Showing papers in "Journal of Spacecraft and Rockets in 2010"
TL;DR: An iterative algorithm for determining density and crosswind from multiaxis accelerometer measurements on satellites is presented, which works independently of the orientation of the instrument in space, and is compared with previously published algorithms using simulated data for the challenging minisatellite payload.
Abstract: An iterative algorithm for determining density and crosswind from multiaxis accelerometer measurements on satellites is presented, which works independently of the orientation of the instrument in space. The performance of the algorithm is compared with previously published algorithms using simulated data for the challenging minisatellite payload. Without external error sources, the algorithm reduces rms density errors from 0.7 to 0.03% andrmswinderrorsfrom38to1 m=sinthistest.However,theeffectsoftheerrorsintheinstrumentcalibrationand the external models that are used in the density and wind retrieval are dominant for the challenging minisatellite payload. These lead to mostly systematic density errors of the order of 10–15%. The accuracy of the wind results whenusingthenewalgorithmisalmostfullydeterminedbythesensitivityofthecross-trackaccelerationcomponent to the calibration and radiation pressure modeling errors. The applicability of the iterative algorithm and the accuracyofits resultsaredemonstrated bypresenting challenging minisatellite payload datafromaperiodin which thesatellitewascommandedto flysidewaysandbycomparingthedensityandwindresultswiththosefromadjacent days for which the satellite was in its nominal attitude mode. These investigations result in recommendations for the design of future satellite accelerometer missions for thermosphere research.
161 citations
TL;DR: In this paper, an ablation and thermal response model was developed for newly manufactured material, including emissivity, heat capacity, thermal conductivity, elemental composition, and thermal decomposition rates.
Abstract: Phenolic Impregnated Carbon Ablator was the heatshield material for the Stardust probe and is also a candidate heatshield material for the Orion Crew Module. As part of the heatshield qualification for Orion, physical and thermal properties were measured for newly manufactured material, included emissivity, heat capacity, thermal conductivity, elemental composition, and thermal decomposition rates. Based on these properties, an ablation and thermal-response model was developed for temperatures up to 3500 K and pressures up to 100 kPa. The model includes transversely isotropic and pressure-dependent thermal conductivity. In this work, model validation is accomplished by comparison of predictions with data from many arcjet tests conducted over a range of stagnation heat flux and pressure from 107 W/cm 2 at 2.3 kPa to 1100 W/cm 2 at 84 kPa. Over the entire range of test
153 citations
TL;DR: In this paper, a multiscale approach is used to model and analyze the ablation of porous materials, including carbon preform and char layer of two Phenolic Impregnated Carbon Ablators (PICA) with the same chemical composition but with different structures.
Abstract: A multiscale approach is used to model and analyze the ablation of porous materials. Models are developed for the oxidation of a carbon preform and of the char layer of two Phenolic Impregnated Carbon Ablators (PICA) with the same chemical composition, but with different structures. Oxygen diffusion through the pores of the materials and in depth oxidation and mass loss are first modeled at microscopic scale. The microscopic model is then averaged in set of partial differential equations describing the macroscopic behavior of the material. Microscopic and macroscopic approaches are applied with progressive degrees of complexity to gain a comprehensive understanding of the ablation process. Porous medium ablation is found to occur in a zone of the char layer, called ablation zone, whose thickness is a decreasing function of the Thiele number. The studied PICA materials are shown to display different ablation behaviors, a fact that is not captured by current models that are based on chemical composition only. Applied to Stardust’s PICA, the models explain and reproduce the unexpected drop in density measured in the char layer during Stardust post-flight analyses [Stackpoole, 2008].
152 citations
TL;DR: In this article, the authors combine the theory that gas-surface interactions in low Earth orbit are driven by adaption of atomic oxygen, with observations of satellite accommodation collected during solar cycle 22.
Abstract: The energy-accommodation coefficient is an important parameter affecting satellite drag and orbit predictions. Previous estimates of this coefficient have been based on interpolation from values tabulated at several altitudes and solar conditions. In an effort to improve drag coefficient accuracy and to compute values of the accommodation coefficient that respond to the real variability of the atmosphere, a first-principles approach is desired. The present work combines the theory that gas–surface interactions in lowEarth orbit are driven byadsorption of atomic oxygen, with observations of satellite accommodation collected during solar cycle 22. The result is a semiempirical model based on Langmuir’s adsorption isotherm, which agrees with the data to within 3%. This model can be used to improve drag predictions during a wide range of space weather conditions, as well as to improve the accuracy for atmospheric densities derived from satellite drag.
89 citations
TL;DR: This assessment has shown that negligible propellant mass fraction benefits are seen for reducing the three-sigma position dispersion at the end of the hypersonic guidance phase (parachute deployment) below approximately 3 km.
Abstract: Landing site selection is a compromise between safety concerns associated with the site’s terrain and scientific interest. Therefore, technologies enabling pinpoint landing performance (sub-100-m accuracies) on the surface of Mars are of interest to increase the number of accessible sites for in situ research, as well as allow placement of vehicles nearby prepositioned assets. A survey of the performance of guidance, navigation, and control technologies that could allow pinpoint landing to occur at Mars was performed. This assessment has shown that negligible propellant mass fraction benefits are seen for reducing the three-sigma position dispersion at the end of the hypersonic guidance phase (parachute deployment) below approximately 3 km. Four different propulsive terminal descent guidancealgorithms were examined. Of these four, a near propellant-optimal analytic guidance law showed promisefortheconceptualdesignofpinpointlandingvehicles.Theexistenceofapropellantoptimumwithregardto theinitiationtimeofthepropulsiveterminaldescentwasshowntoexistforvarious flightconditions.Subsonicguided parachutes were shown to provide marginal performance benefits, due to the timeline associated with descent through the thin Mars atmosphere. This investigation also demonstrates that navigation is a limiting technology for Mars pinpoint landing, with landed performance being largely driven by navigation sensor and map tie accuracy.
87 citations
86 citations
TL;DR: In this article, the authors revisited propulsion and power generation by bare electrodynamic tethers in a unified way and issues and challenges associated with tether temperature, bowing, deployment, and arcing are addressed.
Abstract: Propulsion and power generation by bare electrodynamic tethers are revisited in a unified way and issues and
constraints are addressed. In comparing electrodynamic tethers, which do not use propellant, with other propellantconsuming systems, mission duration is a discriminator that defines crossover points for systems with equal initial masses. Bare tethers operating in low Earth orbit can be more competitive than optimum ion thrusters in missions exceeding two-three days for orbital deboost and three weeks for boosting operations. If the tether produces useful onboard power during deboost, the crossover point reaches to about 10 days. Power generation by means of a bare electrodynamic tether in combination with chemical propulsion to maintain orbital altitude of the system is more efficient than use of the same chemicals (liquid hydrogen and liquid oxygen) in a fuel cell to produce power for missions longer than one week. Issues associated with tether temperature, bowing, deployment, and arcing are also discussed.
Heating/cooling rates reach about 4 K/s for a 0.05-mm-thick tape and a fraction of Kelvin/second for the ProSEDS
(0.6-mm-radius) wire; under dominant ohmic effects, temperatures areover200K (night) and 380 K (day) for the tape and 320 and 415 K for that wire. Tether applications other than propulsion and power are briefly discussed.
82 citations
TL;DR: A novel algorithm is derived from the redefinition of some evolutionary heuristics in the form of a discrete dynamical system that displays very good performance on the particular class of problems presented in this paper.
Abstract: In this paper, we analyze the performance of some global search algorithms on a number of space trajectory design problems. A rigorous testing procedure is introduced to measure the ability of an algorithm to identify the set of ²-optimal solutions. From the analysis of the test results, a novel algorithm is derived. The development of the novel algorithm starts from the redefinition of some evolutionary heuristics in the form of a discrete dynamical system. The convergence properties of this discrete dynamical system are used to derive a hybrid evolutionary algorithm that displays very good performance on the particular class of problems presented in this paper.
82 citations
TL;DR: In this article, the predicted shock-layer emission signatures during the reentry of the Japanese Hayabusa capsule are compared with flight measurements conducted during an airborne observation mission in NASA's DC-8 Airborne Laboratory.
Abstract: Predicted shock-layer emission signatures during the reentry of the Japanese Hayabusa capsule are presented and compared with flight measurements conducted during an airborne observation mission in NASA's DC-8 Airborne Laboratory. For selected altitudes at 11 points along the flight trajectory of the capsule, lines of sight were extracted from flow field solutions computed using the in-house high-fidelity CFD code, DPLR. These lines of sight were used as inputs for the line-by-line radiation code NEQAIR, and emission spectra of the air plasma were computed in the wavelength range from 300 nm to 1600 nm, a range which covers all of the different experiments onboard the DC-8. In addition, the computed flow field solutions were post-processed with the material thermal response code FIAT, and the resulting surface temperatures of the heat shield were used to generate thermal emission spectra based on Planck radiation. Both spectra were summed and integrated over the flow field. The resulting emission at each trajectory point was propagated to the DC-8 position and transformed into incident irradiance to be finally compared with experimental data.
71 citations
TL;DR: In this paper, two-dimensional axisymmetric simulations have been performed to numerically reproduce the ablation of a graphite sphere cone that has been tested in the Interaction Heating Facility at the NASA Ames Research Center.
Abstract: equilibriumablationwithsurfacemassandenergybalancesfullycoupledwiththenumericalsolverandcanaccount for both surface oxidation and sublimation. The surface temperature is obtained from the steady-state ablation approximation. This numerical procedure can predict aerothermal heating, chemical species concentrations, and carbon material ablation rate over the heat-shield surface of reentry vehicles. Two-dimensional axisymmetric simulations have been performed to numerically reproduce the ablation of a graphite sphere cone that has been tested in the Interaction Heating Facility at the NASA Ames Research Center. The freestream conditions of the selected test case are typical for Earth reentry from a planetary mission. The predicted ablation rate and surface temperatureassumingfrozenchemistryinthe flowshowagoodagreementwiththeavailableexperimentaldata.The agreementisfurtherimprovedfreezingthenitrogenrecombinationreactionatthesurfacetobemoreconsistentwith experimental observation, which has shown nitrogen atom recombination not to occur at the graphite surface.
70 citations
TL;DR: Spectroscopic observations of the 2006 Stardust Sample Return Capsule entry are presented in this paper, obtained by means of a slitless miniature echelle spectrograph onboard NASA's DC-8 airborne laboratory.
Abstract: Spectroscopic observations of the 2006Stardust SampleReturnCapsule entry arepresented, obtainedbymeans of a slitless miniature echelle spectrograph onboardNASA’s DC-8 airborne laboratory. The data cover the wavelength range from 336 to 880 nm, at 0.14–0.9 nm resolution, and were obtained during the time interval when radiative heatingwasmost important. Thedata contain abroadband continuum,presumably from the hot heat-shield surface, shock-layer air plasma emissions of N, O, andN 2 , and atomic hydrogen and CNmolecular band emission from the ablating heat-shield material, a form of phenol-impregnated carbon ablator. Early in flight, there were also atomic line emissions of Zn, K, Ca , and Na, presumably from a white Z-93P paint applied to the top of the phenolimpregnated carbon ablator. At each moment along the trajectory, the whole spectrum was recorded simultaneously, but broken into smaller segments. Key issues addressed in the data reduction and calibration are described. The interpretation of these data was given elsewhere.
TL;DR: In this paper, a configuration of an E B layer as a reentry blackout mitigation method was proposed, where the manipulated plasma region provides a possibility to communicate through a plasma layer during a Reentry blackout.
Abstract: Radio blackout that occurs during hypersonic reentry flight is an important issue for the operation of the vehicle. Since the radio blackout problem is caused by a high plasma number density around a vehicle, it is necessary to manipulate the plasma to allow communication. We suggest a configuration of an E B layer as a reentry blackout mitigationmethod.ThesuggestedE Blayerconfigurationwithatwo-dimensionalmagnetic fieldissimulatedusing thethermalized potentialmodelandthePoisson-like modelin ordertoillustrate theeffectiveness of thisapproachas amitigationmethod.Thenumericalmodelusesamagnetohydrodynamicsapproximationandissolvedusinga finite volumemethodwithaRiemannsolver.Theresultsofthenumericalmodelareassessedusingavailableexperimental results.Astrongplasmadensityreductionisobtainedwhenthehighelectricandmagnetic fieldsareappliednearthe cathode. The manipulated plasma region provides a possibility to communicate through a plasma layer during a reentry blackout.
TL;DR: In this paper, the aerodynamic effects of a spike attached to a blunt body were analyzed using schlieren flow visualization and measured aerodynamic forces and moments, and the experimental results of the research on a hemispherical blunt nose body with and without spike at L=D ratio of 1.5 and 2 (where L is the spike length and D is cylinder diameter).
Abstract: A HIGH-SPEED flow over a blunt body generates a bow shock wave in front of it, which causes a rather high surface pressure and, as a result, high aerodynamic drag. The surface pressure on the blunt body can be substantially reduced if a conical shock wave is generated by attaching a forward-facing spike. Thus, the introduction of the spike decreases the drag and increases the lift coefficient. The spike produces a region of recirculating separated flow that shields the blunt-nosed body from the incoming flow. The applicability of the spike is limited due to the possible appearance offlowoscillations in the separation region, which may reduce its positive effects and may cause aerodynamic disturbances during the flight [1]. Many experimental studies focused their attention on the influence of the spike’s length on the aerodynamic characteristics of blunt bodies for various angles of attack at some transonic [2], supersonic [3–5], or even hypersonic [6–8] speeds. This Note contributes to the experimental study of the fluid flow structure and aerodynamic characteristics of a spike attached to blunt body atMach 6. This Note analyzes the aerodynamic effects of the spike attached to the blunt body by using schlieren flow visualization and measured aerodynamic forces and moments. This Note briefly describes the experimental results of the research on a hemispherical blunt nose body with and without spike at L=D ratio of 1.5 and 2 (where L is the spike length and D is cylinder diameter), and angle of attack from 0 to 8 deg, with a 1 deg step. An in-depth description of the experiment conditions and results may be found in [9].
TL;DR: The National Aeronautics and Space Administration (NASA) has flight-tested a flush airdata sensing (FADS) system on the Hyper-X Research Vehicle (X-43A) at hypersonic speeds during the course of two successful flights as mentioned in this paper.
Abstract: The National Aeronautics and Space Administration (NASA) has flight-tested a flush airdata sensing (FADS) system on the Hyper-X Research Vehicle (X-43A) at hypersonic speeds during the course of two successful flights. For this series of tests, the FADS system was calibrated to operate between Mach 3 and Mach 8, and flight test data was collected between Mach 1 and Mach 10. The FADS system acquired pressure data from surface-mounted ports and generated a real-time angle-of-attack (alpha) estimate on board the X-43A. The collected data were primarily intended to evaluate the FADS system performance, and the estimated alpha was used by the flight control algorithms on the X-43A for only a portion of the first successful flight. This paper provides an overview of the FADS system and alpha estimation algorithms, presents the in-flight alpha estimation algorithm performance, and provides comparisons to wind tunnel results and theory. Results indicate that the FADS system adequately estimated the alpha of the vehicle during the hypersonic portions of the two flights.
TL;DR: In this paper, the efiect of ablation or surface blowing is reviewed by summarizing the experimental data, for blunt bodies and for slender bodies at zero and nonzero angles of attack.
Abstract: Hypersonic boundary-layer transition is afiected by many factors, including Mach number, Reynolds number, geometry, roughness, and tunnel noise The efiect of ablation or surface blowing is reviewed by summarizing the experimental data Blowing generally moves transition upstream, with larger mass∞ow rates or lighter gases causing a larger efiect Blowing that occurs farther upstream on the model generally also has a larger efiect It may be feasible to estimate the efiect of blowing using semi-empirical stability-based methods such as e N : Experimental data suitable for comparisons to these methods are summarized, for blunt bodies and for slender bodies at zero and nonzero angles of attack
TL;DR: In this article, the authors provide a set of requirements for the development of a solar sail propelled Interstellar Heliopause Probe mission. But, they do not consider the use of a spinning sail, limiting the direct application of current hardware development activities.
Abstract: This paper provides a set of requirements for the technology development of a solar sail propelled Interstellar Heliopause Probe mission. The mission is placed in the context of other outer solar systems missions, ranging from a Kuiper Belt mission through to an Oort cloud mission. Mission requirements are defined and a detailed parametric trajectory analysis and launch date scan performed. Through analysis of the complete mission trade space a set of critical technology development requirements are identified which include an advanced lightweight composite High-Gain Antenna, a high-efficiency Ka-band travelling-wave tube amplifier and a radioisotope thermoelectric generator with power density of approximately 12 W/kg. It is also shown that the Interstellar Heliopause Probe mission necessitates the use of a spinning sail, limiting the direct application of current hardware development activities. A Kuiper Belt mission is then considered as a pre-curser to the Interstellar Heliopause Probe, while it is also shown through study of an Oort cloud mission that the Interstellar Heliopause Probe mission is the likely end-goal of any future solar sail technology development program. As such, the technology requirements identified to enable the Interstellar Heliopause Probe must be enabled through all prior missions, with each mission acting as an enabling facilitator towards the next.
TL;DR: In this paper, the authors demonstrate that time-accurate solutions for multidimensional ablation and shape change of thermal protection system materials may be obtained by loose coupling of a high-fidelity flow solver with a material thermal response code.
Abstract: The central focus of this study is to demonstrate that time-accurate solutions for multidimensional ablation and shapechange of thermal protection system materials may be obtained by loose coupling of a high-fidelity flow solver with a material thermal response code. In this study, the flow code solves the nonequilibrium Navier–Stokes equations using the data-parallel line-relaxation (DPLR) method. The material response code is the latest version of the Two-dimensional Implicit Thermal Response and Ablation Program (TITAN). In TITAN, the governing equations, which include a three-component decomposition model and a surface energy balance with thermochemical ablation, are solved with a robust moving-grid scheme to predict the shape change caused by surface recession. Coupling between the material response and flow codes is required for many multidimensional ablation simulations, because the magnitude and distribution of the surface heat flux are very sensitive to shape change. This paper demonstrates the application of the TITAN-DPLR system to problems with large-scale recession and shape change.Ablationandthermalresponsesimulationsarepresentedforiso-qand flat-facedarc-jettestmodelsandalso for a wedge with a cylindrical leading edge exposed to hypersonic flow at various angles of attack.
TL;DR: In this article, a finite element modeling method supported by empirically determined cable properties and structural behavior was developed to characterize cable harness impacts on dynamic response in a free-free beam.
Abstract: Powerand data-handling cables, which can account for up to 30% of a satellite’s dry mass, couple with the spacecraft structure and impact dynamic response. Structural dynamicmeasurements suggest that amore complete representation of cable effects is needed to improvemodel predictive accuracy. To that end, a studywas performed to characterize cable harness impacts on dynamic response. From this study, a finite element modeling method supported by empirically determined cable properties and structural behaviorwas developed. Themodelingmethod was validatedwith a considerable amount ofmodel simulation and experimental data for a variety of cables attached to a free–free beam. At low frequencies, the cable effect was dominated bymass and stiffness, changing the apparent stiffness; damping was a secondary effect. At higher frequencies, where the cables themselves were resonant, the cable effect was dissipative, increasing the apparent damping in addition to affecting the overall frequency response. Tiedown stiffness was found to be an important, but difficult to measure, parameter. Finite element models of a cabled beamwere shown to be valid for all cable families studied.As a result, thefinite elementmodelingmethod itself was validated.
TL;DR: In this paper, the results of an analysis of the thermocouple measurements used to infer the dynamics of the transition process during the trajectories for both flights, on both the lower surface and upper surface, were presented.
Abstract: The successful Mach 7 and 10 flights of the first fully integrated scramjet propulsion systems by the Hyper-X (X-43A) program have provided the means with which to verify the original design methodologies and assumptions. As part of Hyper-X s propulsion-airframe integration, the forebody was designed to include a spanwise array of vortex generators to promote boundary layer transition ahead of the engine. Turbulence at the inlet is thought to provide the most reliable engine design and allows direct scaling of flight results to groundbased data. Pre-flight estimations of boundary layer transition, for both Mach 7 and 10 flight conditions, suggested that forebody boundary layer trips were required to ensure fully turbulent conditions upstream of the inlet. This paper presents the results of an analysis of the thermocouple measurements used to infer the dynamics of the transition process during the trajectories for both flights, on both the lower surface (to assess trip performance) and the upper surface (to assess natural transition). The approach used in the analysis of the thermocouple data is outlined, along with a discussion of the calculated local flow properties that correspond to the transition events as identified in the flight data. The present analysis has confirmed that the boundary layer trips performed as expected for both flights, providing turbulent flow ahead of the inlet during critical portions of the trajectory, while the upper surface was laminar as predicted by the pre-flight analysis.
TL;DR: In this article, the impact of 3D surface roughness elements on the laminar boundary layer of a smooth titanium hemispheres was evaluated in the NASA Ames hypersonic ballistic range through quiescent CO2 and air environments.
Abstract: Smooth titanium hemispheres with isolated three-dimensional (3D) surface roughness elements were flown in the NASA Ames hypersonic ballistic range through quiescent CO2 and air environments. Global surface intensity (temperature) distributions were optically measured and thermal wakes behind individual roughness elements were analyzed to define tripping effectiveness. Real-gas Navier-Stokes calculations of model flowfields, including laminar boundary layer development in these flowfields, were conducted predict key dimensionless parameters used to correlate transition on blunt bodies in hypersonic flow. For isolated roughness elements totally immersed within the laminar boundary layer, critical roughness Reynolds numbers for flights in air were found to be higher than those measured for flights in CO2, i.e., it was easier to trip the CO2 boundary layer to turbulence. Tripping effectiveness was found to be dependent on trip location within the subsonic region of the blunt body flowfield, with effective tripping being most difficult to achieve for elements positioned closest to the stagnation point. Direct comparisons of critical roughness Reynolds numbers for 3D isolated versus 3D distributed roughness elements for flights in air showed that distributed roughness patterns were significantly more effective at tripping the blunt body laminar boundary layer to turbulence.
TL;DR: In this paper, an overview of the reconstruction analyses performed for the Stardust capsule entry is described, which indicate that the actual entry was very close to the pre-entry predictions.
Abstract: An overview of the reconstruction analyses performed for the Stardust capsule entry is described. The results indicate that the actual entry was very close to the pre-entry predictions. The capsule landed 8.1 km north-northwest of the desired target at Utah Test and Training Range. Analyses of infrared video footage and radar range data (obtained from tracking stations) during the descent show that drogue parachute deployment was 4.8 s later than the pre-entry prediction, while main parachute deployment was 19.3 s earlier than the pre-set timer indicating that main deployment was actually triggered by the backup baroswitch. Reconstruction of a best estimated trajectory revealed that the aerodynamic drag experienced by the capsule during hypersonic flight was within 1% of pre-entry predications. Observations of the heatshield support the pre-entry estimates of small hypersonic angles of attack, since there was very little, if any, charring of the shoulder region or the aftbody. Through this investigation, an overall assertion can be made that all the data gathered from the Stardust capsule entry were consistent with flight performance close to nominal pre-entry predictions. Consequently, the design principles and methodologies utilized for the flight dynamics, aerodynamics, and aerothermodynamics analyses have been corroborated.
TL;DR: The results of the Monte Carlo analysis show that ATK-RBCC featuring airbreathing propulsion combined with a high lift-over-drag airframe exhibits significant operability benefits.
Abstract: Hypersonic airbreathing propulsion has been considered as an enhancement for access-to-space systems for decades. However, previous research usingmetrics such as takeoff gross weight and payload weight fraction has not shown conclusive benefits for airbreathing systems when compared with all-rocket launch vehicles. The U.S. Air Force Research Laboratory has developed new operability-based metrics relevant to U.S. Air Force missions: time to rendezvous with a target spacecraft, number of launch opportunities per day, and launch-window duration. Computation of the new metrics requires launch vehicle ascent trajectory optimization, orbital transfer solutions, andMonteCarlo analysis. Ascent optimization uses propulsion throttling, aerodynamic turning, andpitch control to command downrange and crossrange at the orbital insertion point while using minimum propellant. Then the twopoint boundary-value problem is solved to find a minimum-propellant transfer orbit to rendezvous with the target. Monte Carlo analysis assigns the orbital target a random starting position over the Earth and then propagates the orbit until rendezvous is accomplished and themetrics can be computed. The Air Force’s ReusableMilitary Launch System all-rocket launch vehicle RMLS 102 is compared against Alliant Techsystems’ rocket-based combined-cycle launch system ATK-RBCC. The results of the Monte Carlo analysis show that ATK-RBCC featuring airbreathing propulsion combined with a high lift-over-drag airframe exhibits significant operability benefits. The developed operability metrics could help to transform access to space by demonstrating clear payoffs from airbreathing propulsion.
TL;DR: In this paper, both computational fluid dynamics and particle simulation Monte Carlo methods were used to analyze the forebody flow of the Stardust sample return capsule at altitudes of 81 and 71 km, where the flow is in the nearcontinuum regime.
Abstract: DOI: 10.2514/1.37360 The Stardust sample return capsule entered the Earth’s atmosphere at a very energetic velocity of 12:6 km=s .I n the present study, both continuum (computational fluid dynamics) and particle (direct simulation Monte Carlo) methods are used to analyze the forebody flow of the Stardust sample return capsule at altitudes of 81 and 71 km, where the flow is in the near-continuum regime. At the higher altitude, direct comparisons between baseline computational fluiddynamicsanddirectsimulationMonteCarlomodelsgiveenormousdifferencesinbasic flowfield properties. To study the discrepancy between the solutions, a modified approach for determining the temperature used by computational fluid dynamics to control the dissociation and ionization reactions is investigated. The modified computational fluid dynamics and direct simulation Monte Carlo results are in significantly better agreement with each other, illustrating the strong sensitivity to chemistry modeling under these highly energetic conditions. Significant differences persist in temperatures near the capsule surface and in surface heat flux. Evaluation of local Knudsen numbers indicates that the flow experiences noncontinuum behavior in the shock front andatthecapsulesurfacethatexplainsthesmallerheat fluxpredictedbydirectsimulationMonteCarlo.Atthelower altitude, the flowfield results become less sensitive to details of the chemistry modeling, although noncontinuum effects are again predicted at the stagnation point.
TL;DR: In this paper, the results of two demonstrations of magnetic flux-pinning technologies implemented on CubeSat-sized spacecraft during microgravity flights as part of the NASA Glenn Research Center Facilitated Access to the Space Environment for Technology Development and Training (FAST) program in August 2009 were reported.
Abstract: M AGNETIC flux pinning, a noncontacting interaction between Type II superconductors andmagneticfields, has been studied at length by the scientific community for its applications to levitating objects in a 1g environment [1–3]. However, due to the unpowered passive stability that flux pinning can provide, it also has many potential applications for the assembly and reconfiguration of modular space structures [4] and spacecraft formations [5]. Current approaches to autonomous docking of space vehicles [6,7], aswell as spacecraft reconfiguration and formation flying [8–10], rely heavily on active controllers. However, a permanent magnet flux-pinned to a superconductor experiences a passive restoring force that attracts it to the position and orientation it held when the superconductor first cooled below its critical temperature. Previous work and laboratory experiments have suggested that this passively stable effect provides sufficient stiffness and damping to bind modular spacecraft together over separation distances up to about 10 cm [11]. Several possible means for actuation of a flux-pinning interface may be superimposed on the passive stability of this interaction, which requires no power to the superconductor except that required for cooling. In addition to actuation by time-varyingmagnetic fields such as those from electromagnet coils [12], aflux-pinned space systemcan exploit symmetries in the pinned magnetic field to form a noncontacting kinematic mechanism in which the modular components do not touch one another but have some specified kinematic degrees of freedom (DOF) [13]. A simple noncontacting mechanism consisting of a single revolute joint on an air-table testbed has been demonstrated in a laboratory setting [14]. This note reports the results of two demonstrations of magnetic flux-pinning technologies implemented onCubeSat-sized spacecraft during microgravity flights as part of the NASA Glenn Research Center Facilitated Access to the Space Environment for Technology Development and Training (FAST) program in August 2009. In the first experiment, a CubeSat mockup was flux pinned to a CubeSatscale vehicle carrying superconductors and was expected to demonstrate low-stiffness, noncontacting, passive station-keeping in 6 degrees of freedom (6 DOF). The second experiment studied the reconfiguration of two CubeSat mockups between equilibrium configurations via a revolute joint formed by a flux-pinned noncontacting kinematic mechanism. It was expected that the spacecraft would move about an axis defined by the flux-pinned interface rather than their respective centers of mass. These microgravity flight results highlight the role magnetic flux pinning might play in future small satellite operations. Each experiment was performed on a microgravity aircraft with two free-floating modules: one containing an array of magnets appropriate to the experiment, and the other containing superconductors in a Dewar of liquid nitrogen. Three experimenters participated in each flight, two equipment managers to monitor the position of the free-floating modules at all times, and one data collector who operated the motion-capture camera. Figure 1 is a diagram of the test setup.
TL;DR: In this article, an assessment of the hypersonic Ludwieg tube of Delft University of Technology is given, and an experimental evaluation is performed to infer the facility performance.
Abstract: An assessment of the hypersonic Ludwieg tube of Delft University of Technology is given. The facility is described theoretically, and an experimental evaluation is performed to infer the facility performance. Experiments are performed using conventional techniques such as static and total head pressuremeasurements andFay–Riddell heat flux evaluations by means of infrared thermography. Furthermore, particle image velocimetry is used to deduce nozzle boundary-layer parameters aswell as the freestreamvelocityfield and the static and total temperatures for the Mach 7 nozzle. For the Mach 9 nozzle, stagnation heat flux measurements are performed to obtain the total temperature of the flow. The freestream values were determined experimentally in two different ways and the results showed good agreement. The application of particle image velocimetry allows the freestream flowfield to be directly obtained and gives a directmeasure for the flowfield uniformity (0.2%) and repeatability (0.4%). The static and total temperatures calculated from the particle image velocimetry results showed that there is a large mismatch between the theoretical total temperature and measured total temperature, which is attributed to heat losses present in the throat and nozzle.
TL;DR: In this article, the effect of the canard on the elevon was investigated and the results showed that higher Mach numbers combined with higher canard detection angles resulted in a greater effect on the elevation of the vehicle.
Abstract: : Airbreathing hypersonic cruise vehicles are typically characterized by long, slender bodies with highly coupled engines and airframes. For a case in which the engine is underslung (below the center of gravity), a large elevon control surface is typically necessary to trim the vehicle. The elevon is usually placed at the rear of the vehicle to yield a large moment arm. However, the drawback is that the elevons can cause large perturbations in lift and other undesirable effects. Canard control surfaces are placed on the forebody of the vehicle to counteract these effects as well as aid in low-speed handling. This study looks at how the canards affect the flow over the elevon control surfaces and, in turn, the controllability of the vehicle in general. A two-dimensional analytical formulation is developed and compared with both a series approximation solution and a computational fluid dynamics Euler flow field solution. The effect of the canard on the elevon, measured using the elevon effectiveness ratio, decreased as the distance between the control surfaces increased. In general, higher Mach numbers combined with higher canard detection angles resulted in a greater effect on the elevon. Adding a thickness correction, as opposed to assuming that the airfoils were at plates, actually decreased, on average, the accuracy of the model when compared with the computational data.
TL;DR: In this paper, a high degree and order spherical-harmonicmagnetic field model is used to partition the space of latitude in a meaningful way, and a successful maneuver developed within this bang-off control framework results in a combined orbital plane change and orbit raising.
Abstract: Anorbital control framework is developed for theLorentz augmented orbit. A spacecraft carrying an electrostatic chargemoves through the geomagnetic field. The resulting Lorentz force is used in the general control framework to evolve the spacecraft’s orbit. The controller operateswith a high degree and order spherical-harmonicmagneticfield model by partitioning the space of latitude in a meaningful way. The partitioning reduces the complexity of the problem to amanageable level. A successful maneuver developed within this bang-off control framework results in a combined orbital plane change and orbit raising. The cost of this maneuver is in electrical power. Reductions in the power usage, at the expense of longermaneuver times, are obtained by using information about local plasma density.