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At least 271 records · Page 15

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NASA NTRS1983

Aerothermal modeling program, phase 1

The physical modeling embodied in the computational fluid dynamics codes is discussed. The objectives were to identify shortcomings in the models and to provide a program plan to improve the quantitative accuracy. The physical models studied were for: turbulent mass and momentum transport, heat release, liquid fuel spray, and gaseous radiation. The approach adopted was to test the models against appropriate benchmark-quality test cases from experiments in the literature for the constituent flows that together make up the combustor real flow.

Sturgess, G. J.↗

NASA NTRS1990

External tank aerothermal design criteria verification, volume 1

The objective of this study was to produce an independent set of ascent environments which would serve as a check on the Rockwell IVBC-3 environments and provide an independent reevaluation of the thermal design criteria for the External Tank (ET). Design heating rates and loads were calculated at 367 acreage body point locations. Ascent flight regimes covered were lift-off, first stage ascent, Solid Rocket Booster (SRB) staging and second stage ascent through ET separation. The purpose here is to document these results, briefly describe the methodology used and present the environments along with a comparison with the Rockwell IVBC-3 counterpart. The methodology and environment summaries are given.

Crain, William K.↗

NASA NTRS1983

Low L/D aerobrake test at Mach 10

An aerobraking Orbital transfer vehicle may be used to increase the Space Shuttle mission capability to and from high orbits. A Mach 10 wind-tunnel test was performed for a low lift-drag aerobrake, to define a preliminary aerothermal environment for this candidate concept. Test hardware simulated the ribs and stretched fabric of conceptual flight hardware. Pressures, paint-melting histories, and oilflow data were measured on the brake. Pressure and thermocouple heating rate data were measured on the payload. Brake peak heating is at the edge at all angles of attack, although the stagnation point is not outboard of 75 percent radius even at 20 degrees angle of attack. Brake ribs show slightly higher heating than flats, although pressures are essentially constant. Payload peak heating occurs near 12 degrees angle of attack, and is 30 percent of the sphere stagnation point heating (for a sphere of brake diameter). Payload pressure distributions follow the heating pattern. Reynolds number effects are small on the brake and large on the payload, for the range of test conditions: 0.4-1.0 million/foot.

Hair, L. M.↗

NASA NTRS1994

Thermoacoustic environments to simulate reentry conditions

Aerothermal environments as encountered during the reentry of spaceplanes or during the cruise of hypersonic aircrafts represent complex loading conditions for the external structures of those vehicles. In order to shield against the aerodynamic heating a special Thermal Protection System (TPS) is required which is designed as a light weight structure to reduce the weight penalty. TPS is therefore vulnerable to vibroacoustic fatigue caused by the pressure fluctuations of the environment. Because of the complex interactions between the loading forces and the resulting structural response which make an analytical treatment difficult and in order to provide means for fatigue testing IABG has designed and built a thermoacoustic facility which recently became operational. The facility is capable to produce surface temperatures up to 1.300 C at sound pressure levels up to 160 dB. This paper describes the design of the facility, some operational test work it also deals with problems associated with the facility instrumentation.

Bayerdoerfer, Gerhard↗

NASA NTRSDate not supplied

Aerosciences Implications for Uranus Aerocapture

Exploration of Uranus has been indicated as the highest priority new flagship planetary science missions for NASA in the 2022 Decadal Survey. Due to Uranus’ location as the second furthest planet from the Sun, interplanetary flight times of a traditional propulsive-capture orbiter can expect to take 13-15 years to reach Uranus while requiring up to two-thirds of the launch mass budget to be allocated to propellant alone. Aerocapture presents an alternative means to capture orbit around Uranus by using aerodynamic forces generated from flight through the planetary atmosphere to decelerate to orbital insertion while reducing time in interplanetary transit time and increasing on-orbit payload mass. To date, only one spacecraft (Voyager 2) has visited Uranus, and there have been no missions which have entered its atmosphere. While knowledge of planetary entry at rocky terrestrial planets with Air (N2/O2) and CO2 dominated atmospheres has greatly advanced over the past decades with a variety of missions to learn from, the H2/He composition and much larger atmosphere of ice giants such as Uranus poses new aerosciences implications which are not fully understood. These implications must be better understood not only for aerocapture, but for planetary entry at Uranus as a whole (such as a Uranus probe, as proposed by the 2022 Decadal Survey), as well as other ice giants like Neptune. To conduct a proper analysis of aerocapture methods, aerodynamic and aerothermal analysis of flight trajectories in Uranus atmosphere are required. Previous work has investigated aerocapture at Neptune (with similar atmosphere to Uranus) using a Mid-L/D configuration. Additional previous work considers the convective heating experienced in H2/He atmospheres. The current work looks to use heritage Low-L/D configurations, adding technical depth and maturation to analysis previously performed by Girija. Atmospheric flight at ice giants presents unique challenges, including but not limited to: fidelity of atmospheric models, large entry velocities due to location in the solar system, extended flight in rarefied flow conditions due to atmospheric scale, and unusual aerodynamic flow properties specific to Uranus and Ice Giant atmospheres.

Eli Shellabarger↗

NASA NTRS1982

Wind tunnel tests of space shuttle solid rocket booster insulation material in the aerothermal tunnel c

Wind tunnel tests of the space shuttle Solid Rocket Booster Insulation were conducted in the von Karman Gas Dynamics Facility Tunnel C. For these tests, Tunnel C was run at Mach 4 with a total temperature of 1100-1440 and a total pressure of 100 psia. Cold wall heating rates were changed by varying the test article support wedge angle. Selected results are presented to illustrate the test techniques and typical data obtained.

Hartman, A. S.↗

NASA NTRS1991

Finite element methods for integrated aerodynamic heating analysis

This report gives a description of the work which has been undertaken during the second year of a three year research program. The objectives of the program are to produce finite element based procedures for the solution of the large scale practical problems which are of interest to the Aerothermal Loads Branch (ALB) at NASA Langley Research Establishment. The problems of interest range from Euler simulations of full three dimensional vehicle configurations to local analyses of three dimensional viscous laminar flow. Adaptive meshes produced for both steady state and transient problems are to be considered. An important feature of the work is the provision of specialized techniques which can be used at ALB for the development of an integrated fluid/thermal/structural modeling capability.

Morgan, K.↗

NASA NTRS2001

Aero-thermal Calibration of the NASA Glenn Icing Research Tunnel (2000 Tests)

Aerothermal calibration measurements and flow quality surveys were made in the test section of the Icing Research Tunnel at the NASA Glenn Research Center. These surveys were made following major facility modifications including widening of the heat exchanger tunnel section, replacement of the heat exchanger, installation of new turning vanes, and installation of new fan exit guide vanes. Standard practice at NASA Glenn requires that test section calibration and flow quality surveys be performed following such major facility modifications. A single horizontally oriented rake was used to survey the flow field at several vertical positions within a single cross-sectional plane of the test section. These surveys provided a detailed mapping of the total and static pressure, total temperature, Mach number, velocity, flow angle and turbulence intensity. Data were acquired over the entire velocity and total temperature range of the facility. No icing conditions were tested; however, the effects of air sprayed through the water injecting spray bars were assessed. All data indicate good flow quality. Mach number standard deviations were less than 0.0017, flow angle standard deviations were between 0.3 deg and 0.8 deg, total temperature standard deviations were between 0.5 and 1.8 F for subfreezing conditions, axial turbulence intensities varied between 0.3 and 1.0 percent, and transverse turbulence intensities varied between 0.3 and 1.5 percent. Measurement uncertainties were also quantified.

Gonsalez, Jose C.↗

NASA NTRS2021

Overview of Dragonfly Entry Aerosciences Measurements (DrEAM)

NASA Ames Research Center (ARC) leads the Dragonfly Entry Aerosciences Measurements (DrEAM) project, which is an aeroshell instrumentation suite on the Dragonfly mission that fulfills the Engineering Science Investigation requirement for the New Frontiers mission. NASA ARC is partnering with NASA Langley Research Center (LaRC) and the German Aerospace Center (DLR) to provide a comprehensive sensor suite, including a DLR-provided Data Acquisition System (DAS). DrEAM will provide key aerothermodynamic data and performance analysis for Dragonfly’s forebody and backshell Thermal Protection System (TPS). Titan’s atmosphere predominantly consists of ni-trogen (~98% by mole) with small amounts of me-thane (~2% by mole) and other trace gases. CN is a strong radiator, and is found in nonequilibrium con-centrations for Titan entry. The accurate modeling of nonequilibrium CN radiation has proven to be a diffi-cult task. Prompted by the Huygens mission, many experimental campaigns and analyses were performed to better understand the aerothermal environments experienced by the probe during Titan entry [1]. How-ever, the Huygens probe carried no heatshield instru-mentation. Therefore, the DrEAM sensor suite will sig-nificantly advance the state-of-the-art not only by documenting the environment and performance of Dragonfly’s entry system but also by making key in situ measurements in Titan’s atmosphere for the first time. Aerothermal environments and TPS response will be measured using sensors whose flight heritage is tak-en from the Mars Entry, Descent, and Landing In-strumentation 2 (MEDLI2) thermocouple plugs and the COMbined Aerothermal and Radiometer Sensor (COMARS) suite [2], with the latter supplied by DLR. The MEDLI2 project used embedded thermocouples to directly measure the in-depth TPS temperature-time history at several locations on the heat shield and backshell of the Mars 2020 entry vehicle. These tem-perature measurements, in turn, can be used to infer surface environments via an inverse analysis proce-dure analogous to that used for MEDLI. For DrEAM, the thermocouple plug subsystem will be known as Dragonfly Sensors for Aero-Thermal Reconstruction (DragSTR). On Schiaparelli, the COMARS suite in-cluded three total surface-mounted heat flux sensors, three pressure sensors, and one radiometer. For DrEAM, the COMARS package will be known as the COmbined Sensor System for Titan Atmosphere (COSSTA). Since the methane concentration in the Titan atmos-phere is directly proportional to the radiative heat flux, the COSSTA measurements will be used to reduce the current uncertainty in the methane volume fraction. Atmospheric density measurements and capsule aero-dynamic data will be obtained through the onboard Inertial Measurement Unit (IMU), supplemented by pressure transducers similar to those used by the MEDLI and MEDLI2 projects. The DrEAM pressure sensors will be known as the Dragonfly Atmospheric Flight Transducers. (DrAFT). The pressure measure-ments, when combined with data from the on-board IMU, will allow for reconstruction of such quantities as vehicle Mach number, freestream density, and atmos-pheric winds. DrAFT measurements will enhance Dragonfly trajectory reconstruction and enable a sepa-ration of the aerodynamics from the atmosphere, as was done for MEDLI [3] and is currently in process for the MEDLI2 flight data set.

EDL↗

NASA NTRS1974

Techniques for aerothermal tests of large, flightweight thermal protection panels in a Mach 7 wind tunnel

Recently developed experimental techniques permit evaluating thermal performance and structural integrity of full-scale panel concepts applicable to reentry and hypersonic vehicles in the Langley 8-foot high-performance structures tunnel. Rate and magnitude of surface heating and differential-pressure loading are independently controlled. Realistic temperature distributions are radiantly preheated into the panel prior to aerodynamic heating, and stream conditions are preselected to sustain the preheat surface heating input during aerodynamic exposure. During tunnel start and shutdown, panels are shielded outside the stream from potentially damaging transient acoustics and buffeting and are then rapidly inserted into the hypersonic flow. Infrared radiometry provides detailed surveys of surface temperatures.

Deveikis, W. D.↗

NASA NTRS1984

Aerothermal loads analysis for high speed flow over a quilted surface configuration

Attention is given to hypersonic laminar flow over a quilted surface configuration that simulates an array of Space Shuttle Thermal Protection System panels bowed in a spherical shape as a result of thermal gradient through the panel thickness. Pressure and heating loads to the surface are determined. The flow field over the configuration was mathematically modeled by means of time-dependent, three-dimensional conservation of mass, momentum, and energy equations. A boundary mapping technique was then used to obtain a rectangular, parallel piped computational domain, and an explicit MacCormack (1972) explicit time-split predictor corrector finite difference algorithm was used to obtain steady state solutions. Total integrated heating loads vary linearly with bowed height when this value does not exceed the local boundary layer thickness.

Olsen, G. C.↗

NASA NTRS1984

Aerothermal loads analysis for high speed flow over a quilted surface configuration

Attention is given to hypersonic laminar flow over a quilted surface configuration that simulates an array of Space Shuttle Thermal Protection System panels bowed in a spherical shape as a result of thermal gradients through the panel thickness. Pressure and heating loads to the surface are determined. The flow field over the configuration was mathematically modeled by means of time-dependent, three-dimensional conservation of mass, momentum, and energy equations. A boundary mapping technique was then used to obtain a rectangular, parallelepiped computational domain, and an explicit MacCormack (1972) explicit time-split predictor-corrector finite difference algorithm was used to obtain steady state solutions. Total integrated heating loads vary linearly with bowed height when this value does not exceed the local boundary layer thickness.

Olsen, G. C.↗

NASA NTRS2015

Development of an Ablative 3D Quartz / Cyanate Ester Composite for the Orion Spacecraft Compression Pad

With the successful flight test of the Orion Multipurpose Crew Vehicle last December, NASA is well on its way to sending humans into deep space with the goal of putting astronauts on Mars in the 2030s. Orion will receive some upgrades for its next launch in 2018 including a newly developed 3D Quartz Cyanate Ester composite material for the compression pad. The compression pads are 10-inches in diameter and 3-inches thick. The pucks are at 4 locations in the heat shield that serve as a part of the mechanism for holding the Crew and Service modules together during most mission phases prior to separation followed by Earth re-entry. Thus the compression pad has structural, aerothermal and thermal protection requirements.This paper describes the approach used for developing the new 3D composite, including continuous 3D weaving on an automated loom followed by resin transfer molding. Mechanical, thermal and arc jet testing of the 3D composite are also described.

Feldman, Jay↗

NASA NTRSDate not supplied

Study of Graphite Ablation at Arc-Jet Conditions using Finite-Rate and Equilibrium Chemistry Models

Moving at hypersonic speeds through the atmosphere generates immense amount of heat that triggers thermo-chemical non-equilibrium effects in the flow approaching the vehicle and leads to surface chemical reactions that cause ablation and consume the material mass. This complex behavior and interaction of the aerothermal environment with a graphite surface is com- putationally modeled in the present work using a fully-coupled simulation between flow solver CHAMPS NBS-Cart and material response solver KATS-MR. In this study, we account for the thermo-chemical non-equilibrium effects occurring in the flow and use an 11-species air-carbon re- action model to simulate the ablation of the material. The reactions at the surface such oxidation, nitridation and recombination are modeled with a newly developed air-carbon reaction model. The coupled simulation is applied to study two experimental cases performed at IHF arc-jet facility at NASA Ames research center. The obtained results are compared against the measured data of surface recession, as well as surface and in-depth temperatures. Additionally, uncoupled sim- ulations, performed with equilibrium chemistry at the wall, point out the substantial differences between the coupled and uncoupled solutions, leading to higher deviation of the predicted temper- atures. Overall, we show excellent agreement between the coupled simulation predictions and the experimental data in all measured quantities.

EDL↗

NASA NTRS1986

Opportunities and challenges in heat transfer - From the perspective of the government laboratory

The technological problems in which heat transfer research is required are briefly reviewed, discussing the role of the government laboratory in this research. The Hot Section Technology Program at NASA Lewis is addressed as an example. A building-block approach to turbine aerothermal research is shown, as is an interactive computational/experimental synergism and an interdisciplinary research methodology for life prediction of turbine engine hot-section components. The main issues for government laboratories as organizers and focusers in this area of research are listed, and challenges and opportunities for the future are indicated.

Simoneau, Robert J.↗

NASA NTRS1986

Aerothermal tests of spherical dome protuberances on a flat plate at a Mach number of 6.5

Aerothermal tests were conducted in the Langley 8-Foot High-Temperature Tunnel at a Mach number of 6.5 on a series of spherical dome protuberances mounted on a flat-plate test apparatus. Detailed surface pressure and heating-rate distributions were obtained for various dome heights and diameters submerged in both laminar and turbulent boundary layers including a baseline geometric condition representing a thermally bowed metallic thermal protection system (TPS) tile. The present results indicated that the surface pressures on the domes were increased on the windward surface and reduced on the leeward surface as predicted by linearized small-perturbation theory, and the distributions were only moderately affected by boundary-layer variations. Surface heating rates for turbulent flow increased on the windward surface and decreased on the leeward surface similar to the pressure; but for laminar boundary layers, the heating rates remained high on the leeward surface, probably due to local transition. Transitional flow effects cause heat load augmentation to increase by 30 percent for the maximum dome height in a laminar boundary layer. However, the corresponding augmentation for a dome with a height of 0.1 in. and a diameter of 14 in. representative of a bowed TPS tile was 14 percent or less for either a laminar or turbulent boundary layer.

Glass, C. E.↗

NASA NTRS1992

An engineering aerodynamic heating method for hypersonic flow

A capability to calculate surface heating rates has been incorporated in an approximate three-dimensional inviscid technique. Surface streamlines are calculated from the inviscid solution, and the axisymmetric analog is then used along with a set of approximate convective-heating equations to compute the surface heat transfer. The method is applied to blunted axisymmetric and three-dimensional ellipsoidal cones at angle of attack for the laminar flow of a perfect gas. The method is also applicable to turbulent and equilibrium-air conditions. The present technique predicts surface heating rates that compare favorably with experimental (ground-test and flight) data and numerical solutions of the Navier-Stokes (NS) and viscous shock-layer (VSL) equations. The new technique represents a significant improvement over current engineering aerothermal methods with only a modest increase in computational effort.

Riley, Christopher J.↗

NASA NTRS1998

Computational Aeroheating Predictions for X-34

Radiative equilibrium surface temperatures, heating rates, streamlines, surface pressures, and flow-field features as predicted by the Langley Aerothermodynamic Upwind Relaxation Algorithm (Laura) are presented for the X-34 Technology Demonstrator. Results for two trajectory points corresponding to entry peak heating and two control surface deflections are discussed. This data is also discussed in context of Thermal Protection System (TPS) design issues. The work presented in this report is part of a larger effort to define the X-34 aerothermal environment, including the application of engineering codes and wind-tunnel studies.

Kelb, William L.↗