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NASA NTRS2020

Toward Transition Modeling in a Hypersonic Boundary Layer at Flight Conditions

An accurate physics-based transition prediction method integrated with computational fluid dynamics (CFD) solvers is pursued for hypersonic boundary layer flows over slender hypersonic vehicles at flight conditions. The geometry and flow conditions are selected to match relevant trajectory locations from the ascent phase of the HIFiRE-1 flight experiment, namely, a 7-degree half-angle cone with 2.5 mm nose radius, freestream Mach numbers in the range of 3.8 – 5.5 and freestream unit Reynolds numbers in the range of 3.3 × 10(exp 6) – 21.4 × 10(exp 6) m(exp -1). Earlier research had shown that the onset of transition during the HIFiRE-1 flight experiment correlated with an amplification factor of N ≈ 13.5 for the planar Mack modes. However, to incorporate the N-factor correlations into a CFD code, we investigate surrogate models for disturbance amplification that avoid the direct computation of stability characteristics. A commonly used approach for low-speed flows is based on an a priori database of stability characteristics for locally similar profiles. However, the results presented in this paper demonstrate that the application of this approach to hypersonic boundary layers over blunt spherical nose-tip cones leads to large, unacceptable errors in the predictions of amplification factors, mainly due to its failure in accounting for the effects of the entropy layer on the boundary-layer profiles along the length of the model. We propose and demonstrate an alternate approach that employs the stability computations for a canonical set of blunt cone configurations to train a physics-informed convolutional neural network model that is shown to provide substantially improved transition predictions for hypersonic flow configurations with entropy-layer effects. Furthermore, the excellent performance of the neural network model is also confirmed for cone configurations with nose radius and half-angle values that do not correspond to those used to build the database. Finally, the convolutional neural network model is shown to outperform the linear stability calculations for underresolved basic states.

Pedro Paredes↗

NASA NTRS1984

Transonic flow calculations using a flux vector splitting method for the Euler equations

A study of the flux vector splitting method of Steger and Warming for the solution of the time dependent Euler equations in strong conservation law form for arbitrary two-dimensional geometries is presented. The procedure employed here differs from that of Buning and Steger in that it uses a different algorithm and employs implicit boundary conditions. Moreover, the method, as implemented here, does not contain any explicit smoothing or any adjustable parameters. Calculations were carried out for an NACA 0012 airfoil at various Mach numbers and angles of attack, and cylinders. Steady symmetric solutions were obtained for the full cylinder at a freestream Mach number of .5 without imposing a symmetry condition. In general, good agreement with other methods was obtained.

Seaford, C. M.↗

NASA NTRS1994

Semi-span model testing in the national transonic facility

The present work was motivated by an ongoing research program at NASA Langley Research Center to develop a semi-span testing capability for the National Transonic Facility (NTF). This test technique is being investigated as a means to design and optimize high-lift devices at flight Reynolds numbers in a ground test facility. Even though the freestream Mach numbers of interest are around .20, the flow around a transport wing with high lift devices deployed may contain regions of compressible flow. Thus to properly model the flow physics, a compressible flow solver may be required. However, the application of a compressible flow solver at low Mach numbers can be problematic. The objective of this phase of the project is to directly compare the performance of two widely used three-dimensional compressible Navier-Stokes solvers at low Mach numbers to both experimental data and to results obtained from an incompressible Navier-Stokes solver. The geometries of interest are two isolated wings with different leading edge sweep angles. The compressible Navier-Stokes solvers chosen, TLNS3D-MB and CFL3D, which were developed at NASA Langley Research Center (LaRC), represent the current state-of-the-art in compressible 3-D Navier-Stokes solvers. The incompressible Navier-Stokes solver, INS3D-UP, developed recently at NASA Ames Research Center (ARC), represents the current state-of-the-art in incompressible Navier-Stokes solvers.

Chokani, Ndaona↗

NASA NTRS1997

Hypersonic Pitching-Moment Shift for Stardust Reentry Capsule Forebody

Aerodynamic coefficients are presented for perfect-gas and equilibrium-air solutions of the Navier-Stokes equations about the Stardust reentry-capsule forebody at Mach numbers of 4.6, 7, 8.5, and 10. A comparison with Newtonian-flow assumptions indicates a divergence of the aerodynamic coefficients from Newtonian-flow for Mach numbers less than 10. The static stability of the forebody is reduced by a factor of 2.5 with decreasing freestream Mach number between Mach 10 and 7.

Wood, William A.↗

NASA NTRS1991

Opportunities for applications of natural laminar flow technology at high-subsonic speeds

The possible profiles of high subsonic speed airfoils with extensive regions of natural laminar flow (NLF) are explored, on the bases of calculations which suggest that high subsonic Mach number NLFs are obtainable for both swept and unswept wing applications at certain Reynolds numbers. Attention is given to the transonic pressure distributions of airfoils for unswept wings at freestream Mach numbers of 0.65-0.80 and chord Reynolds numbers of up to 50 million. The case of 10-30 deg swept-wing NLF airfoils is also investigated for chord Reynolds numbers of 15-50 million.

Viken, Jeff↗

NASA NTRS1979

Reynolds number effects on the turbulence field in compressible boundary layers

Detailed experiments were conducted in a zero pressure gradient, supersonic turbulent boundary layer, including measurements of the three components of velocity fluctuations and the turbulent shear stress, for Reynolds numbers ranging from 11.7 million to 105 million at a freestream Mach number of 2.3. The mean flow measurements established the fully developed and equilibrium nature of the boundary layer. Measurements of the turbulence field show that the vertical and transverse fluctuations are essentially equal throughout the boundary layer at all Reynolds numbers, a feature that is different from observations in incompressible flows. The data show that the boundary layer exhibits similarity in the turbulence profiles for the entire Reynolds number range and agrees with previous compressible and incompressible data using Morkovin's scaling to account for compressibility effects.

Acharya, M.↗

NASA NTRSDate not supplied

Effect of Roughness on Transition on the BOLT-1a Geometry in Supersonic Flow

Infrared thermography measurements on an 8.8% scale BOLT-1a model were made in the NASA Langley Probe Calibration Tunnel for freestream unit Reynolds numbers from 7.7 × 10^ -6 /m to 30.7 × 10^ -6 /m, a freestream Mach number of 3.5, and nominal angle of attack of 0 degrees. The model was printed from polycarbonate and tested before and after fine-grit sanding to evaluate the effect of surface roughness on transition. Global surface roughness measurements were obtained using a 3D optical profilometer. The temperature distributions on the upper and lower experimental surfaces of the test article were simultaneously captured using two infrared cameras. Prior to testing, the facility’s freestream environment was characterized. At a freestream unit Reynolds number of approximately 17 × 10^ -6 /m, the nozzle-wall boundary layer began to transition non-uniformly, inducing transition on the model. The boundary-layer state was determined by analyzing Stanton number distributions computed from the surface temperature measurements. The present supersonic conditions that were examined on BOLT-1a were found to display both similarities and disparities in comparison to previous hypersonic experiments.

boundary layer transition↗

NASA NTRS2004

Real-Time Adaptive Control of Flow-Induced Cavity Tones

An adaptive generalized predictive control (GPC) algorithm was formulated and applied to the cavity flow-tone problem. The algorithm employs gradient descent to update the GPC coefficients at each time step. The adaptive control algorithm demonstrated multiple Rossiter mode suppression at fixed Mach numbers ranging from 0.275 to 0.38. The algorithm was also able t o maintain suppression of multiple cavity tones as the freestream Mach number was varied over a modest range (0.275 to 0.29). Controller performance was evaluated with a measure of output disturbance rejection and an input sensitivity transfer function. The results suggest that disturbances entering the cavity flow are colocated with the control input at the cavity leading edge. In that case, only tonal components of the cavity wall-pressure fluctuations can be suppressed and arbitrary broadband pressure reduction is not possible. In the control-algorithm development, the cavity dynamics are treated as linear and time invariant (LTI) for a fixed Mach number. The experimental results lend support this treatment.

Kegerise, Michael A.↗

NASA NTRS1978

Investigation of jet exhaust and local Mach number effects on the Solid Rocket Booster (SRB) base pressures (fa19)

Two single nozzles with flare angles of 10 and 20 degrees were tested at Mach numbers of 0.5, 0.9, 1.2, 1.46, 1.96 and 3.48 in the presence of gaseous plumes. An attempt was made to determine the local Mach number above the flare by utilizing a pitot probe. This objective was only partially satisfied because the 20 degree flare separated the flow ahead of the flare for Mach numbers of 0.5 to 1.96. An accurate local Mach number could not be determined because of the separated flow. To meet the objective of a data base as a function of freestream Mach number, model surface and base pressures were obtained in the presence of gaseous plumes for a matrix of chamber pressures and temperatures at Mach numbers of 0.5, 0.9, 1.2, 1.46, 1.96 and 3.48.

Love, D. A.↗

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 NTRS1993

Generic hypersonic vehicle performance model

An integrated computational model of a generic hypersonic vehicle was developed for the purpose of determining the vehicle's performance characteristics, which include the lift, drag, thrust, and moment acting on the vehicle at specified altitude, flight condition, and vehicular configuration. The lift, drag, thrust, and moment are developed for the body fixed coordinate system. These forces and moments arise from both aerodynamic and propulsive sources. SCRAMjet engine performance characteristics, such as fuel flow rate, can also be determined. The vehicle is assumed to be a lifting body with a single aerodynamic control surface. The body shape and control surface location are arbitrary and must be defined. The aerodynamics are calculated using either 2-dimensional Newtonian or modified Newtonian theory and approximate high-Mach-number Prandtl-Meyer expansion theory. Skin-friction drag was also accounted for. The skin-friction drag coefficient is a function of the freestream Mach number. The data for the skin-friction drag coefficient values were taken from NASA Technical Memorandum 102610. The modeling of the vehicle's SCRAMjet engine is based on quasi 1-dimensional gas dynamics for the engine diffuser, nozzle, and the combustor with heat addition. The engine has three variable inputs for control: the engine inlet diffuser area ratio, the total temperature rise through the combustor due to combustion of the fuel, and the engine internal expansion nozzle area ratio. The pressure distribution over the vehicle's lower aft body surface, which acts as an external nozzle, is calculated using a combination of quasi 1-dimensional gas dynamic theory and Newtonian or modified Newtonian theory. The exhaust plume shape is determined by matching the pressure inside the plume, calculated from the gas dynamic equations, with the freestream pressure, calculated from Newtonian or Modified Newtonian theory. In this manner, the pressure distribution along the vehicle after body expansion surface is then determined. The aerodynamic modeling, the engine modeling, and the exhaust plume analysis are described in more detail. A description of the computer code used to perform the above calculations is given and an input/output example is then given. The computer code is available on a Macintosh floppy disk.

Chavez, Frank R.↗

NASA NTRS2002

Adaptive Identification and Control of Flow-Induced Cavity Oscillations

Progress towards an adaptive self-tuning regulator (STR) for the cavity tone problem is discussed in this paper. Adaptive system identification algorithms were applied to an experimental cavity-flow tested as a prerequisite to control. In addition, a simple digital controller and a piezoelectric bimorph actuator were used to demonstrate multiple tone suppression. The control tests at Mach numbers of 0.275, 0.40, and 0.60 indicated approx. = 7dB tone reductions at multiple frequencies. Several different adaptive system identification algorithms were applied at a single freestream Mach number of 0.275. Adaptive finite-impulse response (FIR) filters of orders up to N = 100 were found to be unsuitable for modeling the cavity flow dynamics. Adaptive infinite-impulse response (IIR) filters of comparable order better captured the system dynamics. Two recursive algorithms, the least-mean square (LMS) and the recursive-least square (RLS), were utilized to update the adaptive filter coefficients. Given the sample-time requirements imposed by the cavity flow dynamics, the computational simplicity of the least mean squares (LMS) algorithm is advantageous for real-time control.

Kegerise, M. A.↗

NASA NTRS1993

Supersonic flow and shock formation in turbine tip gaps

Shock formation due to overexpansion of supersonic flow at the inlet to the tip clearance gap of a turbomachine has been studied. As the flow enters the tip gap, it accelerates around the blade pressure-side corner creating a region of minimum static pressure. The 'free streamline' separates from the wall at the corner; and, for Mach numbers greater than about 1.3, it curves back to intersect the blade tip. At this point, the freestream flow is abruptly turned parallel to the surface, giving rise to an oblique shock. The results are consistent with compressible sharp-edged orifice flow calculations found in the literature and with the theory of oblique shock wave formation in supersonic flow over a wedge. For freestream Mach numbers of 1.4 to 1.8, wave angles are 43 to 54 deg, and turning angles are 9 to 20 deg; as the Mach number increases, the angle of turn also increases. It appears that in a turbine, after separating from the inlet corner, the flow reattaches on the blade tip and an oblique shock is formed at 0.4-1.4 tip gap heights into the clearance gap. The resulting shock-boundary layer interaction may contribute to further enhancement of already high heat transfer to the blade tip in this region. This in turn could lead to higher blade temperatures and adversely affect blade life and turbine efficiency.

Moore, John↗

NASA NTRS1975

Aerodynamic heating to corrugation stiffened structures in thick turbulent boundary layers

The results of an experimental program to evaluate heat transfer and pressure distributions on corrugation roughened flat plates in thick turbulent boundary layers are presented. The experimental program consisted of tests in the tunnel wall boundary layers of the Langley Unitary Plan Wind Tunnel (UPWT) and Continuous Flow Hypersonic Tunnel (CFHT) at freestream Mach numbers of 2.5, 3.5, 4.5, and 10.3. Tests in the UPWT were conducted at a freestream Reynolds number/m of 10,800,000 and in the CFHT, at Reynolds numbers/m of 1,300,000 to 5,800,000. The test configurations consisted of 50.8 x 50.8 cm panels with corrugated beads of two different peak amplitudes, 0.61 and 0.29 cm. The angle of the corrugated beads relative to the flow direction was varied between 0 (aligned) and 90 deg (normal). The measured peak and average heat transfer are analyzed and correlated in terms of the bulk boundary layer, internal boundary layer, and geometric parameters.-

Brandon, H. J.↗

NASA NTRS1980

Complete viscous flowfield solutions about a blunt parabolic body in a supersonic stream

A hybrid computational technique which splits the flowfield into inviscid and viscous regions is used to investigate the complete flowfield about axisymmetric parabolic blunt bodies in a supersonic stream. The solutions are carried out on the CDC CYBER-203 computer which, with its extensive memory, allows for the use of a large number of finite-difference mesh points, allowing resolution of important flowfield features. A range of freestream Mach number of 2-5 and a range of Re number based on nose radius of 500-125,000 was run for a sonic corner body. Contour plots of density, pressure, and Mach number, velocity vector plots, and surface distributions of pressure, heat transfer, and shear stress are presented. Also, correlations of the downstream extent of the base recirculation region with Re number based on nose radius are given.

Weilmuenster, K. J.↗

NASA NTRS1981

Computation of airfoil buffet boundaries

The ILLIAC IV computer has been programmed with an implicit, finite-difference code for solving the thin layer compressible Navier-Stokes equation. Results presented for the case of the buffet boundaries of a conventional and a supercritical airfoil section at high Reynolds numbers are found to be in agreement with experimentally determined buffet boundaries, especially at the higher freestream Mach numbers and lower lift coefficients where the onset of unsteady flows is associated with shock wave-induced boundary layer separation.

Levy, L. L., Jr.↗

NASA NTRS2009

Aerodynamics of a Gulfstream G550 Nose Landing Gear Model

In this paper we discuss detailed steady and unsteady aerodynamic measurements of a Gulfstream G550 nose landing gear model. The quarter-scale, high-fidelity model includes part of the lower fuselage and the gear cavity. The full model configuration allowed for removal of various gear components (e.g. light cluster, steering mechanism, hydraulic lines, etc.) in order to document their effects on the local flow field. The measurements were conducted at a Reynolds number of 7.3 x 10(exp 4) based on the shock strut (piston) diameter and a freestream Mach number of 0.166. Additional data were also collected at lower Mach numbers of 0.12 and 0.145 and correspondingly lower Reynolds numbers. The boundary layer on the piston was tripped to enable turbulent flow separation, so as to better mimic the conditions encountered during flight. Steady surface pressures were gathered from an extensive number of static ports on the wheels, door, fuselage, and within the gear cavity. To better understand the resultant flow interactions between gear components, surface pressure fluctuations were collected via sixteen dynamic pressure sensors strategically placed on various subcomponents of the gear. Fifteen of the transducers were flush mounted on the gear surface at fixed locations, while the remaining one was a mobile transducer that could be placed at numerous varying locations. The measured surface pressure spectra are mainly broadband in nature, lacking any local peaks associated with coherent vortex shedding. This finding is in agreement with off-surface flow measurements using PIV that revealed the flow field to be a collection of separated shear layers without any dominant vortex shedding processes.

Neuhart, Dan H.↗

NASA NTRS2014

Direct Numerical Simulations of High-Speed Turbulent Boundary Layers over Riblets

Direct numerical simulations (DNS) of spatially developing turbulent boundary layers over riblets with a broad range of riblet spacings are conducted to investigate the effects of riblets on skin friction at high speeds. Zero-pressure gradient boundary layers under two flow conditions (Mach 2:5 with T(sub w)/T(sub r) = 1 and Mach 7:2 with T(sub w)/T(sub r) = 0:5) are considered. The DNS results show that the drag-reduction curve (delta C(sub f)/C(sub f) vs l(sup +)(sub g )) at both supersonic speeds follows the trend of low-speed data and consists of a `viscous' regime for small riblet size, a `breakdown' regime with optimal drag reduction, and a `drag-increasing' regime for larger riblet sizes. At l l(sup +)(sub g) approx. 10 (corresponding to s+ approx 20 for the current triangular riblets), drag reduction of approximately 7% is achieved at both Mach numbers, and con rms the observations of the few existing experiments under supersonic conditions. The Mach- number dependence of the drag-reduction curve occurs for riblet sizes that are larger than the optimal size, with smaller slopes of (delta C(sub f)/C(sub f) for larger freestream Mach numbers. The Reynolds analogy holds with 2(C(sub h)=C(sub f) approximately equal to that of at plates for both drag-reducing and drag-increasing configurations.

Duan, Lian↗