Showing papers in "Journal of Spacecraft and Rockets in 2001"
TL;DR: In this paper, the authors provide physical insight into satellite formation e ying design and outline the effects of realistic dynamics on those designs, including in-plane, in-track, circular, and projected circular designs.
Abstract: Several satellite formation e ying designs and their evolution through time are investigated. Satellite formation e ying designs arederived from the linearized equations of relative motion under two-body dynamics better known as Hill’ s equations (Hill, G. W., “ Researches in the Lunar Theory,”American Journal of Mathematics , Vol. 1, No. 1, 1878, pp. 5‐26). The formations are then propagated forward in time in the presence of realistic perturbations to determine the stability of each design. Formations considered include in-plane, in-track, circular, and projected circular designs. The Draper Semianalytic Satellite Theory is used to propagate mean elements of the satellites. When perturbations disrupt the satellite formations, an effort is made to quantify the cost of formation-keeping maneuvers. The goal of this effort is to provide physical insight into satellite formation e ying design and outline the effects of realistic dynamics on those designs.
375 citations
TL;DR: In this article, the authors reviewed the effect of facility noise on the trend in transition Reynolds numbers in conventional ground-test facilities, of both conventional and quiet design, at hypersonic and high supersonic speeds.
Abstract: It is well known that the high levels of noise present in conventional hypersonic ground-test facilities cause transition to occur earlier than in e ight. Flight measurements of incoming noise are reviewed and compared with measurements in ground-test facilities, of both conventional and quiet design, at hypersonic and high supersonic speeds. The low noise present in e ight is apparently the reason for the very large transition Reynolds numbers sometimes measured in e ight, when roughness, crosse ow, and other factors are controlled. Design will usually involve consideration of the trend in transition when a parameter is varied. The effect of facility noise on these trends is reviewed. In some cases, the trend of conventional-tunnel data is opposite to the trend in quiet-tunnel data. Thus, transition measurements in conventional ground-test facilities are not reliable predictors of e ight performance, except perhaps in special cases.
344 citations
TL;DR: In this article, boundary-layer trip devices for the Hyper-X forebody have been experimentally examined in several wind tunnels, including the NASALangleyResearch Center 20-Inch Mach 6 Air and 31-inch Mach 10 Air tunnels and in the HYPULSE Reeected Shock Tunnel at the General Applied Sciences Laboratory.
Abstract: Boundary-layer trip devices for the Hyper-X forebody have been experimentally examined in several wind tunnels.Fivedifferenttripconegurationswerecomparedinthreehypersonicfacilities:theNASALangleyResearch Center 20-Inch Mach 6 Air and 31-Inch Mach 10 Air tunnels and in the HYPULSE Reeected Shock Tunnel at the General Applied Sciences Laboratory. Heat-transfer distributions, utilizing the phosphor thermography and thin-elm techniques, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. Parametric variations include angles of attack of 0, 2, and 4 deg; Reynolds numbers based on model length of 1.2 ££ 10 6‐15.4 £ 10 6 ; and inlet cowl door simulated in both open and closed positions. Comparisons of boundary-layer transition as a result of discrete roughness elements have led to the selection of a trip coneguration for the Hyper-X Mach 7 eight vehicle.
186 citations
TL;DR: In this paper, the TITAN program for predicting charring material ablation and shape change of thermal protection materials is presented. But the authors focus on predicting the shape change due to surface recession.
Abstract: The TITANprogramforpredicting charringmaterial ablationand shapechangeof thermal protectionmaterials is presented. The governing equations include energy conservation and a three-component decomposition model. The surface energy balance condition is solved with a moving grid to calculate the shape change due to surface recession. The governing equations are discretized with a nite volume approximation with a general body- tted coordinate system. A time-accurate solution is achieved by an implicit time-marching techniquewith Gauss–Seidel line relaxationwith alternating sweeps. Benchmark solutionsare calculated and comparedwith availablesolutions to check code consistency and accuracy. For fully coupled solid– uid simulation, this technique has been directly integrated with both a high- delity Navier–Stokes solver and an aerothermal ow eld engineering correlation code. Representative computations, including a slender hypersonic reentry vehicle and a at-faced cylinder model in an arcjet test, are presented and discussed in detail.
126 citations
TL;DR: In this article, the present status of computer modeling of Hall thruster plumes is reviewed in the context of being able to address spacecraft integration concerns and a simple, empirical approach is described that can be used as a quick engineering tool.
Abstract: Hall thrusters are an attractive form of electric propulsion that are being developed to replace chemical systems for many orbit propulsion tasks on communications satellites. A concern in the use of these devices is the possible damage their plumes may cause to the host spacecraft. The present status of computer modeling of Hall thruster plumes is reviewed in the context of being able to address spacecraft integration concerns. A simple, empirical approach is described that can be used as a quick engineering tool. However, accurate modeling of Hall thruster plumes requires use of kinetic-based simulation techniques. In particular, particle methods are discussed with respect to the physical modeling required to accurately simulate the plasma and collision processes that are signi cant in Hall thruster plumes. An assessment is made of the computer models through direct comparison between simulation results and detailed experimental measurements.
123 citations
TL;DR: The first autonomous rendezvous docking (RVD) between uninhabited spacecraft was performed by ETS-VII on 7 July 1998 as mentioned in this paper, and the performance was better than required.
Abstract: On 7 July 1998, Engineering Test Satellite-VII (ETS-VII) successfully performed the e rst autonomous rendezvous docking (RVD) between uninhabited spacecraft ever. RVD technology is essential for future space activities, such aslogisticsupportfortheInternational SpaceStation.TheNationalSpaceDevelopment Agency ofJapan developed ETS-VII to demonstrate autonomous RVD technologies. ETS-VII carried two RVD experiment e ights successfully in 1998. We introduce the autonomous RVD system of ETS-VII and show the results and evaluations of two RVD experiment e ights. We realized autonomous uninhabited RVD, and its performance was better than required.
118 citations
TL;DR: In this paper, an unsteady scheme based on grid movement was proposed to solve the problem of estimating the Magnus force and moment of a yawing and spinning axisymmetric projectile.
Abstract: Computations of steady e ows over yawing and spinning axisymmetric projectiles are largely carried out by numerical algorithms using steady methods. Of particular interest is the prediction of the Magnus force and moment. However this axisymmetric characteristic is lost with e ns addition, and the e ow becomes unsteady whatever the framework is. ONERA and GIAT Industries have developed a new unsteady scheme, based on grid movement, that allows such a turbulent unsteady e ow to be solved. This scheme has been used successfully over a spinning and yawed body-tail cone guration. The Magnus effect is generated on the body by the spin-induced boundary-layer distortion at moderate incidences, whereas asymmetric vortices tend to invert this effect at upper incidences. Fins contribute to an opposite and greater lateral force. The total Magnus force appeared to be linear with respect to angle of attack and spin rate, but the range of linearity of angle of attack is much smaller than for a none nned body.
113 citations
TL;DR: In this article, the effects of discrete and distributed roughness elements on boundary layer transition, which included trip height, size, location, and distribution, both on and off the windward centerline, were investigated.
Abstract: Boundary layer and aeroheating characteristics of several X-33 configurations have been experimentally examined in the Langley 20-Inch Mach 6 Air Tunnel. Global surface heat transfer distributions, surface streamline patterns, and shock shapes were measured on 0.013-scale models at Mach 6 in air. Parametric variations include angles-of-attack of 20-deg, 30-deg, and 40-deg; Reynolds numbers based on model length of 0.9 to 6.6 million; and body-flap deflections of 0, 10 and 20-deg. The effects of discrete and distributed roughness elements on boundary layer transition, which included trip height, size, location, and distribution, both on and off the windward centerline, were investigated. The discrete roughness results on centerline were used to provide a transition correlation for the X-33 flight vehicle that was applicable across the range of reentry angles of attack. The attachment line discrete roughness results were shown to be consistent with the centerline results, as no increased sensitivity to roughness along the attachment line was identified. The effect of bowed panels was qualitatively shown to be less effective than the discrete trips; however, the distributed nature of the bowed panels affected a larger percent of the aft-body windward surface than a single discrete trip.
110 citations
TL;DR: In this paper, a particle-in-cell simulation model was developed to compute the ion propulsion induced plasma environment for the Deep Space 1 (DS1) spacecraft, and the results showed good agreement.
Abstract: Afully three-dimensionalparticle-in-cell simulationmodelwas developed to compute the ionpropulsion induced plasma environment for the Deep Space 1 (DS1) spacecraft. Simulationsare comparedwith in- ightmeasurements of charge-exchange plasma from the ion propulsion diagnostics subsystem on DS1, and the results show good agreement. It is found that the plasma environment of DS1 is dominated by the charge-exchange plasma from the plume. For a typical ion thruster operating condition, the charge exchange plasma near the spacecraft surface has a density ranging from 106 cmi 3 near the thruster to 104 cmi 3 at the opposite end of the spacecraft and a current density ranging from 10i 7 A/cm2 to 10 i 9 A/cm2. It is shown that, for an interplanetary spacecraft with a moderate charging potential, charge-exchange ion back ow is through an expansion process similar to that of the expansion of a mesothermal plasma into a vacuum.
92 citations
TL;DR: In this paper, it is demonstrated that a sun-synchronous apse-line precession can be achieved using only a small, low-cost solar sail, which can be configured to permanently station a science payload within the geomagnetic tail.
Abstract: Conventional geomagnetic tail missions require a spacecraft to be injected into a long elliptical orbit to explore the spatial structure of the geomagnetic tail. However, because the elliptical orbit is inertially fixed and the geomagnetic tail is directed along the sun-Earth line, the apse line of the elliptical orbit is precisely aligned with the geomagnetic tail only once every year. To artificially precess the apse line of the elliptical orbit in a sun-synchronous manner, which would keep the spacecraft in the geomagnetic tail during the entire year, would require continuous low-thrust propulsion or periodic impulses from a high-thrust propulsion system. Both of these options require reaction mass that will ultimately limit the mission lifetime. It is demonstrated that sun-synchronous apse-line precession can be achieved using only a small, low-cost solar sail. Because solar sails do not require reaction mass, a geomagnetic tail mission can be configured that provides a continuous science return by permanently stationing a science payload within the geomagnetic tail.
91 citations
TL;DR: In this article, the concept of symmetrical double-ended motorized spinning tethers for use as orbital transfer vehicles is introduced. But the performance of these tethers is limited by their high motor torque and a safety factor close to unity.
Abstract: The concept of symmetrical double-ended motorized spinning tethers for use as orbital transfer vehicles is introduced. The orbital elements of a payload released from above and below a hanging, prograde librating, and prograde spinning tether are derived and employed to evaluate the effectiveness of the three tether types along with their optimum configurations for payload transfer. A new ratio, the efficiency index, is defined as the altitude gain or loss half an orbit after tether release per tether length. The motorized tether is found to perform best and also most efficiently, improving by two orders of magnitude on the librating tether, which, in turn, improves on the hanging tether by a factor of two. A long motorized tether on a circular orbit can transfer an upper payload from a low to a geostationary Earth orbit by employing relatively high motor torque and a safety factor on the tether strength close to unity.
TL;DR: This work finds both hovering points and orbiting trajectories about various sized asteroids using equations of motion for a solar sail spacecraft that are stable and offer good coverage of the asteroid surface, although re- strictions on sail acceleration are needed for smaller asteroids.
Abstract: The inherent capabilities of solar sails and the fact that they need no onboard supplies of fuel for propul- sion make them well suited for use in long-term, multiple-objective missions. They are especially well suited for the exploration of asteroids, where one spacecraft could rendezvous with a number of aster- oids in succession. The orbital mechanics of solar sail operations about an asteroid, however, have not yet been studied in detail. Building on previous stud- ies that consider the equations of motion, we find both hovering points and orbiting trajectories about various sized asteroids using equations of motion for a solar sail spacecraft. These hovering points are stabilizable using feedback control to sail attitude alone. The orbiting trajectories are stable and offer good coverage of the asteroid surface, although re- strictions on sail acceleration are needed for smaller asteroids.
TL;DR: In this paper, the Spalart-Allmaras one-equation model was originally developed in substantial derivative form and when rewritten in conservation form, a density gradient term appears in the source term.
Abstract: Many Navier-Stokes codes require that the governing equations be written in conservation form with a source term. The Spalart-Allmaras one-equation model was originally developed in substantial derivative form and when rewritten in conservation form, a density gradient term appears in the source term. This density gradient term causes numerical problems and has a small influence on the numerical predictions. Further work has been performed to understand and to justify the neglect of this term. The transition trip term has been included in the one-equation eddy viscosity model of Spalart-Allmaras. Several problems with this model have been discovered when applied to high-speed flows. For the Mach 8 flat plate boundary layer flow with the standard transition method, the Baldwin-Barth and both k-{omega} models gave transition at the specified location. The Spalart-Allmaras and low Reynolds number k-{var_epsilon} models required an increase in the freestream turbulence levels in order to give transition at the desired location. All models predicted the correct skin friction levels in both the laminar and turbulent flow regions. For Mach 8 flat plate case, the transition location could not be controlled with the trip terms as given in the Spalart-Allmaras model. Several other approaches have been investigated to allow the specification of the transition location. The approach that appears most appropriate is to vary the coefficient that multiplies the turbulent production term in the governing partial differential equation for the eddy viscosity (Method 2). When this coefficient is zero, the flow remains laminar. The coefficient is increased to its normal value over a specified distance to crudely model the transition region and obtain fully turbulent flow. While this approach provides a reasonable interim solution, a separate effort should be initiated to address the proper transition procedure associated with the turbulent production term. Also, the transition process might be better modeled with the Spalart-Allmaras turbulence model with modification of the damping function f{sub v1}. The damping function could be set to zero in the laminar flow region and then turned on through the transition flow region.
TL;DR: In this article, a ring of appropriately sized lateral pulse jets coupled with a trajectory tracking engine was used to reduce the impact point dispersion of a direct-e re rocket, which was shown to work well against uncertainty in the form of initial off-axis angular velocity perturbations as well as atmospheric winds.
Abstract: Impact point dispersion of a direct e re rocket can be drastically reduced with a ring of appropriately sized lateral pulse jets coupled to a trajectory tracking e ight control system. The system is shown to work well against uncertainty in the form of initial off-axis angular velocity perturbations as well as atmospheric winds. For an example case examined, dispersion was reduced by a factor of 100. Dispersion reduction is a strong function of the number of individual pulse jets, the pulse jet impulse, and the trajectory tracking window size. Proper selection of these parameters for a particular rocket and launcher combination is required to achieve optimum dispersion reduction. Forrelatively lowpulsejetimpulse, dispersion steadily decreasesasthenumberofpulsejetsisincreased or as the pulse jet impulse is increased. For a e xed total pulse jet ring impulse, a single pulse is the optimum pulse jet cone guration when the pulse jet ring impulse is small because the effect of a pulse on the trajectory of a rocket decreases as the round e ies downrange. Nomenclature CDD = e n cant roll moment aerodynamic coefe cient CLP = roll damping aerodynamic coefe cient CMQ = pitch damping aerodynamic coefe cient CNA = normal force aerodynamic coefe cient CX0 = zero yaw axial force aerodynamic coefe cient CX2 = yaw axial force aerodynamic coefe cient D = rocket reference diameter ethres = trajectory tracking window size L; M; N = total applied moments about rocket mass center expressed in the aft body reference frame nJ = number of individual lateral pulse jets nRXi ;nRYi , = ith main rocket motor direction cosines in the nRZi body frame p;q;r = components of the angular velocity vector of the projectile in the body reference frame T = P ° time constant TJi = ith lateral pulse jet thrust TRi = ith main rocket motor thrust t ¤ = time of the most recent pulse jet e ring u;v;w = components of the velocity vector of the mass center of the composite body in the body reference frame uA;vA;wA = components of the velocity of the mass center of the projectile with mean wind expressed in the body reference frame VA = magnitude of the velocity vector of the mass center of the projectile experienced with mean wind expressed in the body reference frame VMW;aeMW = magnitude and wind factor of the mean atmospheric wind expressed in the initial reference frame X;Y; Z = total applied force components in the aft body reference frame
TL;DR: In this article, the experimental investigation of the original method of providing radio communications with a reentry vehicle along the trajectory section when a vehicle is surrounded by plasma of the ionized shock layer was given.
Abstract: Consideration is given to the experimental investigation of the original method of providing radio communications with a reentry vehicle along the trajectory section when a vehicle is surrounded by plasma of the ionized shock layer. The method suggests placing of the antennas in special small containers. The containers are located at about zero angle of attack on a pylon ahead of the bow shock wave generated near a vehicle. Because of the small container bluntness low level of ionization near the antennas could be provided. But many problems connected with heating and thermal protection of the container and pylon arise in this case. Therefore along with the measurements of ionization near the container, heating and thermal protection of the container and pylon and their ine uenceon thevehicleaerodynamicswereinvestigated.Theexperiments wereperformed in fourhypersonicwind tunnels of different types in Mach-number range from 6.5 to 20.5. Various measurement methods were used. The investigation shows that the remote antenna assembly can provide uninterrupted radio communication with the reentry vehicle.
TL;DR: An overview of the activities associated with the aerodynamic database that is being developed in support of NASA's Hyper-X flight experiments is provided in this paper, including wind-tunnel test activities and parallel computational fluid dynamics analysis efforts.
Abstract: An overview of the activities associated with the aerodynamic database that is being developed in support of NASA's Hyper-X scramjet flight experiments is provided. Three flight tests are planned as part of the Hyper-X program. Each will utilize a small, non-recoverable research vehicle with an airframe-integrated scramjet propulsion engine. The research vehicles will be individually rocket boosted to the scramjet engine test points at Mach 7 and 10. The research vehicles will then separate from the first stage booster vehicle and the scramjet engine test will be conducted before the terminal decent phase of the flight. An overview is provided of the activities associated with the development of the Hyper-X aerodynamic database, including wind-tunnel test activities and parallel computational fluid dynamics analysis efforts for all phases of the Hyper-X flight tests. A brief summary of the Hyper-X research vehicle aerodynamic characteristics is provided, including the direct and indirect effects of the airframe-integrated scramjet propulsion system operation on the basic airframe stability and control characteristics. Brief comments on the planned postflight data analysis efforts are also included.
TL;DR: In this paper, the authors report recently completed structural dynamics experimental activities with new ultra-lightweight and inflatable space structures (a.k.a., "Gossamer" spacecraft) at NASA Langley Research Center, NASA Marshall Space Flight Center, and NASA Goddard Space Flight center.
Abstract: This paper reports recently completed structural dynamics experimental activities with new ultra-lightweight and inflatable space structures (a.k.a., "Gossamer" spacecraft) at NASA Langley Research Center, NASA Marshall Space Flight Center, and NASA Goddard Space Flight Center. Nine aspects of this work are covered: 1) inflated, rigidized tubes, 2) active control experiments, 3) photogrammetry, 4) laser vibrometry, 5) modal tests of inflatable structures, 6) in-vacuum modal tests, 7) tensioned membranes, 8) deployment tests, and 9) flight experiment support. Structural dynamics will play a major role in the design and eventually in-space deployment and performance of Gossamer spacecraft. Experimental research and development such as this is required to validate new analysis methods. The activities discussed in the paper are pathfinder accomplishments, conducted on unique components and prototypes of future spacecraft systems.
TL;DR: In this article, the results of velocity and temperature boundary-layer profile measurements made at the Dresden University of Technology were used to characterize the influence of foreign gas transpiration on skin friction and heat transfer.
Abstract: Blowing through the wall is one of the most efficient methods to influence the characteristics of a turbulent boundary layer such as heat transfer, skin friction, boundary-layer profiles, and flow separation. The most important technical application is to cool a permeable wall with a coolant mass flow through the wall, known as transpiration cooling. Designing transpiration-cooled devices requires the knowledge of empirical correlations for the calculation of skin friction and heat transfer in the case of blowing. The results of velocity and temperature boundary-layer profile measurements made at the Dresden University of Technology were used to characterize the influence of foreign gas transpiration on skin friction and heat transfer. Simple empirical correlations describing skin friction and heat transfer reduction as functions of the blowing ratio and especially the blowing gas properties were found. Further investigations deal with flow separation due to blowing, and a universal critical blowing parameter is defined. The empirical correlations have also been verified by means of a numerical calculating procedure. Measurements made at the German Aerospace Center in Lampoldshausen using an H2/O2 combustion chamber delivered experimental heat transfer results in the case of hot gas parameters according to practical applications.
TL;DR: Airframe-integrated scramjet engine testing has been completed at Mach 7 flight conditions in the NASA Langley 8-Foot High Temperature Tunnel and the subsystems that were subjected to flight-like conditions are described and supporting data is presented.
Abstract: Airframe-integrated scramjet engine testing has been completed at Mach 7 flight conditions in the NASA Langley 8-Foot High Temperature Tunnel as part of the NASA Hyper-X program. This test provided engine performance and operability data, as well as design and database verification, for the Mach 7 flight tests of the Hyper-X research vehicle (X-43), which will provide the first-ever airframe-integrated scramjet data in flight. The Hyper-X Flight Engine, a duplicate Mach 7 X-43 scramjet engine, was mounted on an airframe structure that duplicated the entire three-dimensional propulsion flowpath from the vehicle leading edge to the vehicle trailing edge. This model was also tested to verify and validate the complete flight-like engine system. This paper describes the subsystems that were subjected to flight-like conditions and presents supporting data. The results from this test help to reduce risk for the Mach 7 flights of the X-43.
TL;DR: In this paper, the plume model of a pulsed plasma thruster (PPT) onboard a spacecraft is presented, which is based on a hybrid (particle-e uid) methodology.
Abstract: Integration of pulsed plasma thruster (PPT) onboard spacecraft requires the evaluation of potential plume/spacecraft interactions. Important e ndings are summarized of our experimental and modeling plume investigations of rectangular-geometry Tee on ® PPTs. Initial studies of the Lincoln Experimental Satellite 8 /9 PPT plume used time-of-e ight analysis of single langmuir probe data and found two ion populations with approximately 30 and 60 km /s, respectively. A residual gas analyzer identie ed C, F, C xFy, and various thruster materials. Fast ionization gauges detected the presence of slow neutral particles up to 1 ms after the end of the discharge. Subsequent studies used triple langmuir probes and obtained electron temperature and density in the plume of a laboratory PPT operating at 5, 20, and 40 J. Plume properties showed large angular density variation on the perpendicular to the electrodes plane but small variation on the parallel plane. Electron density and temperature were found to decrease with increasing radial distance from the Tee on surface. Time-average temperatures were between 1 and 3 eV. Time-averageelectron density increased with increasing dischargeenergy and are in therange between 10 19 and 2 ££ 10 20 m i 3 for 5 J, 6 ££ 10 20 to 10 21 m i 3 for 20 J, and 2 ££ 10 20 to 1.4 ££ 10 21 m i 3 for 40 J. The PPT plume model is based on a hybrid (particle‐e uid) methodology. Neutrals and ions were modeled with a combination of the direct simulation Monte Carlo and a hybrid-particle-in-cell method. Electrons were modeled as a massless e uid with a momentum equation that includes collisional contributions from ions and neutrals and an energy equation. Simulations of the laboratory PPT operating at discharge energies of 5, 20, and 40 J showed the expansion of the neutral and ion components of the plume during a pulse, the generation of low-energy ions and high-energy neutrals dueto chargeexchange reactions, and the generation ofbacke ow. Numerical predictions showed good quantitative agreements with data.
TL;DR: In this paper, a qualitative assessment of lateral-directional stability characteristics was made through a series of tip-to-tail inviscid calculations, including a simulation of the powered scramjet eight-test condition.
Abstract: Computational e uid dynamics tools have been used extensively in the analysis and development of the X-43A Hyper-X Research Vehicle. A signiecant element of this analysis is the prediction of integrated vehicle aeropropulsive performance,which includes an integration of aerodynamic and propulsion eoweelds. Thedevelopment of the Mach7X-43Arequiredapree ightassessmentoflongitudinalandlateral-directionalaeropropulsivecharacteristics nearthetargeteight-testcondition.Thedevelopmentofthispree ightdatabasewasaccomplishedthroughextensive aerodynamic wind-tunnel testing and a combination of three-dimensional inviscid airframe calculations and cowlto-tail scramjet cycle analyses togenerate longitudinal performance increments between mission sequences. These increments were measured directly and validated through tests of the Hyper-X e ight engine and vehicle e owpath simulator in the NASA Langley Research Center 8-Foot High Temperature Tunnel. Predictions were reened with tip-to-tailNavier‐Stokescalculations,which alsoprovidedinformation onscramjetexhaustplumeexpansion inthe aftbody region. A qualitative assessment of lateral-directional stability characteristics was made through a series of tip-to-tail inviscid calculations, including a simulation of the powered scramjet eight-test condition. Additional comparisons with wind-tunnel forceand momentdataas well as surfacepressuremeasurements from theHyper-X e ight engine and vehicle eowpath simulator model and wind-tunnel testing were made to assess solution accuracy.
TL;DR: In this paper, a comprehensive investigation of Hall thruster plume plasma ion energy distribution functions and ion charge state has been made on both e ight-and laboratory-model engines.
Abstract: A comprehensive investigation of Hall thruster plume plasma ion energy distribution functions and ion charge state has been made on both e ight- and laboratory-model engines. An energy analyzer, mass spectrometer, and E £ B probe were used to characterize Hall thruster plume ions. The results of this investigation show that the distribution function of the ion beam exhibits both Maxwellian and Druyvesteyn traits and that the Hall thruster plumecontainsa nontrivial amountofenergetic,multiplycharged particlesthat mustbeaccounted forinmodeling the erosion rate of solar array cover glass and interconnect material. Detection of these multiply charged ions by energy analyzers has been hampered in the past by charge exchange and elastic collisions. The high-energy tail seen in numerous energy analyzer data is thought to result from charge exchange and elastic collisions between singly and multiply charged ions and neutrals. The role of facility pressure was also investigated and was found to have an ine uence mainly on the width of the ion energy distribution function. This pressure broadening is caused by elastic collisions between beam ions and background chamber gas particles. Nomenclature B = magnetic e eld vector, T E = electric e eld, V /m E = electric e eld vector, V /m Eb = ion beam energy, eV Ei = ion energy, eV f = ion distribution function f .Ei/ = ion energy distribution function f .v/ = ion velocity distribution function Ii = E£B probe collector current, A ni = ion number density, m i3 q = elementary charge, C qI = charge state of ion r = position vector, m t = time, s ui = ion speed, m /s Vi = beam ion acceleration potential, V v = ion velocity vector, m /s
TL;DR: The onset of spacecraft charging in a geosynchronous-satellite environment is independent of the ambient electron density, ion density, and ion temperature but depends solely on a critical, or threshold, electron temperature.
Abstract: The onset of spacecraft charging in a Maxwellian space environment is independent of the ambient electron density, ion density, and ion temperature but depends solely on a critical, or threshold, electron temperature. Below it, no spacecraft charging occurs; above it, spacecraft charging occurs. The spacecraft-charging potential is determined by the balance of all currents, including the incoming electrons, outgoing electrons, and ambient ions. Abundant evidence from the Los Alamos National Laboratory (LANL) geosynchronous-satellite data supports the existence of critical temperature for the onset of spacecraft charging. Comparison of the theoretical curve with observations on the LANL-1994-084 satellite is encouraging. Whereas the electron-induced secondary-electron coefficient controls the onset of spacecraft charging at about from - 1 to -2 kV, the backscattered-electron emission coefficient and the ion-induced electron emission coefficient, which is commonly neglected, control the charging level at higher potentials beyond about from -3 to -4 kV.
TL;DR: In this paper, an overview of the experimental aerodynamics test program to ensure mission success for the autonomous flight of the Hyper-X research vehicle, a 12ft-long, 2700-lb lifting body technology demonstrator designed to flight demonstrate for the first time a fully airframe-integrated scramjet propulsion system.
Abstract: An overview is provided of the experimental aerodynamics test program to ensure mission success for the autonomous flight of the Hyper-X research vehicle, a 12-ft-long, 2700-lb lifting body technology demonstrator designed to flight demonstrate for the first time a fully airframe-integrated scramjet propulsion system. Three flights are planned, two at Mach 7 and one at Mach 10. The research vehicles will be boosted to the prescribed scramjet engine test point, where they will separate from the booster, stabilize, and initiate engine test. Following more than 5 s of powered flight and 15 s of cowl-open tares, the cowl will close, and the vehicle will fly a controlled deceleration trajectory, which includes numerous control doublets for in-flight aerodynamic parameter identification. The preflight testing activities, wind-tunnel models, test rationale, risk reduction activities, and sample results from wind-tunnel tests supporting the flight trajectory from hypersonic engine test point through subsonic flight termination are reviewed.
TL;DR: Analytical and numerical investigations are performed that model the motion of a bouncing target marker across the surface of a rotating asteroid and some target limits on the target marker coef cient of restitution are developed.
Abstract: The proposed asteroid sample return mission MUSES-C calls for a spacecraft to approach an asteroid, touch down on its surface, and collect samples that will be returned to Earth. During the touchdownand samplingphase, the spacecraft will navigate relative to the asteroid surface using optical target markers placed on the asteroid surface before the nal approach. By using the target marker as a reference point, navigation during the landing phase will be much more reliable and precise. Because of the microgravity environment on the asteroid surface, the settling time and dynamics of the target markers are items of interest. Thus, it is important to design the target marker with as small a coef cient of restitution as possible to minimize the settling time, which in turn minimizes the time the spacecraft must hover above the asteroid surface. To achieve this small coef cient of restitution, the target marker will be constructed out of a bag with balls stored internally. On impact, the balls will dissipate energy relative to each other and, hence, will dissipate the total energy of the target marker. To better predict the performance of such a target marker, analytical and numerical investigations are performed that model the motionof a bouncing targetmarker across the surface of a rotating asteroid. As a result of the analysis, some target limits on the target marker coef cient of restitution are developed. A series of microgravity tests are reported that con rm the basic design and show that the target value of coef cient of restitution can be reached.
TL;DR: The early planning activity, background, and chronology that developed the series of wind-tunnel tests to support multi-degree-of-freedom simulation of the separation process are discussed in this article.
Abstract: NASA's Hyper-X research program was developed primarily to flight demonstrate a supersonic combustion ramjet engine, fully integrated with a forebody designed to tailor inlet flow conditions and a free expansion nozzle/afterbody to produce positive thrust at design flight conditions. With a point-designed propulsion system the vehicle must depend on some other means for boost to its design flight condition. Clean separation from this initial propulsion system stage within less than a second is critical to the success of the flight. This paper discusses the early planning activity, background, and chronology that developed the series of wind-tunnel tests to support multi-degree-of-freedom simulation of the separation process. Representative results from each series of tests are presented, and issues and concerns during the process and current status are highlighted.
TL;DR: An overview of the hypersonic aeroheating wind tunnel program conducted at the NASA Langley Research Center in support of the ground-based testing activities is provided in this article, where global surface heat transfer images, surface streamline patterns, and shock shapes are measured on 0.013 scale (10in.) ceramic models of the proposed X-33 configuration in Mach 6 air.
Abstract: The goal of the NASA Reusable Launch Vehicle (RLV) technology program is to mature and demonstrate essential, cost effective technologies for next generation launch systems. The X-33 flight vehicle presently being developed by Lockheed Martin is an experimental Single Stage to Orbit (SSTO) demonstrator that seeks to validate critical technologies and insure applicability to a full scale RLV. As with the design of any hypersonic vehicle, the aeroheating environment is an important issue and one of the key technologies being demonstrated on X-33 is an advanced metallic Thermal Protection System (TPS). As part of the development of this TPS system, the X-33 aeroheating environment is being defined through conceptual analysis, ground based testing, and computational fluid dynamics. This report provides an overview of the hypersonic aeroheating wind tunnel program conducted at the NASA Langley Research Center in support of the ground based testing activities. Global surface heat transfer images, surface streamline patterns, and shock shapes were measured on 0.013 scale (10-in.) ceramic models of the proposed X-33 configuration in Mach 6 air. The test parametrics include angles of attack from -5 to 40 degs, unit Reynolds numbers from 1x106 to 8x106/ft, and body flap deflections of 0, 10, and 20 deg. Experimental and computational results indicate the presence of shock/shock interactions that produced localized heating on the deflected flaps and boundary layer transition on the canted fins. Comparisons of the experimental data to laminar and turbulent predictions were performed. Laminar windward heating data from the wind tunnel was extrapolated to flight surface temperatures and generally compared to within 50 deg F of flight prediction along the centerline. When coupled with the phosphor technique, this rapid extrapolation method would serve as an invaluable TPS design tool.