Showing papers in "Journal of Spacecraft and Rockets in 1997"

Collaborative Approach to Launch Vehicle Design

Abstract: Collaborative optimization is a new design architecture specie cally created for large-scale distributed-analysis applications. In this approach, a problem is decomposed into a user-dee ned number of subspace optimization problems that are driven toward interdisciplinary compatibility and the appropriate solution by a system-level coordination process. This decentralized design strategy allows domain-specie c issues to be accommodated by disciplinary analysts while requiring interdisciplinary decisions to be reached by consensus. This investigation focuses on application of the collaborative optimization architecture to the multidisciplinary design of a singlestage-to-orbit launch vehicle. Vehicle design, trajectory, and cost issues are directly modeled in this problem, which is characterized by 95 design variables and 16 constraints. Numerous collaborative solutions are obtained. Comparison of these solutions demonstrates the ine uencethat an a priori ascent-abort criterion has on the vehicle design and the distinction between minimum weight and minimum cost concepts. The operational advantages of the collaborative optimization architecture include minimal framework integration requirements, the ability to use domain-specie c analyses, which already provide optimization without modie cation, inherent system e exibility and modularity, a distributed analysis and optimization capability, and a signie cant reduction in interdisciplinary communication requirements.

Galileo Probe Heat Shield Ablation Experiment

Abstract: The Galileo probe deceleration module contained an experiment which measured the recession of the forebody heat shield during the ablative probe entry into the Jovian atmosphere. A detailed description of the experiment, reduction of the probe data, reconstruction of the heat-shield shape, and comparisons with preflight predictions are presented. Data quality is good during the second half of recession, but excessive noise levels at the onset of ablation prevented the experiment from meeting two of three performance requirements. The recession distribution is surprisingly dissimilar from preflight predictions. Measured recession was less than predicted near the stagnation point, but exceeded predictions over most of the frustum. (Author)

Advanced Propulsion for Geostationary Orbit Insertion and North-South Station Keeping

Abstract: Solar electric propulsion technology is currently being used for geostationary satellite station keeping. Analyses show that electric propulsion technologies can be used to obtain additional increases in payload mass by using them to perform part of the orbit transfer. Three electric propulsion technologies are examined at two power levels for geostationary insertion of an Atlas IIAS class spacecraft. The onboard chemical propulsion apogee engine fuel is reduced in this analysis to allow the use of electric propulsion. A numerical optimizer is used to determine the chemical burns that will minimize the electric propulsion transfer times. For a 1550-kg Atlas IIAS class payload, increases in net mass (geostationary satellite mass less wet propulsion system mass) of 150-800 kg are enabled by using electric propulsion for station keeping, advanced chemical engines for part of the transfer, and electric propulsion for the remainder of the transfer. Trip times are between one and four months.

Three-Dimensional Aerodynamic Shape Optimization Using Genetic and Gradient Search Algorithms

Abstract: This study introduces various gradient search methods as well as hybrid genetic techniques that achieve impressive convergence rates on constrained problems. These methods are applied to 3D shape optimization of ogive-shaped, star-shaped, and spiked projectiles and lifting bodies in a hypersonic flow. Flow-field analyses are performed using Newtonian flow theory and in certain cases verified using a parabolized Navier-Stokes (PNS) flow analysis algorithm. Three-dimensional geometrical rendering is achieved using a variety of techniques including beta-splines from the computer graphics industry. (Author)

Aerodynamics of Stardust Sample Return Capsule

Abstract: Successful return of interstellar dust and cometary material by the Stardust Sample Return Capsule requires an accurate description of the Earth entry vehicle''s aerodynamics. This desciption must span the hypersonic-rarefied, hypersonic-continuum, supersonic, transonic, and subsonic flow regimes. Data from numerous sources are compiled to accomplish this objective. These include Direct Simulation Monte Carlo analyses, thermochemical nonequilibrium computational fluid dynamics, transonic computational fluid dynamics, existing wind tunnel data, and new wind tunnel data. Four observations are highlighted: 1) a static instability is revealed in the free-molecular and early transitional-flow regime due to aft location of the vehicle''s center-of-gravity, 2) the aerodynamics across the hypersonic regime are compared with the Newtonian flow approximation and a correlation between the accuracy of the Newtonian flow assumption and the sonic line position is noted, 3) the primary effect of shape change due to ablation is shown to be a reduction in drag, and 4) a subsonic dynamic instability is revealed which will necessitate either a change in the vehicle''s center-of-gravity location or the use of a stabilizing drogue parachute.

High-Enthalpy Aerothermodynamics of a Mars Entry Vehicle Part 1: Experimental Results

Abstract: Aerodynamic heating tests were conducted on a 70-deg sphere ‐cone Mars entry vehicle cone guration in a high-enthalpy impulse facility in both carbon dioxide and air test gases. The purpose of these tests was to obtain heat transfer data for comparison with results of Navier ‐Stokes computations. Surface heat transfer rates were determined for both the forebody and afterbody of the test models and for the stings that supported the models in the facility test section. Little difference was observed between normalized heating distributions for the air and carbon dioxide test conditions. For both cases, peak sting heating was on the order of 4 ‐5% of the forebody stagnation-point heating, and it was concluded that the wake e ow remained laminar. The wake e ow establishment process was quantie ed and was found to require approximately 40 ‐70 e ow path lengths, which corresponded to approximately 75% of the available facility test time. The repeatability of facility test conditions was estimated to vary between § 3% and § 10%. The overall experimental uncertainty of the data was estimated to be § 10‐11% for forebody heating and § 17‐22% for wake heating.

Evolution of Secondary Electron Emission Characteristics of Spacecraft Surfaces

Abstract: Helical Motion In this motion, the center of gravity of the  ight vehicle traverses a helical  ight path.The longitudinalaxis of the vehicle is orientated in the same directionas the axis of the helix but displacedfrom it by a constantdistance.A Ž xed body axial spin rate equal to the angular rate of rotation of the helical  ight path is present. The resulting motion is sometimes referred to as lunar motion because the same surface of the body is always orientated toward the central axis of the maneuver; in aircraft maneuver terminology it is equivalent to a barrel roll. A three-dimensional view of the motion is given in Fig. 4. Again, with gravity neglected, we can write u D w D 0; P u D P v D P w D 0, and q D r D 0. It follows that XF D YF D 0 and that ZF D pv. Because is now 0 deg, d3⁄4=dt ; P ®, and R are all zero. This result is again consistent with that predicted by Eqs. (4–6), (10), and (11). Note that, because of the presence of axial spin, this is not the same case as that considered in Ref. 2.

Low-Density Aerodynamics of the Stardust Sample Return Capsule

Abstract: The aerodynamics of the Stardust Sample Return Capsule are analyzed in the low-density, transitional flow regime using free-molecular, Direct Simulation Monte Carlo, Navier-Stokes, and Newtonian methods to provide inputs for constructing a transitional flow bridging relation. The accuracy of this bridging relation in reconstructing the aerodynamic coefficients given by the more exact methods is presented for a range of flight conditions and vehicle attitudes. There is good agreement between the various prediction methods, and a simple sine-squared bridging relation is shown to provide a reasonably good description of the axial force, normal force, and pitching moment over a range of Knudsen numbers from 0.001 to 10. The predictions show a static instability of the Stardust capsule in the free-molecular regime that persists well into the transitional flow. The addition of a thin disk to the base of the capsule is shown to remove this static instability. However, the extremely high entry velocity of 12.6 km/s for the proposed trajectory introduces difficult design issues for incorporating this disk caused by the high aerothermal loads that occur even under relatively rarefied conditions.

Enthalpy effects on hypervelocity boundary-layer transition : ground test and flight data

Abstract: Boundary-layer-transitionexperiments on a 5-deg half-anglecone at 0-deg angle of attackwere performed in the T5 hypervelocity shock tunnel. The test gases investigated included air, nitrogen, and carbon dioxide. Reservoir enthalpies were varied from 3 to 27MJ/kg and reservoir pressures from 10 to 95MPa, depending on the gas and tunnel settings. No clear relationship is found to exist between the transition Reynolds number based on the boundary-layer-edge conditions and the reservoir enthalpy. However, when the reference temperature conditions are used instead, the different test gases are distinguishable and ordered according to their dissociation energy. Data from a free- ight experiment are also compared with the shock tunnel experiments. When the transition Reynolds numbers are evaluated at the boundary-layer-edge conditions, they are an order of magnitude higher than the tunnel results. However, when the reference conditions are used, the  ight data fall within the same range as the experiments, although the trend with reservoir enthalpy is reversed.

Navier-Stokes Predictions of Pitch Damping for Axisymmetric Projectiles

Abstract: An approach for predicting the pitch-damping coefe cient sum for axisymmetric e ight bodies is presented. The approach utilizes a specie c combination of spinning and coning motions that allows the pitch-damping force and moment coefe cient to be directly related to the aerodynamic side force and moment. The use of combined spinning and coning motion represents an improvement over existing techniques that utilize lunar coning motion for predicting the pitch-damping coefe cients. A parabolized Navier ‐Stokes approach that utilizes a missile-e xed, noninertialrotatingcoordinateframeisappliedtopredictthee owe eldsaboutaxisymmetricprojectilesundergoing steady coning motion. Thegoverning equationsaremodie edto includethecentrifugal and Coriolisforcetermsdue to the rotating coordinate frame. From the computed e owe eld, the side force and moment due to coning motion, spinning motion, and combined spinning and coning motion are used to determine the pitch-damping coefe cients. Computationsareperformedfora genericshell cone guration (with and withoutboattail ), andthepredictionsshow good agreement with an existing inviscid code. The comparisons of computational results for a family of ogivecylindercone gurationswithaerodynamicsrangedatashowexcellentagreementandfurthervalidatetheapproach.

Periodic Optimal Cruise for a Hypersonic Vehicle with Constraints

Abstract: Anew approach is presented to solve periodic optimal control problems. As an applicationof the approach, fueloptimal periodic control problems for a hypersonic vehicle are solved. The model of the vehicle was constructed by using numerical data and Ž gures from available space plane literature. In particular, heating-rate and load-factor constraints are considered tomake thismodelmore realistic thanotherpreviousmodels.These constraints increase the difŽ culty in obtaining a numerical solution and also increase the sensitivity to the initial guess for convergence. By assuming the shape of an altitude proŽ le as a sinusoidal function of range and by using a bang-bang thrust control, a suboptimal solution is obtained for the vehicle without any constraints. This suboptimal solution serves as a very good initial guess for the optimal solution generated by theminimizing-boundary-conditionmethod. The optimal solution shows a fuel saving of 8.12% over the steady-state cruise, with a maximumheating rate of 1202.4 W/cm2 , andwith a maximumload factor of 8.27. Constraints for a heating rate and a load factor are then added to the problem.With amaximumheating rate of 400W/cm2 , the fuel saving reduces to 2.45%.With a load factor of 7, the fuel saving does not change much from the nonconstrained solution. An optimal periodic-cruise solution with maximum heating rate of 1158.0 W/cm2 and simultaneously with maximum load factor of 7 is also determined with a fuel saving of 8.09%.

Fluid Dynamics of Hypersonic Forward-Facing Cavity Flow

Abstract: Hypersonic  ow over the nose of a blunt body with a forward-facing cylindrical cavity is studied. Extensive numerical results involvinga wide range of cavity depths are veriŽ ed by experimental runs including a new set of runs performed in a quiet  ow supersonic tunnel at Mach 4. It is shown that freestream noise is the mechanism that drives resonant pressure oscillations within relatively shallow cavities. Numerical results and conventional tunnel experiments show that deeper cavities oscillate strongly without freestream noise. For shallow cavities the sensitivity of cavity ampliŽ cation, i.e., the relative strength of resonant pressure oscillations, to the characteristics of freestreamnoise (frequency, amplitude,perturbationvariable), cavitygeometry (depth, geometric scale, lip radius), Mach number, viscous effects, and thermal wall boundary condition is studied. There is a strong dependence of oscillation strength on freestream noise frequency. Oscillation strength increases nearly proportionally to input noise amplitude, increases rapidly with cavity depth, and increases with Mach number but levels off at highMach numbers. The pressure oscillations exhibit behavior analogous to that of a damped harmonic oscillator.

Three-Dimensional Hypersonic Laminar Boundary-Layer Computations for Transition Experiment Design

Abstract: The stability of boundary layers on sharp-nosed cones with elliptical cross sections is assessed using linear stability theory and crosse ow correlations. The objective is to identify a cone guration for wind-tunnel testing that exhibits signie cant crosse ow but also possesses a sufe cient laminar region for boundary-layer stability probing. Parabolized Navier ‐Stokes computer codes were used to calculate the mean e ow about cones with eccentricities of 1.5:1, 2.0:1, and 4.0:1 at a freestream Mach number of 7.95 and freestream unit Reynolds number of 3:3 £ 10 6 m i 1 . Correlations indicated that transition was possible on each cone guration at the above conditions. All three cone gurations showed unstable, ine ectional velocity proe les and boundary-layer thickening along the centerline (minor axis) due to the ine ux of low-momentum e uid. Crosse ow separation was observed on the 2.0:1 cone guration. Linear stability theory was used to calculate stationary crosse ow N factors on all three cone gurations, and to calculate traveling-wave N factors on the 1.5:1 and 2.0:1 cone gurations. All three cone gurations showed crosse ow instability, with the 4.0:1 cone guration attaining the highest N factors. The 1.5:1 and 2.0:1 cone gurationswere unstable to a broad spectrum of traveling waves, with thehighest Nfactors attained on centerline, due to the unstable proe les there.

Cassini/Huygens Science Instruments, Spacecraft, and Mission

Abstract: The Cassini spacecraft will take 18 scientific instruments to Saturn. After launch and a seven-year cruise, Cassini will arrive at Saturn and separate into a Saturn orbiter and an atmospheric probe, called Huygens, which will descend to the surface of Titan. The orbiter will orbit the planet for four years, making close flybys of five satellites, including multiple flybys of Titan. Communication with Earth is at X-band; the maximum downlink rate from Saturn is 166 x 10(exp 3) bps. Orbiter instruments are body mounted; the spacecraft must be turned to point some of them toward objects of interest. The orbiter carries 12 instruments. Optical instruments provide imagery and spectrometry. Radar supplies imaging, altimetry, and radiometry. Radio links contribute information about intervening material and gravity fields. Other instruments measure electromagnetic fields and the properties of plasma, energetic particles, and dust particles. The probe is spin stabilized. It returns data via an S-band link to the orbiter. The probe's six instruments include sensors to determine atmospheric physical properties and composition. Radiometric and optical sensors will produce data on thermal balance and obtain images of Titan's atmosphere and surface. Doppler measurements between probe and orbiter will provide wind profiles. Surface sensors will measure impact acceleration, thermal and electrical properties, and, if the surface is liquid, density and refractive index. This design will enable Cassini to determine the composition; the physical, morphological, and geological nature; and the physical and chemical processes of the atmospheres, surfaces, and magnetosphere of the Saturnian system. This paper briefly describes the Cassini mission and spacecraft and, in somewhat more detail, the scientific instruments.

Systems Engineering Trades for the IRIDIUM Constellation

Abstract: The customer imperative has driven the system design for the IRIDIUM® constellation and ground support infrastructure. Although the goal of the system is to provide global, mobile telephone service, the two principal drivers for the trade studies are 16-dB link margin and commercial success. The systems engineering and constellation architecture trade studies that determined the design of the IRIDIUM telecommunications system were comprehensive and are continuously evaluated. The complexity of the total system drove the engineers to critical evaluations of major factors such as complexity of payload, size/weight of spacecraft, altitude of constellation, size of individual beams, space environment (radiation belts), size of telephone (pocket size), quality of voice/data, number of telephone calls, type of commercial launch vehicles, number of satellites, orbits and inclination of orbits, number of gateways, crosslink types, and type of modulation. The complexity of the decisions and how they trade off to one of the two basic principles of customer satisfaction, 16-dB link margin and commercial success, is discussed.

Experimental Rarefied Heat Transfer at Hypersonic Conditions over 70-Degree Blunted Cone

Abstract: Surface heat transfer rates have been measured over a 70-deg spherically blunted cone chosen as a test case model to provide both experimental and computational databases. Under raree ed and hypersonic conditions, heating rate distributions are measured along the model and presented at angles of attack varying from 0 to 30 deg. Experiments have been conducted in the SR3 facility at a freestream Mach number close to 20. Three e ow rarefactions have been considered, which correspond to Reynolds numbers of 1420, 4175, and 36,265. Reynolds numbers are calculated using freestream conditions and the model base diameter. In parallel to the experimental work, e ow calculations were executed by an international group of researchers for identical test conditions. Comparisons between experimental and computational heating rates are also presented.

Aerodynamics of the Mars Microprobe Entry Vehicles

Abstract: The selection of the unique aeroshell for the Mars Microprobes is discussed. A description of its aerodynamics in hypersonic rarefied, hypersonic continuum, supersonic, and transonic flow regimes is then presented. This description is compiled from Direct Simulation Monte Carlo simulations, computational fluid dynamics, wind tunnel data, and ballistic range data. The aeroshell is shown to possess the correct combination of aerodynamic stability and drag to convert the probe''s initial tumbling attitude at atmospheric-interface into the desired surface-impact orientation and velocity.

High-Enthalpy Aerothermodynamics of a Mars Entry Vehicle Part 2: Computational Results

Abstract: Numerical solutions for hypersonic  ows of carbon dioxide and air around a 70-deg sphere–cone Mars entry vehicle conŽ guration were computed using a laminar, axisymmetric, nonequilibrium Navier–Stokes solver with freestream  ow conditions equivalent to those of aerothermodynamic tests conducted in a high-enthalpy impulse facility The wake  owŽ eld computations were found to be much more sensitive to both grid resolution and grid adaptationthan the forebody results The wake computations showed the existence of a region of separated, steady, recirculating  ow behind the vehicle Whereas the rapid expansion of the  ow around the corner of the vehicle resulted in a wake that was mostly frozen both chemically and vibrationally, the degree of  ow expansion was not great enough to produce noncontinuum ow behavior Comparisons between computational and experimental surface heating distributionswere within the estimated experimental uncertainty for both cases except around the forebody stagnation point and the free-shear-layer reattachment point for the air case and within a small portion of the wake recirculation vortex for the carbon dioxide case

Control Effectiveness of a Jet-Slender Body Combination at Hypersonic Speeds

Abstract: The effects of jet control on the aerodynamic characteristics, performance, and stability of a 5-deg semiangle slender conemissilecone guration arestudied. Tests were made in theCrane eld University College of Aeronautics hypersonicgun tunnelusingboth sharp and blunted cones. Thestudywasconducted at a Mach number of 8.2and a Reynolds number of 392 700, based on base diameter, at pitch angles of 15 to 15 deg. The boundary layer was laminar.Airwasusedastheworkinggasforboththefreestreamandthesonicjet.Thetestsemployedschlierenphotographytostudytheoveralle owe eld.Quantitativestudiesoftheeffect of thejethavebeenmadebypressuremeasurements.Theforcesweremeasured withathree-component balanceequippedwith semiconductorstraingauges.

of Plasma Contactor Performance

Abstract: A hollow cathode-based plasma contactor will bee own on the international spacestation to control the station’ s potential to within 40 V of the local ionosphere. Extensive testing of the plasma contactor has been conducted in vacuum facilitiesat theNASA LewisResearch Center. Signie cant performance differenceswereobserved between testsofthesameplasmacontactorindifferentfacilities.Whymeasuredplasmacontactorperformancediffersinthe laboratory in different tank environmentsand how the plasma contactor performance measured in the laboratory relatestoexpected performanceinspaceisaddressed. Presented aremodelsof plasmacontactorplasma generation and interaction in a laboratory environment, including anode area limiting. These models were integrated using the Space Station Environment Work Bench to predict plasma contactor operation, and the results are compared with the laboratory measurements. Nomenclature F = gas e ow rate, standard cubic centimeter per minute ID = total orie ce electron current, A Iemission = orie ce current emitted, A Ikeeper = keeper electrode current, A Iloss = ion loss rate, A Imax = maximum possible electron current, A Iprod = total ion production rate, A

Numerical Investigation of Supersonic Base Flow with Base Bleed

Abstract: : A zonal, implicit, time-marching Navier-Stokes computational technique has been used to compute the turbulent supersonic base flow over a cylindrical afterbody with base bleed. A critical element of calculating such flows is the turbulence model. Two eddy viscosity turbulence models have been used in the base region flow computations. These models include an algebraic turbulence model and a two-equation k-e model. The k-e equations are solved using an implicit algorithm, and calculations with the k-e model are extended up to the wall. Flow field computations have been performed for a cylindrical afterbody at M approx. = 2.46 and at an angle of attack of a = 0. The results are compared to the experimental data for the same conditions and the same configuration. Details of the mean flow field as well as the turbulence quantities have been presented. In addition, the computed base pressure distribution has been compared with the experiment. In general, the k-e turbulence model performs better in the near wake than the algebraic model and predicts the base pressure much better.

Solid Rocket Motor Exhaust Model for Alumina Particles in the Stratosphere

Abstract: Based on available and new data, a unie ed model is presented for the particle size distribution, particle density, andgeometricaldispersionforthealuminaparticlesintheexhaustofsolidrocketmotorplumesinthestratosphere. The particle size distribution is trimodal with Sauter mean diameters of 0.056, 1.0, and 3.6 πm. Nearly all of the particles lie within the small-size mode but nearly all of the mass lies in the large-size mode. Approximately twothirds of the particle surface area available for heterogeneous chemical reactions is due to the large particle mode while most of the remaining surface area is due to the small particle mode. The early horizontal dispersion rate of the plume is found to be about an order of magnitude greater than the dispersion rates used in several recent models of stratospheric ozone-plume chemistry.

Experimental Rarefied Density Flowfields at Hypersonic Conditions over 70-Degree Blunted Cone

Abstract: At raree ed e ow regimes, e owe eld investigations have been conducted in the SR3 wind tunnel on a 70-deg sphericallybluntedconeanddensitye owe eldsobtained bynonintrusiveelectronbeam e uorescencemeasurements. Theblunted cone, chosen as thetest case model, has been thesubject of extensive studiesduring thepast few years. In addition to some limited results already presented at the Fourth European High-Velocity Database Workshop, the present research gathersdensity e owe elds obtained experimentally at two raree ed hypersonice ow conditions. Experiments have been performed at a freestream Mach number close to 20 and for two Reynolds numbers, 1420 and 4175, calculated using freestream conditions and the cone base diameter. Density e owe elds are presented for the two angles of attack, 0 and 10 deg, of the cone. In parallel to the experimental work presented, a number of e owe eld calculations were executed by an international group of researchers for test conditions identical to SR3 test conditions. Flowe elds were calculated using direct simulation Monte Carlo solutions, leading to comparisons between experimental and computational e owe elds.

Computational Analysis and Validation for Lateral Jet Controlled Missiles

Abstract: Comparison of three-dimensional, viscous, turbulent Navier–Stokes simulation for generic missile bodies with wind-tunnel tests have been performed with and without lateral jet thrusters for  ow Mach number range of 2–5 and angles of attack of 0–20 deg. Computationalresults show good overall engineering predictive capability for the surface pressure, normal force coefŽ cient, and jet interaction effects. Further analysisof the computationalresults shows that the favorable upstream pressure zone (lambda zone) created by the lateral jet, unfavorable pressure loss behind the jet caused by its blockage effect, and the jet wraparound effect are the three principal competing physical mechanisms that inhibit or enhance the jet ampliŽ cation factor. Canted jet studies to enhance the Ž rst of these effects show substantial increase in favorable pressure; however, it does not recover axial component of the thrust vector. Qualitative computations for multijets and hot/binary gas thruster jets have been presented to demonstrate the overall computational capability for missile design applications.

Fundamentals and Issues in Launch Vehicle Design

Abstract: The fundamentals of launch vehicle design are examined using simpliŽ ed single-stage, two-stage, and Space Shuttle performance equations. The single-stage-to-orbit launch vehicle is very sensitive to the performancecritical parameters of mass efŽ ciency, propulsion efŽ ciency, and loss management. Cost and operations coupled in the performance equation further complicates the design process. Launch vehicle design is optimized when the performance and programmaticdrivers are balanced. Programmaticdrivers include affordability, reusability, operability,abort/safety, and reliability. The issues, disciplines, and potential problems that characterize the building of a future launch system are presented. The history of the Space Shuttle is used as the benchmark example. Robustness is the key to uncoupling the design factors so that optimization can occur, but typically robust designs deŽ ne low-performance systems. Future space launch vehiclesmust develop new technologies to reshape the design parameter sensitivities of the robustness and performance functions.

Trajectory based validation of the Shuttle heating environment

Abstract: Chemically reacting,three-dimensional,fullNavier ‐Stokescalculationsaregenerated around theshuttleorbiter and are compared with the STS-2 e ight database at eight trajectory locations. Numerical estimates of quantities necessary for thermal protection system design, surface temperature and heating proe les, integrated heat load, bond-line temperatures, and thermal protection system thicknesses arecompared with theSTS-2 shuttledata. The effects of surface kinetics, turbulence, and grid resolution are investigated. It is concluded that trajectory-based thermal protection system sizing, the use of a Navier ‐Stokes e ow solver combined with a conduction analysis applied over an entry trajectory, is a benee cial tool for future thermal protection system design. This conclusion is based on a reasonable agreement between the e ight data and numerical predictions of surface heat transfer and temperature proe les, integrated heat loads and bond-line temperatures at most of the wind-side thermocouples. Theeffectsofturbulentheating on thermalprotection system designareillustrated.Forfuturelargeentry vehicles, it is concluded that the prediction of turbulent transition will be a major driver in the thermal protection system design process. Finally,onepotential payoff ofusingtrajectory-based thermalprotection system sizing,a reduction in thermal protection system mass, is illustrated. Nomenclature CT = heat transfer coefe cient, W/m 2 -K cp = heat capacity of solid cs

Whole-Spacecraft Passive Launch Isolation

Abstract: : A spacecraft is subjected to very large dynamic forces from its launch vehicle during its ascent into orbit. These large forces place stringent design requirements on the spacecraft and its components to assure that the trip to orbit will be survived. The severe launch environment accounts for much of the expense of designing, qualifying, and testing satellite components. Reduction of launch loads would allow more sensitive equipment to be included in missions, reduce risk of equipment or component failure, and possibly allow the mass of the spacecraft bus to be reduced. These benefits apply to military as well as commercial satellites. This paper reports the design and testing of a prototype whole-spacecraft isolation system which will replace current payload attach fittings, is passive-only in nature, and provides lateral isolation to a spacecraft which is mounted on it. This isolation system is being designed for a medium launch vehicle and a 6500 lb spacecraft, but the isolation technology is applicable to practically all launch vehicles and spacecraft, small and large. The isolator significantly reduces the launch loads seen by the spacecraft. Follow-on contracts will produce isolating payload attach fittings for commercial and government launches.

Steady-State and Frequency Response of a Thin-Film Heat Flux Gauge

Abstract: Anewand simplerdesign ofthin-e lm heate ux gaugehasbeendeveloped forusein high-heat-e ux environments. Heat e ux gauges of the same design were fabricated on three different substrates and tested. The heat e ux gauge comprises a thermopile and a thermocouple junction, which measures the surface temperature. The thermopile has 40 pairs of S-type thermocouples and is covered by two thermal resistance layers. Calibration and testing of thesegaugesweree rst carriedoutin an arc-lamp calibrationfacility.Sensitivityofthegaugewasdiscussed in terms of the relative conductivity and surface temperature. The heat e ux calculated from the gauge output was in good agreement with the precalibrated standard sensor. The steady-state and the transient response characteristics of the heat e ux gauge were also investigated using a carbon dioxide pulse laser as a heat source. The dynamic frequency response was evaluated in terms of the nondimensional amplitude ratio with respect to the frequency spectrum of a chopped laser beam. The frequency response of the gauge was determined to be about 3 kHz. The temperature proe les in the thin-e lm heat e ux gauge were obtained numerically in steady-state conditions using FLUENT and compared with the experimental results. Nomenclature d = thickness of thermal resistance layer, m Es = thermopile voltage output, mV K = thermal conductivity, W/m C K = relative conductivity, W/m C N = number of thermocouple pairs in the thermopile Q = heat e ux, W/m 2 S T = absolute thermoelectric power at temperature T, mV/ C T = temperature, C t = time, s = absorptivity d = thickness difference in thermal resistance layers between 1 and 2 Subscripts 1, 2 = thermal resistance layers 1 and 2, respectively

Validation Database for Propulsion Computational Fluid Dynamics

Abstract: Validation of computational e uid dynamic (CFD) models appropriate for subsonic through hypersonic e ow applicationsrequirescarefulconsiderationofthephysicalprocessesencounteredinthesee ightregimesanddetailed comparisons of the calculated results with experimental data sets that include these processes. The work reported involved two efforts: 1 ) identie cation of quality data sets for establishment of a standard validation database and 2) direct comparison of the CFD model results with the measurement database. Three Navier ‐Stokes-based CFD models were selected for comparison: the three-dimensional, upwind-differenced, e nite volume e ow solver; the generalized implicite owsolver;and thegeneralaerodynamicsimulationprogram.Computede owpropertiesfrom these three models were compared to experimental data from seven selected databases. These databases include measurements from the following experiments: 1 ) supersonic e ow over a rearward-facing step, 2 ) supersonic, two-dimensional nozzle e ow, 3 ) low subsonic, reacting nozzle e ow, 4 ) combustion in two-dimensional, supersonic e ow with tangential hydrogen injection, 5 ) shear-layer combustion in a supersonic concentric hydrogen/air e ow, 6) hypersonic e ow over a biconic model with perpendicular nitrogen injection, and 7 ) sonic, normal injection of staged N2 jets behind a rearward-facing step into a Mach number 2 airstream. The results of the study indicate that all three models compared reasonably well with e ow measurements from the seven validation cases.