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NASA NTRS2015

Boundary Layer Transition and Trip Effectiveness on an Apollo Capsule in the JAXA High Enthalpy Shock Tunnel (HIEST) Facility

Computational assessments were performed to size boundary layer trips for a scaled Apollo capsule model in the High Enthalpy Shock Tunnel (HIEST) facility at the JAXA Kakuda Space Center in Japan. For stagnation conditions between 2 MJ/kg and 20 MJ/kg and between 10 MPa and 60 MPa, the appropriate trips were determined to be between 0.2 mm and 1.3 mm high, which provided kappa/delta values on the heatshield from 0.15 to 2.25. The tripped configuration consisted of an insert with a series of diamond shaped trips along the heatshield downstream of the stagnation point. Surface heat flux measurements were obtained on a capsule with a 250 mm diameter, 6.4% scale model, and pressure measurements were taken at axial stations along the nozzle walls. At low enthalpy conditions, the computational predictions agree favorably to the test data along the heatshield centerline. However, agreement becomes less favorable as the enthalpy increases conditions. The measured surface heat flux on the heatshield from the HIEST facility was under-predicted by the computations in these cases. Both smooth and tripped configurations were tested for comparison, and a post-test computational analysis showed that kappa/delta values based on the as-measured stagnation conditions ranged between 0.5 and 1.2. Tripped configurations for both 0.6 mm and 0.8 mm trip heights were able to effectively trip the flow to fully turbulent for a range of freestream conditions.

Kirk, Lindsay C.↗

NASA NTRS1995

Effects of Chemistry on Blunt-Body Wake Structure

Results of a numerical study are presented for hypersonic low-density flow about a 70-deg blunt cone using direct simulation Monte Carlo (DSMC) and Navier-Stokes calculations. Particular emphasis is given to the effects of chemistry on the near-wake structure and on the surface quantities and the comparison of the DSMC results with the Navier-Stokes calculations. The flow conditions simulated are those experienced by a space vehicle at an altitude of 85 km and a velocity of 7 km/s during Earth entry. A steady vortex forms in the near wake for these freestream conditions for both chemically reactive and nonreactive air gas models. The size (axial length) of the vortex for the reactive air calculations is 25% larger than that of the nonreactive air calculations. The forebody surface quantities are less sensitive to the chemistry than the base surface quantities. The presence of the afterbody has no effect on the forebody flow structure or the surface quantities. The comparisons of DSMC and Navier-Stokes calculations show good agreement for the wake structure and the forebody surface quantities.

Dogra, Virendra K.↗

NASA NTRS2021

Boundary-layer Instability Measurements on a Cone at Freestream Mach 3.5

An experimental study was conducted in the NASA (National Aeronautics and Space Administration) Langley Supersonic Low-Disturbance Tunnel to investigate naturally-occurring instabilities in a supersonic boundary layer on a 7◦ half-angle cone at nominal freestream conditions: Mach 3.5, total temperature of 299.8K, and unit Reynolds numbers (millions per m) of 9.89, 13.85, 21.77, and 25.73. Instability measurements were acquired under noisy-flow and quiet-flow conditions. Pitot-pressure and calibrated hot-wire measurements were obtained using a model-integrated traverse system to document the model flow field. In noisy-flow conditions, growth rates and mode shapes achieved good agreement between the measured results and linear stability theory (LST). The corresponding N factor at transition from LST is N ≈ 3.9. Under quiet-flow conditions, the most unstable first-mode instabilities as predicted by LST were measured, but this mode was not the dominant instability measured in the boundary layer. Instead, the dominant instabilities were less-amplified, low-frequency disturbances predicted by LST, and grew according to linear theory. These low-frequency unstable disturbances were initiated by freestream acoustic disturbances through a receptivity process believed to occur near the branch I location of the cone. Under quiet-flow conditions, the boundary layer remained laminar up to the last measurement station for the largest unit Reynolds number, implying a transition N factor of N > 8.5.

Boundary Layer↗

NASA NTRS1975

Numerical computation of three-dimensional blunt body flow fields with an impinging shock

A time-marching finite-difference method was used to solve the compressible Navier-Stokes equations for the three-dimensional wing-leading-edge shock impingement problem. The bow shock was treated as a discontinuity across which the exact shock jump conditions were applied. All interior shock layer detail such as shear layers, shock waves, jets, and the wall boundary layer were automatically captured in the solution. The impinging shock was introduced by discontinuously changing the freestream conditions across the intersection line at the bow shock. A special storage-saving procedure for sweeping through the finite-difference mesh was developed which reduces the required amount of computer storage by at least a factor of two without sacrificing the execution time. Numerical results are presented for infinite cylinder blunt body cases as well as the three-dimensional shock impingement case. The numerical results are compared with existing experimental and theoretical results.

Holst, T. L.↗

NASA NTRSDate not supplied

Application of a Detached Eddy Simulation Approach with Finite-Rate Chemistry to Mars-Relevant Retropropulsion Operating Environments

Human-scale Mars vehicles will require retropropulsion for descent and landing, replacing heritage supersonic parachute systems with an extended phase of powered flight. Due to the limitations of terrestrial testing in Mars-relevant conditions, design and analysis will increasingly rely on computational modeling and simulation. This paper provides an overview of a computational campaign investigating the aerodynamics of a Mars lander concept along various points on a powered descent trajectory including supersonic, transonic, and subsonic conditions using finite-rate chemistry. Simulations using unstructured grids containing billions of elements are performed at scale using thousands of Graphics Processing Units, enabling run-times of a few days for each simulation presented. At each freestream condition, significant minor species concentrations are observed external to the nozzles in the large mixing region upstream of the vehicle. While the flowfields are highly non-stationary in all cases, the mean integrated forces and moments on the vehicle remain small in comparison to the deceleration provided through retropropulsion.

Ashley Korzun↗

NASA NTRS2019

Measurement and Analysis of Terminal Shock Oscillation and Buffet Forcing Functions on a Launch Vehicle Payload Fairing

The buffet loads on a launch vehicle payload shroud can be impacted by the unsteadiness associated with a terminal shock at high subsonic speeds. At these conditions, flow accelerates to supersonic speeds on the nose of the payload fairing and is terminated by a normal shock on the cylindrical section downstream of the nose cone/cylinder shoulder. The location of the terminal shock and associated separated boundary layer is affected by the freestream Mach number, Reynolds number, and the pitch/yaw of the launch vehicle. Furthermore, even when the freestream conditions and vehicle attitude are constant, this terminal shock oscillates on the surface of the vehicle. The time-varying surface pressure associated with the terminal shock results in unsteady aerodynamic loads that may interact with vehicle structural dynamic modes and the guidance and control of the vehicle. Buffet testing of a 3-percent scale rigid buffet model of a launch vehicle cargo configuration with a tangent-ogive payload shroud was conducted in 2012 and in 2016. Initial buffet forcing functions (BFFs) utilized a coarse pressure sensor distribution on the vehicle surface in which a single longitudinal station with eight sensors observed the terminal shock environment at Mach 0.90. An examination of these circumferential pressures reveal large impulse-like pressure fluctuations and an asymmetry in pressure when the vehicle is at a nonzeroangle of attack that result in high BFFs. Revisions to the shock integration region were made based on computational fluid dynamics and shadowgraph video of shock motion to better represent the BFFs and reduce the high loads resulting from this environment. To more clearly understand this terminal shock environment, a second wind tunnel test was conducted with a dense distribution of 256 sensors at the terminal shock location. These sensor arrays presents a unique opportunity to observe the unsteady terminal shock environment and to characterize the impact of various integration schemes on the BFFs. This paper presents a summary of the development of BFFs for this terminal shock and a detailed analyses of shock region pressure coefficients, coherence, BFFs, shock location time histories, and power spectral density to help guide development of BFFs for other launch vehicle test and analysis programs.

Piatak, David J.↗

NASA NTRS1995

Arcjets for Aerodynamic and Materials Testing: Flow Characterization

Potential use of Arcjets as a hypersonic ground test facility for aerodynamic testing of future space vehicles is examined. Since high fidelity simulation of flight freestream conditions is a basic requirement for any useful ground test facility, it is imperative that the Arcjet flow field be thoroughly investigated in a systematic and orderly manner. At the same time we must know how and to what extent an inaccurate simulation of the flight freestream will effect the test data. The paper after discussing these two topics, describes various experimental techniques for Arcjet flow characterization. Results from an on-going Arcjet flow characterization program are also presented.

Sharma, Surendra P.↗

NASA NTRS1979

Significance of shock and body slip conditions on Jovian entry heating

The influence of the body and shock slip conditions on the heating of a Jovian entry body is investigated. The flow in the shock layer is considered to be axisymmetric, steady, laminar, viscous, and in chemical equilibrium. Realistic thermophysical and step-function spectral models are employed and results are obtained by implicit finite-difference and iteractive procedures. The freestream conditions correspond to a typical Jovian entry trajectory point. The results indicate that the effect of the slip conditions is significant when the altitudes are higher than 225 km and that the contribution of a radiative heat-flux term in the energy equation should not be neglected at any altitude.

Tiwari, S. N.↗

NASA NTRS1996

Experimental and Computational Aerothermodynamics of a Mars Entry Vehicle

An aerothermodynamic database has been generated through both experimental testing and computational fluid dynamics simulations for a 70 deg sphere-cone configuration based on the NASA Mars Pathfinder entry vehicle. The aerothermodynamics of several related parametric configurations were also investigated. Experimental heat-transfer data were obtained at hypersonic test conditions in both a perfect gas air wind tunnel and in a hypervelocity, high-enthalpy expansion tube in which both air and carbon dioxide were employed as test gases. In these facilities, measurements were made with thin-film temperature-resistance gages on both the entry vehicle models and on the support stings of the models. Computational results for freestream conditions equivalent to those of the test facilities were generated using an axisymmetric/2D laminar Navier-Stokes solver with both perfect-gas and nonequilibrium thermochemical models. Forebody computational and experimental heating distributions agreed to within the experimental uncertainty for both the perfect-gas and high-enthalpy test conditions. In the wake, quantitative differences between experimental and computational heating distributions for the perfect-gas conditions indicated transition of the free shear layer near the reattachment point on the sting. For the high enthalpy cases, agreement to within, or slightly greater than, the experimental uncertainty was achieved in the wake except within the recirculation region, where further grid resolution appeared to be required. Comparisons between the perfect-gas and high-enthalpy results indicated that the wake remained laminar at the high-enthalpy test conditions, for which the Reynolds number was significantly lower than that of the perfect-gas conditions.

Hollis, Brian R.↗

NASA NTRSDate not supplied

Fluidic Fence Flow Control on A 30° Swept Wing in Compressible Freestreams

A 30° swept wing with a NACA 64(3)-618 airfoil cross-section was evaluated in four different freestream conditions with Mach numbers increasing from 0.05 to 0.7 using computational fluid dynamics simulations. Active flow control (AFC) through a streamwise row of vortex generating jets was applied at 70% span, blowing inboard to obstruct the development of spanwise flow and increase lift outboard of the control location. The effectiveness of the AFC was evaluated at α = 5° and 10° as freestream Mach number increased and the AFC mass flow coefficient was held constant. Increases to lift coefficient were reduced at higher Mach numbers, but the ratio of change in lift coefficient and the AFC momentum coefficient remained roughly constant. The fluidic fence caused decreased pressure coefficients on the outboard wing sections, but this region of decreased pressure coefficient moved aft on the chord with increasing Mach number. Mach 0.7 is higher than the critical Mach number of the wing and produced interesting interplay with the AFC, as the suction surface shock moved forward on the inboard wing section but moved aft on the outboard wing section.

Evan J McFadden↗

NASA NTRSDate not supplied

Fluidic Fence Flow Control on A 30° Swept Wing in Compressible Freestreams

Evan J McFadden↗

NASA NTRS1996

CFD Analysis of Modular Thrusters Performance

The effective performance of modular thrusters in an aerospike configuration is difficult to determine. Standard analytical tools are applicable to conventional nozzle shapes, but are limited when applied to an aerospike nozzle (An aerospike nozzle is an altitude compensating external nozzle). Three baseline nozzle shapes are derived using standard analytical procedures. The baseline nozzle sizes are restricted to fill a volume envelope. The three shapes are an axi-symmetric round nozzle, a two dimensional planar square exit nozzle, and a super elliptic round to nearly square nozzle. The integrated (thruster/aerospike) performance of the three nozzles is determined through the use of three dimensional viscous computational fluid dynamic (CFD) calculation where complex features of the flow field can be accurately captured. The resulting installed performance is then used to evaluate the efficiency of these nozzle shapes for aerospike applications. The determination of effective performance of a thruster nozzle integrated into an aerospike nozzle require the solution of the three dimensional turbulent Navier-Stokes equations. The model used in this study consisted of two zones; one of the upstream thruster cowl surface so freestream conditions can be accurately predicted, and two, the aerospike surface beginning with the thruster outflow and extending to the end of the aerospike surface. The numerical grid consisted of over 120,000 nodes and used symmetry on the thruster centerline and edge. A two species non-reacting chemistry model was used to capture the variation of fluid properties between the hot plume base and freestream air. From the results of the three baseline nozzle aerospike calculations, the effictive performance of the nozzle was determined. The flow fields of these calculations do show some variation between the cases. Recirculation zones on the cowl surface is predicted for the two dimensional planar nozzle and a smaller one for the super elliptic nozzle. The recirculation is caused by the strong pressure gradient between the plume and freestream flows. The axi-symmetric nozzle results indicate recirculation zones on the thruster face. These recirculation zones smooth the pressure gradient between the plume and freestream flow limiting the formation of recirculation on the cowl surface. Thruster to thruster interaction is evident for the axi-symmetric and supper elliptic calculation while the two dimensional planar nozzle did not have any lateral expansion in the nozzle, so thruster to thruster interaction is limited. The integrated performance results, at the altitude choosen, show very little variation between the three thruster shapes. This result allows for nozzle shape determination based on additional considerations (thermal, structural, weight) besides performance.

Ungewitter, Ronald J.↗

NASA NTRSDate not supplied

Boundary Layer Stability and Transition in a Chemically Reacting Martian Atmosphere using LASTRAC

The Thermal Protection System (TPS) is a significant portion of the mass of reentry vehicles, planetary probes, and Martian entry vehicles. Reducing this mass has benefits in terms of decreased fuel requirements and increased payload; however, due to the high risk and un-certainty, the TPS or heat shields are designed conservatively by assuming fully turbulent flow. Laminar flow results in reduced heat flux, and improved transition prediction has the potential to reduce TPS mass and uncertainty in aerothermodynamic predictions. Limited previous research exists examining the problem of transition prediction for Martian atmospheric entry, with studies available on transition on the Mars Science Laboratory TPS. Although transition was demonstrated in wind-tunnel tests, uncertainty in the transition location resulted in a TPS designed for fully turbulent flow, and therefore greater mass than required for a partially-laminar condition. In this work, we extend the boundary layer stability code LASTRAC, recently modified to include chemical and thermal nonequilibrium capabilities, to a model of the Martian atmosphere. LASTRAC provides Parabolized Stability Equations (PSE) as well as Linear Stability Theory (LST) to predict the stability of a boundary layer and transition with semi-empirical eN methods. Results included in this work compare disturbance growth characteristics between air and Martian atmosphere at similar nondimensional freestream conditions on a simple flat plate geometry. Both chemical nonequilibrium and thermochemical nonequilibrium, as well as both PSE and LST, are used.

Heather L Kline↗

NASA NTRS2002

Cartesian-Grid Simulations of a Canard-Controlled Missile with a Free-Spinning Tail

The proposed paper presents a series of simulations of a geometrically complex, canard-controlled, supersonic missile with free-spinning tail fins. Time-dependent simulations were performed using an inviscid Cartesian-grid-based method with results compared to both experimental data and high-resolution Navier-Stokes computations. At fixed free stream conditions and canard deflections, the tail spin rate was iteratively determined such that the net rolling moment on the empennage is zero. This rate corresponds to the time-asymptotic rate of the free-to-spin fin system. After obtaining spin-averaged aerodynamic coefficients for the missile, the investigation seeks a fixed-tail approximation to the spin-averaged aerodynamic coefficients, and examines the validity of this approximation over a variety of freestream conditions.

Murman, Scott M.↗

NASA NTRS2019

Boundary Layer Stability and Transition in a Chemically Reacting Martian Atmosphere Using LASTRAC

The Thermal Protection System (TPS) is a significant portion of the mass of reentry vehicles, planetary probes, and Martian entry vehicles. Reducing this mass has benefits in terms of decreased fuel requirements and increased payload; however, due to the high risk and uncertainty, the TPS or heat shields are designed conservatively by assuming fully turbulent flow. Laminar flow results in reduced heat flux, and improved transition prediction has the potential to reduce TPS mass and uncertainty in aerothermodynamic predictions. Limited previous research exists examining the problem of transition prediction for Martian atmospheric entry, with studies available on transition on the Mars Science Laboratory TPS. Although transition was demonstrated in wind-tunnel tests, uncertainty in the transition location resulted in a TPS designed for fully turbulent flow, and therefore greater mass than required for a partially laminar condition. In this work, we extend the boundary layer stability code LASTRAC, recently modified to include chemical and thermal nonequilibrium capabilities, to a model of the Martian atmosphere. LASTRAC provides Parabolized Stability Equations (PSE) as well as Linear Stability Theory (LST) to predict the stability of a boundary layer and transition with semi-empirical e(sup N) methods. Results included in this work compare disturbance growth characteristics between air and Martian atmosphere at similar non-dimensional freestream conditions on a simple flat plate geometry. Both chemical nonequilibrium and thermochemical nonequilibrium, as well as both PSE and LST, are used.

Kline, H. L.↗

NASA NTRS2011

USM3D Simulations of Saturn V Plume Induced Flow Separation

The NASA Constellation Program included the Ares V heavy lift cargo vehicle. During the design stage, engineers questioned if the Plume Induced Flow Separation (PIFS) that occurred along Saturn V rocket during moon missions at some flight conditions, would also plague the newly proposed rocket. Computational fluid dynamics (CFD) was offered as a tool for initiating the investigation of PIFS along the Ares V rocket. However, CFD best practice guidelines were not available for such an investigation. In an effort to establish a CFD process and define guidelines for Ares V powered simulations, the Saturn V vehicle was used because PIFS flight data existed. The ideal gas, computational flow solver USM3D was evaluated for its viability in computing PIFS along the Saturn V vehicle with F-1 engines firing. Solutions were computed at supersonic freestream conditions, zero degree angle of attack, zero degree sideslip, and at flight Reynolds numbers. The effects of solution sensitivity to grid refinement, turbulence models, and the engine boundary conditions on the predicted PIFS distance along the Saturn V were discussed and compared to flight data from the Apollo 11 mission AS-506.

Deere, Karen↗

DOE OSTI2021

Quantifying wind plant blockage under stable atmospheric conditions

Wind plant blockage reduces the wind velocity upstream undermining turbine performance for the first row of the plant. We assess how atmospheric stability modifies the induction zone of a wind plant in flat terrain. We also explore different approaches to quantifying the magnitude and extent of the induction zone from field-like observations. To investigate the influence from atmospheric stability, we compare simulations of two stable boundary layers using the Weather Research and Forecasting model in large-eddy simulation mode, representing wind turbines using the generalized actuator disk approach. We find a faster cooling rate at the surface, which produces a stronger stably stratified boundary layer, amplifies the induction zone of both an isolated turbine and of a large wind plant. A statistical analysis on the hub-height wind speed field shows wind slowdowns only extend far upstream (up to 15D) of a wind plant in strong stable boundary layers. To evaluate different ways of measuring wind plant blockage from field-like observations, we consider various ways of estimating the freestream velocity upstream of the plant. Sampling a large area upstream is the most accurate approach to estimating the freestream conditions, and thus of measuring the blockage effect. Also, the choice of sampling method may induce errors of the same order as the velocity deficit in the induction zone.

54 ENVIRONMENTAL SCIENCES↗

DOE OSTI2022

Calibrating hypersonic turbulence flow models with the HIFiRE-1 experiment using data-driven machine-learned models.

In this paper we study the efficacy of combining machine-learning methods with projection-based model reduction techniques for creating data-driven surrogate models of computationally expensive, high-fidelity physics models. Such surrogate models are essential for many-query applications e.g., engineering design optimization and parameter estimation, where it is necessary to invoke the high-fidelity model sequentially, many times. Surrogate models are usually constructed for individual scalar quantities. However there are scenarios where a spatially varying field needs to be modeled as a function of the model’s input parameters. Here we develop a method to do so, using projections to represent spatial variability while a machine-learned model captures the dependence of the model’s response on the inputs. The method is demonstrated on modeling the heat flux and pressure on the surface of the HIFiRE-1 geometry in a Mach 7.16 turbulent flow. The surrogate model is then used to perform Bayesian estimation of freestream conditions and parameters of the SST (Shear Stress Transport) turbulence model embedded in the high-fidelity (Reynolds-Averaged Navier–Stokes) flow simulator, using shock-tunnel data. The paper provides the first-ever Bayesian calibration of a turbulence model for complex hypersonic turbulent flows. We find that the primary issues in estimating the SST model parameters are the limited information content of the heat flux and pressure measurements and the large model-form error encountered in a certain part of the flow.

42 ENGINEERING↗