the Servicing Mission


    NICMOS enabling it to resume operation, and install a new set of solar panels. Replacement of the thermal insulation will continue and the telescope will be reboosted to a higher orbit. The plans for the fourth Servicing Mission are preliminary at this time, but two new science instruments are being developed for that mission: Cosmic Origins Spectrograph (COS), which will replace COSTAR, and Wide Field Camera 3 (WFC3), which will replace WFPC2. It is planned to retrieve Hubble at the end of its life (around 2010) and bring it back to Earth. In the future ESA may have the opportunity to continue its collaboration with NASA on the Next Generation Space Telescope (NGST), which in many ways can be seen as Hubble's successor. The plan is to launch NGST in 2008, and ESA is currently considering a possible role in the project. Piero Benvenuti concludes: "The European Space Agency, in deciding to join NASA on the HST Project, made a very successful investment on behalf of European science. Today, NASA would not consider proceeding alone on the continued operation of HST or on the design of NGST. Not just because of the benefit of shared cost, but mainly because of the intellectual contribution by the European astronomers, who have made such effective scientific use of HST." Hubble Space Telescope - Fact sheet Description The Hubble Space Telescope (HST) is a co-operation between ESA and NASA. It is a long-term space-based observatory. Its observations are carried out in visible, infrared and ultraviolet light. HST has in many ways revolutionised modern astronomy, being a highly efficient tool for making new discoveries, but also by driving astronomical research in general. Objective HST was designed to take advantage of being above the Earth's disturbing atmosphere, and thereby providing astronomers with observations of very high resolution - opening new windows on planets, stars and galaxies. HST was designed as a flagship mission of the highest standard, and has served to pave

  3. Obituary: David Fulmer Bender, 1913-2004

    Miller, Sylvia L.


    of having him create a comprehensive set of ephemerides that would be available for the Advanced Projects Group to use for mission planning purposes. Dave worked with Phil Roberts, Carl Sauer, and others who were creating mission design software at the time to ensure that the asteroid file would be compatible with these computer programs. Dave, himself, authored many papers documenting trajectories he discovered to various asteroids, comets, and Lagrange points, along with the search techniques he used. He also documented surveys of opportunities, some for use with low-thrust propulsion as well as the more common chemical propulsion. Along with Raymond Jurgens, Dave published opportunities for radar astronomers to view asteroids passing close to the Earth. Dave did not restrict his investigations to small bodies. He also published papers on Venus missions, lunar swingby techniques, Jupiter gravity assist trajectories to Kuiper belt objects, and multibody-assist trajectories for missions to Jupiter's satellite Europa (the latter two in the 1990s!). Brian Marsden recalls that in1980 Dave visited him at the new facilities of the Minor Planet Center in Massachusetts and left with a box of new computer cards punched with the orbital elements of the 2000 asteroids known at that time. Colleagues at JPL remember how excited he was when he returned from that trip. As more asteroids were discovered, he would add their orbital elements to the file. Because of Dave's pioneering work in making the asteroid orbits accessible for mission studies before most people cared about these bodies, he can be credited in part for the mission Galileo's close flyby of both Gaspra (in 1991) and Ida (in 1993), along with the discovery of Dactyl, the first confirmed asteroid satellite. Dave eventually passed the responsibility of maintaining the small body file to Donald Yeomans and Ravenel (Ray) Wimberly at JPL. Now called DASTCOM, it includes elements for over 260,000 bodies, most of them

  4. People


    several quick passes, leaving an incomplete image of only half of the planet. This, and observations made from Earth, provide almost all of our knowledge of Mercury. Earth observations, however, are hampered by the planet's proximity to the Sun, making observations possible only at dawn and dusk. A mosaic of images of Mercury from the NASA Mariner 10 spacecraft. ©NASA In the mid-80s improved radar equipment allowed high resolution mapping of surface features from the Earth. Amongst the results were two tantalising mysteries: a large dome feature, similar in some ways to shield volcanoes seen on Mars, observed on the unimaged side of the planet and complex scattering of returned radar from distinct areas around the poles, suggesting that water ice may exist in craters there. Both NASA and the European Space Agency (ESA) are now planning missions to Mercury. The US team are using a newly discovered trajectory that will allow them to reach Mercury using traditional chemical propulsion, incorporating various planetary flybys so-called 'gravity assist' manoeuvres. The European team, on the other hand, has proposed a much more complex mission. In order to get to Mercury, ESA have adopted a novel technology knows as 'solar electric propulsion' (SEP). The basic principle is that electrical energy is produced using solar cells, and this is used to accelerate ions of gas, producing a continuous, if low thrust. The upshot is that the mission is much less constrained by the alignment of the planets and other trajectory concerns and can complete the journey in only two and a half years. BepiColombo, ESA's Mercury mission, will actually consist of three spacecraft! The planetary orbiter will stay close to Mercury and perform remote sensing and mapping of the surface environment. The magnetospheric orbiter, now going to be built by the Institute for Space and Astronautical Science (ISAS) in Japan, will fly in a highly eccentric orbit that takes it from within a few hundred kilometres