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At least 253 records · Page 14

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NASA NTRS2020

Physical Simulation Of Rocket Exhaust Aerodynamics Using Heated Ethane: Prototypical Experiments

An inaugural experimental investigation into the use of heated ethane as a motive fluid for subscale simulation of rocket exhaust plume aerodynamics has been conducted. A small, self-contained, hydrocarbon aerodynamics test stand was designed and fabricated for this purpose. Testing was performed on an aerodynamic system comprised of a 77:1 upper stage nozzle contour coupled with a passive second-throat supersonic diffuser, specifically selected for its known LOX/GH2 hot-fire performance and complex shock structure. Effects of ethane purity and stagnation temperature on the accuracy of flow field replication have been examined. All major flow features present in hot-fire tests were reproduced in heated ethane with start and unstart pressure ratio errors of approximately +4% and -5%, respectively. Steady-state test cell pressure ratio errors of ±0.5% were shown to be readily achievable when the stagnation temperature was well-controlled. Ambient-temperature nitrogen testing was also conducted for the purpose of direct comparison of the ethane method to conventional cold-flow techniques. However, its higher isentropic exponent prevented the diffuser from achieving start.

rocket↗

NASA NTRS1974

Viking entry aerodynamics and heating

The characteristics of the Mars entry including the mission sequence of events and associated spacecraft weights are described along with the Viking spacecraft. Test data are presented for the aerodynamic characteristics of the entry vehicle showing trimmed alpha, drag coefficient, and trimmed lift to drag ratio versus Mach number; the damping characteristics of the entry configuration; the angle of attack time history of Viking entries; stagnation heating and pressure time histories; and the aeroshell heating distribution as obtained in tests run in a shock tunnel for various gases. Flight tests which demonstrate the aerodynamic separation of the full-scale aeroshell and the flying qualities of the entry configuration in an uncontrolled mode are documented. Design values selected for the heat protection system based on the test data and analysis performed are presented.

Polutchko, R. J.↗

NASA NTRS1979

External tank laminar and rarefied aeroheating

A practical method has been developed for calculation of laminar and rarefied aerodynamic heating to the external tank of the Space Shuttle during the ascent phase of flight. A modification of laminar boundary theory is used as the basis for the heating relation. This relation is used in conjunction with a correlation for surface pressures to obtain undisturbed aerodynamic heating. Interference effects on the external tank due to the orbiter are correlated using the rarefaction parameter which was developed as part of the heating relation. The effects of rarefaction on the undisturbed and interference heating are illustrated with test data.

Engel, C. D.↗

NASA NTRSDate not supplied

Lessons Learned from Aerothermal Flight Data on NASA Mars Entry Vehicles

The NASA Mars Science Laboratory (MSL) and Mars 2020 entry vehicles included temperature measurements inside and direct heat flux on the surface of the thermal protection system (TPS) material. The instrumentation was included to begin addressing predictive capabilities for aerodynamic heating and ablative TPS material response. The MSL aeroshell included temperature measurements at seven heatshield locations, from which total aerodynamic heating was reconstructed. The Mars 2020 instrumentation suite included eleven heatshield measurement locations and new backshell measurements at nine locations: six with temperature sensors, two with sensors to directly measure total heat flux, and one to measure radiative heat flux. Both instrumentation suites returned full sets of data, with all temperature sensors nearest to the surface surviving entry, indicating minimal material recession. Heatshield boundary layer transition (BLT) was observed for both vehicles prior to the time of peak heating, with transition starting near the shoulder and progressing towards the stagnation area. MSL’s higher speed on approach to Mars resulted in higher total heating compared to Mars 2020, with the maximum occurring near the heatshield’s leeside shoulder. The Mars 2020 backshell heating data indicate significant contributions from shock layer radiation. Heating from Navier-Stokes flowfield calculations are in family with reconstructed heating. The paper covers the instrumentation, BLT and reconstructed heating on both heatshields, backshell heating for Mars 2020, TPS performance, and predicted versus reconstructed heat fluxes. Opportunities for improved design approaches for future Mars missions are discussed, especially for the upcoming Mars Sample Return mission.

Mars Science Laboratory↗

NASA NTRSDate not supplied

Lessons Learned from Aerothermal Flight Data on NASA Mars Entry Vehicles

Mars Science Laboratory↗

NASA NTRS1942

Aerodynamic Heat-Power Engine Operating on a Closed Cycle

Hot-air engines with dynamic compressors and turbines offer new prospects of success through utilization of units of high efficiencies and through the employment of modern materials of great strength at high temperature. Particular consideration is given to an aerodynamic prime mover operating on a closed circuit and heated externally. Increase of the pressure level of the circulating air permits a great increase of limit load of the unit. This also affords a possibility of regulation for which the internal efficiency of the unit changes but slightly. The effect of pressure and temperature losses is investigated. A general discussion is given of the experimental installation operating at the Escher Wyss plant in Zurich for a considerable time at high temperatures.

Ackeret, J.↗

NASA NTRS1993

Space Station trash removal system

A trash removal system for space stations is described. The system is comprised of a disposable trash bag member and an attached, compacted large, lightweight inflatable balloon element. When the trash bag member is filled, the astronaut places the bag member into space through an airlock. Once in the vacuum of space, the balloon element inflates. Due to the large cross-sectional area of the balloon element relative to its mass, the combined balloon element and the trash bag member are slowed by atmospheric drag to a much greater extent than the Space Station's. The balloon element and bag member lose altitude and re-enter the atmosphere, and the elements and contents are destroyed by aerodynamic heating. The novelty of this system is in the unique method of using the vacuum of space and aerodynamic heating to dispose of waste material with a minimum of increase in orbital debris.

Petro, Andrew J.↗

NASA NTRS1992

Heat transfer characteristics of hypersonic waveriders with an emphasis on leading edge effects

The present analysis of the heat-transfer characteristics of a family of viscous-optimized, 60 m-long waverider hypersonic vehicles gives attention to the transition from laminar to turbulent flow, and to how the transition affects aerodynamic heating distributions over the waverider surface. Two different constant-dynamic-pressure flight trajectories are considered, at 0.2 and 1.0 freestream atmospheres. For Mach numbers below 10, it is found that passive radiative cooling of the surface is sufficient. The degree of leading-edge bluntness required by aerodynamic heating constraints does not significantly degrade the aerodynamic performance of these waveriders.

Vanmol, Denis O.↗

NASA NTRS1971

Feasibility study for automatic reduction of phase change imagery

The feasibility of automatically reducing a form of pictorial aerodynamic heating data is discussed. The imagery, depicting the melting history of a thin coat of fusible temperature indicator painted on an aerodynamically heated model, was previously reduced by manual methods. Careful examination of various lighting theories and approaches led to an experimentally verified illumination concept capable of yielding high-quality imagery. Both digital and video image processing techniques were applied to reduction of the data, and it was demonstrated that either method can be used to develop superimposed contours. Mathematical techniques were developed to find the model-to-image and the inverse image-to-model transformation using six conjugate points, and methods were developed using these transformations to determine heating rates on the model surface. A video system was designed which is able to reduce the imagery rapidly, economically and accurately. Costs for this system were estimated. A study plan was outlined whereby the mathematical transformation techniques developed to produce model coordinate heating data could be applied to operational software, and methods were discussed and costs estimated for obtaining the digital information necessary for this software.

Nossaman, G. O.↗

NASA NTRS1990

Application of the LAURA code for slender-vehicle aerothermodynamics

Application of a three-dimensional thin-layer Navier-Stokes code (LAURA) to compute the laminar, perfect gas aerothermodynamics of slender vehicles is demonstrated. Comparison of solutions with experimental data enabled an assessment of the code's aerodynamic heating prediction capabilities. The results show radial grid refinement to a wall cell Reynolds number of 2 to be accurate for predicting aerodynamic heating to a slender cone. Additional comparisons for an 80 deg slab-delta wing and a generic Aero-Space Plane vehicle show the method to be generally accurate using this radial cell spacing. Discrepancies between the predicted and measured heating in a region of crossflow separation on the Aero-Space Plane geometry are noted which indicates the need for further study. While additional applications and comparisons are required to fully test the LAURA program, the results of this study demonstrate the current level of confidence and the engineering utility of this particular code.

Thompson, Richard A.↗

NASA NTRS1984

On the formation of meteoritic chondrules by aerodynamic drag heating in the solar nebula

During the formation of the solar nebula, interstellar grains were falling into the nebula with velocities of the order of 10 km/s at the radial distance where the meteorites were to form. This kinetic energy is 20 times the amount of thermal energy needed to melt the grains. The grains were decelerated by aerodynamic drag in the nebula. Where grain-rich parcels of interstellar material fell into the nebula, heat generated by drag could not be radiated away because of the opacity imparted to the system by the grains, and high temperatures were reached. In this situation presolar aggregations of grains would melt to form chondrules. Many of the properties of chondrules (and also Ca/Al-rich inclusions) are consistent with their formation by this means. The infall-heating concept provides a new framework in which the formation and significance of chondritic meteorites can be understood.

Wood, J. A.↗

NASA NTRS1990

Active control of aerothermoelastic effects for a conceptual hypersonic aircraft

Procedures for and results of aeroservothermoelastic studies are described. The objectives of these studies were to develop the necessary procedures for performing an aeroelastic analysis of an aerodynamically heated vehicle and to analyze a configuration in the classical cold state and in a hot state. Major tasks include the development of the structural and aerodynamic models, open loop analyses, design of active control laws for improving dynamic responses and analyses of the closed loop vehicles. The analyses performed focused on flutter speed calculations, short period eigenvalue trends and statistical analyses of the vehicle response to controls and turbulence. Improving the ride quality of the vehicle and raising the flutter boundary of the aerodynamically-heated vehicle up to that of the cold vehicle were the objectives of the control law design investigations.

Heeg, Jennifer↗

NASA NTRS1990

Active control of aerothermoelastic effects for a conceptual hypersonic aircraft

This paper describes the procedures for an results of aeroservothermoelastic studies. The objectives of these studies were to develop the necessary procedures for performing an aeroelastic analysis of an aerodynamically heated vehicle and to analyze a configuration in the classical 'cold' state and in a 'hot' state. Major tasks include the development of the structural and aerodynamic models, open loop analyses, design of active control laws for improving dynamic responses and analyses of the closed loop vehicles. The analyses performed focused on flutter speed calculations, short period eigenvalue trends and statistical analyses of the vehicle response to controls and turbulence. Improving the ride quality of the vehicle and raising the flutter boundary of the aerodynamically-heated vehicle up to that of the cold vehicle were the objectives of the control law design investigations.

Heeg, Jennifer↗

NASA NTRS1987

The effects of inlet turbulence and rotor/stator interactions on the aerodynamics and heat transfer of a large-scale rotating turbine model, volume 1

A combined experimental and analytical program was conducted to examine the effects of inlet turbulence on airfoil heat transfer. Heat transfer measurements were obtained using low conductivity airfoils with miniature thermocouples welded to a thin, electrically heated surface skin. Heat transfer data were acquired for various combinations of low or high inlet turbulence intensity, flow coefficient (incidence), first-stator/rotor axial spacing, Reynolds number, and relative circumferential position of the first and second stators. Aerodynamic measurements include distributions of the mean and fluctuating velocities at the turbine inlet and, for each airfoil row, midspan airfoil surface pressures and circumferential distributions of the downstream steady state pressures and fluctuating velocities. Analytical results include airfoil heat transfer predictions and a examination of solutions of the unstead boundary layer equipment.

Dring, R. P.↗

NASA NTRS1986

The effects of inlet turbulence and rotor/stator interactions on the aerodynamics and heat transfer of a large-scale rotating turbine model. Volume 2: Heat transfer data tabulation. 15 percent axial spacing

A combined experimental and analytical program was conducted to examine the effects of inlet turbulence on airfoil heat transfer. The experimental portion of the study was conducted in a large-scale (approx 5X engine), ambient temperature, rotating turbine model configured in both single stage and stage-and-a-half arrangements. Heat transfer measurements were obtained using low-conductivity airfoils with miniature thermcouples welded to a thin, electrically heated surface skin. Heat transfer data were acquired for various combinations of low or high inlet turbulence intensity, flow coefficient, first-stator/rotor axial spacing, Reynolds number and relative circumferential position of the first and second stators. Aerodynamic measurements obtained as part of the program include distributions of the mean and fluctuating velocities at the turbine inlet and, for each airfoil row, midspan airfoil surface pressures and circumferential distributions of the downstream steady state pressures and fluctuating velocities. Analytical results include airfoil heat transfer predictions produced using existing 2-D boundary layer computation schemes and an examination of solutions of the unsteady boundary layer equations. The results are reported in four separate volumes, of which this is Volume 2: Heat Transfer Data Tabulation; 15 Percent Axial Spacing.

Dring, R. P.↗

NASA NTRS1978

A test of the comet hypothesis of the Tunguska Meteor Fall - Nature of the meteor 'thermal' explosion paradox

The hypothesis that a comet was responsible for the Tunguska Meteor Fall is rejected because the hypothesis does not seem to account for the intense terminal spherical shock. A porous meteoroid model is proposed, and an analysis indicates that an entity of this type might produce an aerodynamic heat flux large enough to account for the terminal meteor explosion. It is suggested that the presence of olivine and of highly irregular macrostructure in meteors might indicate the presence of some porosity. For a highly porous meteoroid, it is postulated that during entry into the atmosphere the aerodynamic heat transfer at its external or pore walls would become so intensified as to cause either complete ablation with popping or a solid-liquid-vapor phase transition accompanied by an explosion.

Liu, V. C.↗

NASA NTRS1977

CAVE: A computer code for two-dimensional transient heating analysis of conceptual thermal protection systems for hypersonic vehicles

A digital computer code CAVE (Conduction Analysis Via Eigenvalues), which finds application in the analysis of two dimensional transient heating of hypersonic vehicles is described. The CAVE is written in FORTRAN 4 and is operational on both IBM 360-67 and CDC 6600 computers. The method of solution is a hybrid analytical numerical technique that is inherently stable permitting large time steps even with the best of conductors having the finest of mesh size. The aerodynamic heating boundary conditions are calculated by the code based on the input flight trajectory or can optionally be calculated external to the code and then entered as input data. The code computes the network conduction and convection links, as well as capacitance values, given basic geometrical and mesh sizes, for four generations (leading edges, cooled panels, X-24C structure and slabs). Input and output formats are presented and explained. Sample problems are included. A brief summary of the hybrid analytical-numerical technique, which utilizes eigenvalues (thermal frequencies) and eigenvectors (thermal mode vectors) is given along with aerodynamic heating equations that have been incorporated in the code and flow charts.

Rathjen, K. A.↗

NASA NTRS1976

Results of phase change paint tests of 0.040 scale 50% forebody models (82-0) of the space shuttle orbiter in the AEDC VKF B hypersonic wind tunnel (OH75)

Post-test information and data are presented from phase change paint, aerodynamic heating wind tunnel tests of a Rockwell International space shuttle orbiter forebody model. These tests were conducted in the Arnold Engineering and Development Center von Karman Facility Tunnel B Hypersonic Wind Tunnel. The purpose of these tests was to determine the effect of simulated orbiter protuberances and penetrations (including RCS nozzles) on aerodynamic heating rates during simulated entry conditions.

Dye, W. H.↗