Search NASA⌕ Search

SEARCH · Search NASA

Results for “HYPERSONIC BOUNDARY LAYER”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Numerical Simulations of A Conceptual MSR-EES Shoulder Recession

The present study demonstrates our in-house material response solver, Icarus's capability to simulate the ablative conditions, including pyrolysis effects due to the interaction between hypersonic boundary layers and the thermal protection system (TPS). A conceptual aeroshell shoulder design, which undergoes mission-relevant flow and material conditions, is selected for the demonstration purpose. LAURA, a structured flow solver, is used to solve flow around the shoulder at several trajectory points of a flight path. The aerothermal dataset obtained from LAURA is used to enforce boundary conditions on the aeroshell wall to simulate ablation processes. The two-layered material system is stacked with HEEET (TPS) and Aluminum (actual material). A set of base parameters, such as the angle of material orientation with respect to the flow direction, convective heat-transfer co-efficient, and aerothermal boundary conditions on the rear end of the HEEET material layer, is selected to conduct a base-case simulation. A small amount of recession was observed, indicating that the design should be fine to go through the selected flow and material conditions. Furthermore, the parameters mentioned above are individually altered and are observed to affect the shoulder design's recession compared to the base case. Verification of our Icarus setup was also carried out using a one-dimensional grid and FIAT, a one-dimensional material-response solver, to build credibility for our results. As the current work is an uncoupled fluid-material-response simulation, a local sharp mesh deformation in the recessed surface near the shoulder corner is observed.

Prakash Shrestha↗
NASA NTRS1976

Real gas scale effects on hypersonic laminar boundary-layer parameters including effects of entropy-layer swallowing

Inviscid and viscous (laminar boundary-layer) flow-field calculations under perfect gas hypersonic wind tunnel and equilibrium real gas flight conditions are presented for the windward centerline of the Rockwell International 139 Space Shuttle Orbiter at a 30-deg angle of attack. Correlation parameters for laminar boundary-layer edge quantities and surface heat transfer are developed which properly account for entropy-layer-swallowing effects under both real and perfect gas conditions. Some implications of the proposed correlation parameters on boundary-layer transition are discussed.

Adams, J. C., Jr.↗
NASA NTRS2019

Effect of Mass Transfer on Aeroheating in Hypersonic Chemically Reacting Boundary Layers

Heat flux characterization of high-enthalpy boundary layer flows is key to optimize the performance and design of Thermal Protection System of next generation aerospace vehicles [1]. At atmospheric entry hypersonic speeds, ablation as well as surface catalycity impact boundary layer aeroheating. Out-gassing occurring from an ablative surface in planetary entry environment introduces a rich set of problems in thermodynamic, fluid dynamic, and material pyrolysis. Ablation leads to out-gassing and surface roughness, both of which are known to affect surface heating in hypersonic chemically reacting boundary layers via three main routes: gas blowing into the boundary layer from the wall, changing the surface heat transfer due to wall-flow chemical reactions, and modifying surface roughness via ablative processes.

Karimi, Mona↗
NASA NTRS1988

Hypersonic turbulent wall boundary layer computations

The Baldwin-Lomax algebraic turbulence model was modified for hypersonic flow conditions. Two coefficients in the outer layer eddy viscosity model were determined as functions of Mach number and temperature ratio. By matching the solutions from the Baldwin-Lomax model to those from the Cebeci-Smith model for a flat plate at hypersonic speed, the new values of the coefficient were obtained. The results show that the values of C sub cp and C sub kleb are functions of both Mach number and wall temperature ratio. The C sub cp and C sub kleb variations with Mach number and wall temperature were used for the calculations of both a 4 deg wedge flow at Mach 18 and an axisymmetric Mach 20 nozzle flow. The Navier-Stokes equations with thin layer approximation were solved for the above hypersonic flow conditions and the results were compared with existing experimental data. The agreement between the numerical solutions and the existing experimental data were good. The modified Baldwin-Lomax model thus is useful in the computations of hypersonic flows.

Kim, S. C.↗
NASA NTRS1975

Shock-wave-induced turbulent boundary-layer separation at hypersonic speeds

An experiment is described that tests and guides computations of the interaction of a shock wave with a turbulent boundary layer. Numerical solutions of the time-averaged Navier-Stokes equations for the entire flow field employing algebraic eddy viscosity and turbulent Prandtl number models for shear stress and heat flux are presented and used to illustrate the dependence of the computations on the particulars of the turbulence models. To guide modifications in the models, the mean flow profiles and surface measurements of pressure, shear, and heat flux are analyzed critically. The results show that the models of eddy viscosity require substantial modifications in the interaction region. Improved solutions employing the experimentally modified models are presented.

Horstman, C. C.↗
NASA NTRS1975

An experimental and numerical investigation of shock-wave induced turbulent boundary-layer separation at hypersonic speeds

An experiment designed to test and guide computations of the interaction of an impinging shock wave with a turbulent boundary layer is described. Detailed mean flow-field and surface data are presented for two shock strengths which resulted in attached and separated flows, respectively. Numerical computations, employing the complete time-averaged Navier-Stokes equations along with algebraic eddy-viscosity and turbulent Prandtl number models to describe shear stress and heat flux, are used to illustrate the dependence of the computations on the particulars of the turbulence models. Models appropriate for zero-pressure-gradient flows predicted the overall features of the flow fields, but were deficient in predicting many of the details of the interaction regions. Improvements to the turbulence model parameters were sought through a combination of detailed data analysis and computer simulations which tested the sensitivity of the solutions to model parameter changes. Computer simulations using these improvements are presented and discussed.

Marvin, J. G.↗
NASA NTRS1988

Hypersonic turbulent wall boundary layer computations

The Baldwin-Lomax (1978) algebraic turbulence model was modified for hypersonic flow conditions. Two coefficients in the outer-layer eddy-viscosity model were determined as functions of Mach number and temperature ratio. By matching the solutions from the Baldwin-Lomax model to those from the Cebeci-Smith (1974) model for a flat plate at hypersonic speed, the new values of the coefficients were obtained. The results show that the values of C(cp) and C(kleb) are functions of both Mach number and wall temperature ratio. The C(cp) and C(kleb) variations with Mach number and wall temperature were used for the calculations of both a 4-deg wedge flow at Mach 18 and an axisymmetric Mach 20 nozzle flow. The Navier-Stokes equations with thin-layer approximation were solved for the above hypersonic flow conditions and the results were compared with existing experimental data. The agreement between the numerical solutions and the existing experimental data were good. The modified Baldwin-Lomax model thus is useful in the computations of hypersonic flows.

Kim, S. C.↗
NASA NTRS1990

Boundary layer transition in hypersonic flows

This paper summarizes some of the recent progress made at NASA Langley Research Center in the understanding, prediction and modeling of high speed boundary-layer transition. Linear and nonlinear theories together with large-eddy and direct numerical simulations have been used to understand various aspects of the transition problem while low disturbance 'quiet' tunnels provide means for validating the theoretical results.

Malik, M. R.↗
NASA NTRS1975

An experimental and numerical investigation of shock wave induced turbulent boundary layer separation at hypersonic speeds

A thoroughly documented experiment is reported that was specifically designed to test and guide computations of the interaction of an impinging shock wave with a turbulent boundary layer. Detailed mean flow field and surface data are presented for two shock strengths which resulted in attached and separated flows, respectively. Numerical computations are used to illustrate the dependence of the computations on the particulars of the turbulence models. Models appropriate for zero pressure gradient flows predicted the overall features of the flow fields, but were deficient in predicting many of the details of the interaction regions. Improvements to the turbulence model parameters were sought through a combination of detailed data analysis and computer simulations which tested the sensitivity of the solutions to model parameter changes. Computer simulations using these improvements are presented and discussed.

Marvin, J. G.↗
NASA NTRS1990

On the instability of hypersonic flow past a flat plate

The instability of hypersonic boundary-layer flows over flat plates is considered. The viscosity of the fluid is taken to be governed by Sutherland's law, which gives a much more accurate representation of the temperature dependence of fluid viscosity at hypersonic speeds than Chapman's approximate linear law; although at lower speeds the temperature variation of the mean state is less pronounced so that the Chapman law can be used with some confidence. Attention is focussed on the so-called (vorticity) mode of instability of the viscous hypersonic boundary layer. This is thought to be the fastest growing inviscid disturbance at hypersonic speeds; it is also believed to have an asymptotically larger growth rate than any viscous or centrifugal instability. As a starting point the instability of the hypersonic boundary layer which exists far downstream from the leading edge of the plate is investigated. In this regime the shock that is attached to the leading edge of the plate plays no role, so that the basic boundary layer is non-interactive. It is shown that the vorticity mode of instability of this flow operates on a significantly different lengthscale than that obtained if a Chapman viscosity law is assumed. In particular, it is found that the growth rate predicted by a linear viscosity law overestimates the size of the growth rate by O(M(exp 2). Next, the development of the vorticity mode as the wavenumber decreases is described, and it is shown that acoustic modes emerge when the wavenumber has decreased from it's O(1) initial value to O(M (exp -3/2). Finally, the inviscid instability of the boundary layer near the leading edge in the interaction zone is discussed and particular attention is focussed on the strong interaction region which occurs sufficiently close to the leading edge. It is found that the vorticity mode in this regime is again unstable, and that it is concentrated in the transition layer at the edge of the boundary layer where the temperature adjusts from its large, O(M(exp 2), value in the viscous boundary layer, to its O(1) free stream value. The existence of the shock indirectly, but significantly, influences the instability problem by modifying the basic flow structure in this layer.

Blackaby, Nicholas↗
NASA NTRS1993

On the instability of hypersonic flow past a flat plate

Qualitative features of the inviscid instability characteristics of hypersonic boundary-layer flows over a flat plate are considered. The instability of a viscous hypersonic boundary layer which exists far downstream from the leading edge of the plate. It is shown that the vorticity mode of instability operates on a different lengthscale from that obtained using a Chapman viscosity law. The growth rate predicted by a linear viscosity law is found to overestimate the size of the growth rate. The inviscid instability of the boundary layer near the leading edge interaction zone is discussed focusing on the strong-interaction zone which occurs sufficiently close to the leading edge. The vorticity mode in this regime is found to be unstable.

Blackaby, Nicholas D.↗
NASA NTRS2001

Infrared Sensing Aeroheating Flight Experiment: STS-96 Flight Results

Major elements of an experiment called the Infrared Sensing Aeroheating Flight Experiment are discussed. The primary experiment goal is to provide reentry global temperature images from infrared measurements to define the characteristics of hypersonic boundary-layer transition during flight. Specifically, the experiment is to identify, monitor, and quantity hypersonic boundary layer windward surface transition of the X-33 vehicle during flight. In addition, the flight data will serve as a calibration and validation of current boundary layer transition prediction techniques, provide benchmark laminar, transitional, and fully turbulent global aeroheating data in order to validate existing wind tunnel and computational results, and to advance aeroheating technology. Shuttle Orbiter data from STS-96 used to validate the data acquisition and data reduction to global temperatures, in order to mitigate the experiment risks prior to the maiden flight of the X-33, is discussed. STS-96 reentry midwave (3-5 micron) infrared data were collected at the Ballistic Missile Defense Organization/Innovative Sciences and Technology Experimentation Facility site at NASA-Kennedy Space Center and subsequently mapped into global temperature contours using ground calibrations only. A series of image mapping techniques have been developed in order to compare each frame of infrared data with thermocouple data collected during the flight. Comparisons of the ground calibrated global temperature images with the corresponding thermocouple data are discussed. The differences are shown to be generally less than about 5%, which is comparable to the expected accuracy of both types of aeroheating measurements.

Blanchard, Robert C.↗
NASA NTRS2001

Infrared Sensing Aeroheating Flight Experiement: STS-96 Flight Results

Major elements of an experiment called the Infrared Sensing Aeroheating Flight Experiment are discussed. The primary experiment goal is to provide reentry global temperature images from infrared measurements to define the characteristics of hypersonic boundary-layer transition during flight. Specifically, the experiment is to identify, monitor, and quantify hypersonic boundary layer windward surface transition of the X-33 vehicle during flight. In addition, the flight data will serve as a calibration and validation of current boundary layer transition prediction techniques, provide benchmark laminar, transitional, and fully turbulent global aeroheating data in order to validate existing wind tunnel and computational results, and to advance aeroheating technology. Shuttle Orbiter data from STS-96 used to validate the data acquisition and data reduction to global temperatures, in order to mitigate the experiment risks prior to the maiden flight of the X-33, is discussed. STS-96 reentry mid-wave (3-5 Pm) infrared data were collected at the Ballistic Missile Defense Organization/Innovative Sciences and Technology Experimentation Facility site at NASA-Kennedy Space Center and subsequently mapped into global temperature contours using ground calibrations only. A series of image mapping techniques have been developed in order to compare each frame of infrared data with thermocouple data collected during the flight. Comparisons of the ground calibrated global temperature images with the corresponding thermocouple data are discussed. The differences are shown to be generally less than about 5%, which is comparable to the expected accuracy of both types of aeroheating measurements.

Blanchard, Robert C.↗