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At least 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Effect of the Reynolds Number on the Freestream Disturbance Environment in a Mach 6 Nozzle

To understand the impact of unit Reynolds number on the acoustic disturbance field inside a high-speed wind tunnel, we use Direct Numerical Simulations (DNS) to model the turbulent boundary layers along the walls of a quasi-two-dimensional nozzle configuration. Intended as a stepping stone to fully three-dimensional simulations of freestream noise inside the NASA 20-Inch Mach 6 Wind Tunnel, the present simulations are based on periodic boundary conditions across the spanwise width of the computational domain that corresponds to about one third of the actual tunnel width. These simulations are performed at four different unit Reynolds numbers, ranging from 3.56e6 to 14.0e6 per meter. The predominantly hydrodynamic fluctuations inside the boundary layer are shown to be nearly unaffected by the presence of freestream forcing associated with the impinging acoustic radiation from the opposite wall. Thus, the Reynolds number trends associated with boundary-layer quantities are consistent with previously published DNS of flat-plate boundary layers at similar Mach numbers and wall temperature ratios. The unsteady disturbance environment within the nozzle core region is found to be approximately spatially homogeneous and purely acoustic in nature. The numerical results are used to make comparisons with tunnel noise measurements by Chou et al. [1,2]. Unlike previous comparisons involving static-pressure fluctuations based on the DNS and pitot-pressure fluctuations measured in the wind tunnel, direct comparisons involving the fluctuations in the same physical quantity, namely, the streamwise mass flux, have been reported for the first time. The predicted decrease in the root-mean-square fluctuations in pressure and mass flux with an increasing unit Reynolds number is in agreement with the measurements in the NASA 20-Inch Mach 6 Wind Tunnel. Additional details of the acoustic radiation field are quantified and should be useful toward a digital synthesis of the tunnel disturbance environment that would enable realistic simulations of the natural transition process.

Turbulence↗
NASA NTRSDate not supplied

Effect of the Reynolds Number on the Freestream Disturbance Environment in a Mach 6 Nozzle

To understand the impact of unit Reynolds number on the acoustic disturbance field inside a high-speed wind tunnel, we use Direct Numerical Simulations (DNS) to model the turbulent boundary layers along the walls of a quasi-two-dimensional nozzle configuration. Intended as a stepping stone to fully three-dimensional simulations of freestream noise inside the NASA 20-Inch Mach 6 Wind Tunnel, the present simulations are based on periodic boundary conditions across the spanwise width of the computational domain that corresponds to about one third of the actual tunnel width. These simulations are performed at four different unit Reynolds numbers, ranging from 3.56e6 to 14.0e6 per meter. The predominantly hydrodynamic fluctuations inside the boundary layer are shown to be nearly unaffected by the presence of freestream forcing associated with the impinging acoustic radiation from the opposite wall. Thus, the Reynolds number trends associated with boundary-layer quantities are consistent with previously published DNS of flat-plate boundary layers at similar Mach numbers and wall temperature ratios. The unsteady disturbance environment within the nozzle core region is found to be approximately spatially homogeneous and purely acoustic in nature. The numerical results are used to make comparisons with tunnel noise measurements by Chou et al. [1,2]. Unlike previous comparisons involving static-pressure fluctuations based on the DNS and pitot-pressure fluctuations measured in the wind tunnel, direct comparisons involving the fluctuations in the same physical quantity, namely, the streamwise mass flux, have been reported for the first time. The predicted decrease in the root-mean-square fluctuations in pressure and mass flux with an increasing unit Reynolds number is in agreement with the measurements in the NASA 20-Inch Mach 6 Wind Tunnel. Additional details of the acoustic radiation field are quantified and should be useful toward a digital synthesis of the tunnel disturbance environment that would enable realistic simulations of the natural transition process.

Acoustics↗
NASA NTRS2018

Direct Numerical Simulation of Acoustic Disturbances in the Rectangular Test Section of a Hypersonic Wind Tunnel

Direct numerical simulations (DNS) of the full-scale rectangular nozzle of a hypersonic wind tunnel are conducted to study the acoustic freestream fluctuations radiating from turbulent boundary layers (TBLs) along the nozzle walls. The nozzle geometry and the flow conditions of the DNS match those of the NASA 20-Inch Mach 6 Tunnel, and the DNS has been completed for a domain without spanwise sidewall boundary conditions. The turbulent boundary layer parameters based on the DNS compare well with those derived from Reynolds Averaged Navier-Stokes (RANS) calculations as well as with the predictions based on Pate’s correlation. A similarly good comparison is observed for both the Mach number distribution and the Reynold’s stresses obtained from the DNS and RANS calculations, respectively. Various characteristics of the acoustic pressure fluctuations within the inviscid core of the nozzle flow are compared with those associated with a single flat plate at a similar freestream Mach number. The frequency spectrum and bulk propagation speeds match well between the nozzle and the flat plate, but the rms pressure fluctuation is higher for the nozzle configuration, likely due to the combined effect of acoustic radiation from the top and bottom walls. Spatial contours of the two-point correlation coefficient display elliptical tails with approximately equal but opposite angles corresponding to the preferred directionality of acoustic structures radiated from both walls. Future work will focus on DNS of the full nozzle configuration, including the effects of the nozzle side walls.

Deegan, Cole P.↗
NASA NTRS1984

Prediction of transonic separated flows

Johnson et al. (1982) have provided a detailed comparison between a thoroughly documented transonic flow with shock-induced separations and solutions of the flow using the Navier-Stokes equations. According to this comparison, there were several deficiencies in the computations. The present investigation takes into account new experimental data which have been obtained in a larger wind tunnel with the same test model for a wider range of freestream Mach numbers. The results of new Navier-Stokes computations using more compatible boundary conditions are shown, and the effects of the turbulence model choice on predicting Mach number trends are assessed.

Horstman, C. C.↗
NASA NTRS1988

Prediction of unsteady loads on maneuvering delta wings using time-accurate Euler schemes

Three-dimensional steady and unsteady vortex-dominated flows around sharp-edged delta wings are considered in this paper. The problem is formulated by using the unsteady conservative Euler equations for the flow relative motion with respect to a moving frame of reference. An implicit approximately-factored finite volume scheme is used to solve the resulting equations on a three-dimensional computational grid which is generated by using a modified Joukowski transformation in cross-flow planes at the grid chord stations. The scheme is applied to a delta wing undergoing pitching oscillation around a large angle of attack. The initial conditions correspond to a steady flow around a delta wing of aspect ratio of one, freestream Mach number of 0.3 and mean angle of attack of 20.5. The steady flow results are compared with those of an explicit computational scheme and the experimental data, and they are in good agreement.

Kandil, Osama A.↗
NASA NTRS1991

Simultaneous measurements of temperature and density in air flows using UV laser spectroscopy

The simultaneous measurement of temperature and density using laser-induced fluorescence of oxygen in combination with Q-branch Raman scattering of nitrogen and oxygen is demonstrated in a low-speed air flow. The lowest density and temperature measured in the experiment correspond to the freestream values at Mach 5 in the Ames 3.5-Foot Hypersonic Wind Tunnel for stagnation conditions of 100 atm and 1000 K. The experimental results demonstrate the viability of the optical technique for measurements that support the study of compressible turbulence and the validation of numerical codes in supersonic and hypersonic wind tunnel flows.

Fletcher, D. G.↗
NASA NTRSDate not supplied

Stagnation Point Convective Heating Correlations for Entry Into H2/He Atmospheres

New convective heating correlations have been developed over a broad range of entry conditions into H2/He atmospheres. The correlations are based on calculations of stagnation point convective heat fluxes over a large range of velocities, densities, and effective nose radii. The heat fluxes and corresponding flowfields are computed using NASA’s hypersonic computational fluid dynamics code, DPLR. Cases in this study have been computed using both a temperaturelimited, radiative equilibrium wall and a cold wall temperature boundary condition. The results for the scaled heat flux, as a function of the difference in freestream and wall enthalpies, show a clear deviation from the linear trend predicted by the method of Sutton and Graves. Therefore, updated non-linear correlations are proposed for the prediction of stagnation point convective heat fluxes. Generally, these updated correlations predict lower heat loads and therefore smaller required TPS thicknesses when compared to previous correlations for a range of test cases in the Neptune, Saturn, and Uranus atmospheres.

Alex T. Carroll↗
NASA NTRSDate not supplied

Stagnation Point Convective Heating Correlations for Entry Into H2/He Atmospheres

New convective heating correlations have been developed over a broad range of entry conditions into H2/He atmospheres. The correlations are based on calculations of stagnation point convective heat fluxes over a large range of velocities, densities, and effective nose radii. The heat fluxes and corresponding flow fields are computed using NASA’s hypersonic computational fluid dynamics code, DPLR. Cases in this study have been computed using both a temperature limited, radiative equilibrium wall and a cold wall temperature boundary condition. The results for the scaled heat flux, as a function of the difference in freestream and wall enthalpies, show a clear deviation from the linear trend predicted by the method of Sutton and Graves. Therefore, updated non-linear correlations are proposed for the prediction of stagnation point convective heat fluxes. Generally, these updated correlations predict lower heat loads and therefore smaller required TPS thicknesses when compared to previous correlations for a range of test cases in the Neptune, Saturn, and Uranus atmospheres.

Alex T. Carroll↗
NASA NTRS1971

Detailed investigation of flowfields within large scale hypersonic inlet models

Analytical and experimental investigations were conducted to determine the characteristics of the internal flows in model passages representative of hypersonic inlets and also sufficiently large for meaningful data to be obtained. Three large-scale inlet models, each having a different compression ratio, were designed to provide high performance and approximately uniform static-pressure distributions at the throat stations. A wedge forebody was used to simulate the flowfield conditions at the entrance of the internal passages, thus removing the actual vehicle forebody from consideration in the design of the wind-tunnel models. Tests were conducted in a 3.5 foot hypersonic wind tunnel at a nominal test Mach number of 7.4 and freestream unit Reynolds number of 2,700,000 per foot. From flowfield survey data the inlet entrance, the entering inviscid and viscous flow conditions were determined prior to the analysis of the data obtained in the internal passages. Detailed flowfield survey data were obtained near the centerlines of the internal passages to define the boundary-layer development on the internal surfaces and the internal shock-wave configuration. Finally, flowfield data were measured across the throats of the inlet models to evaluate the internal performance of the internal passages. These data and additional results from surface instrumentation and flow visualization studies were utilized to determine the internal flowfield patterns and the inlet performance.

Seebaugh, W. R.↗
NASA NTRS2016

Acoustic Radiation From a Mach 14 Turbulent Boundary Layer

Direct numerical simulations (DNS) are used to examine the turbulence statistics and the radiation field generated by a high-speed turbulent boundary layer with a nominal freestream Mach number of 14 and wall temperature of 0:18 times the recovery temperature. The flow conditions fall within the range of nozzle exit conditions of the Arnold Engineering Development Center (AEDC) Hypervelocity Tunnel No. 9 facility. The streamwise domain size is approximately 200 times the boundary-layer thickness at the inlet, with a useful range of Reynolds number corresponding to Re 450 􀀀 650. Consistent with previous studies of turbulent boundary layer at high Mach numbers, the weak compressibility hypothesis for turbulent boundary layers remains applicable under this flow condition and the computational results confirm the validity of both the van Driest transformation and Morkovin's scaling. The Reynolds analogy is valid at the surface; the RMS of fluctuations in the surface pressure, wall shear stress, and heat flux is 24%, 53%, and 67% of the surface mean, respectively. The magnitude and dominant frequency of pressure fluctuations are found to vary dramatically within the inner layer (z/delta 0.< or approx. 0.08 or z+ < or approx. 50). The peak of the pre-multiplied frequency spectrum of the pressure fluctuation is f(delta)/U(sub infinity) approx. 2.1 at the surface and shifts to a lower frequency of f(delta)/U(sub infinity) approx. 0.7 in the free stream where the pressure signal is predominantly acoustic. The dominant frequency of the pressure spectrum shows a significant dependence on the freestream Mach number both at the wall and in the free stream.

Zhang, Chao↗
NASA NTRSDate not supplied

Testing of Mars Retropropulsion Concepts in the Langley Unitary Plan Wind Tunnel

Future Mars human lander missions will require using retrorockets beginning at supersonic conditions, something which has never been done before at Mars. Significant aerosciences challenges exist due to interactions between the retrorocket exhaust plumes, the freestream flow, and the vehicle that alter the aerodynamic behavior. Historical retropropulsion ground and flight data do not provide a basis for calibrating the computational uncertainties for aerodynamic interference forces and moments on proposed Mars descent vehicles, either because of insufficient data or dissimilar vehicles geometries and/or conditions. A new retropropulsion wind tunnel test was conducted in the NASA Langley Unitary Plan Wind Tunnel in order to address some of the deficiencies of previous tests and to provide a basis for computational model comparisons on relevant geometries. The test included improvements in model design and measured data. Two sub-scale model geometries derived from full-scale vehicle concepts were tested: a blunt low lift-to-drag vehicle and a more slender geometry. The first model includes variations in nozzle parameters that were estimated to have first-order effects on the measurements: location, pointing direction, size, and exit-to-throat area ratio. Both models were instrumented with discrete steady and unsteady pressure sensors on the aerodynamic and interior flow path surfaces. The heats shields of both models also were covered with pressure sensitive paint that effectively measures a steady pressure distribution in an otherwise unsteady flow field. This pressure distribution was integrated over the surface geometry in order to estimate the steady aerodynamic force and moment coefficient contribution from the painted surfaces; this approach was used as a replacement for using a flow-through balance. This paper will cover the test objectives, model design, test matrix, and available test results for each model configuration at different freestream Mach numbers, model attitude, and non-dimensional thrust magnitude. A subsequent publication will cover uncertainty quantification of the test data and comparisons to computational fluid dynamics model results.

Powered Descent↗
NASA NTRS1998

Experimental Study of Transitional Flow Behavior in a Simulated Low Pressure Turbine

A detailed investigation of the flow physics occurring on the suction side of a simulated Low Pressure Turbine (LPT) blade was performed. A contoured upper wall was designed to simulate the pressure distribution of an actual LPT airfoil onto a flat lower plate. The experiments were carried out for the Reynolds numbers of 35,000, 70,000, 100,000 and 250,000 with four levels of freestream turbulence ranging from 1% to 4%. For the three lower Reynolds numbers, the boundary layer on the flat plate was separated and formed a bubble. The size of laminar separation bubble was measured to be inversely proportional to the freestream turbulence levels and Reynolds numbers. However, no separation was observed for the Re = 250,000 case. The transition on a separated flow was found to proceed through the formation of turbulent spots in the free shear layer as evidenced in the intermittency profiles for Re = 35,000, 70,000 and 100,000. Spectral data show no evidence of Kelvin-Helmholtz or Tollmien-Schlichting instability waves in the free shear layer over a separation bubble (bypass transition). However, the flow visualization revealed the large vortex structures just outside of the bubble and their development to turbulent flow for Re = 50,000, which is similar to that in the free shear layer (separated-flow transition). Therefore, it is fair to say that the bypass and separated-flow transition modes coexist in the transitional flows over the separation bubble for certain conditions. Transition onset and end locations and length determined from intermittency profiles decrease as Reynolds number and freestream turbulence levels increase.

Sohn, Ki Hyeon↗
NASA NTRS2007

Experimental Study of Transitional Flow Behavior in a Simulated Low Pressure Turbine

A detailed investigation of the flow physics occurring on the suction side of a simulated Low Pressure Turbine (LPT) blade was performed. A contoured upper wall was designed to simulate the pressure distribution of an actual LPT airfoil onto a flat lower plate. The experiments were carried out for the Reynolds numbers of 35,000, 70,000, 100,000, and 250,000 with four levels of freestream turbulence ranging from 1 to 4 percent. For the three lower Reynolds numbers, the boundary layer on the flat plate was separated and formed a bubble. The size of laminar separation bubble was measured to be inversely proportional to the freestream turbulence levels and Reynolds numbers. However, no separation was observed for the Re = 250,000 case. The transition on a separated flow was found to proceed through the formation of turbulent spots in the free shear layer as evidenced in the intermittency profiles for Re = 35,000, 70,000, and 100,000. Spectral data show no evidence of Kelvin-Helmholtz of Tollmien-Schlichting instability waves in the free shear layer over a separation bubble (bypass transition). However, the flow visualization revealed the large vortex structures just outside of the bubble and their development to turbulent flow for Re = 50,000, which is similar to that in the free shear layer (separated-flow transition). Therefore, it is fair to say that the bypass and separated-flow transition modes coexist in the transition flows over the separation bubble of certain conditions. Transition onset and end locations and length determined from intermittency profiles decreased as Reynolds number and freestream turbulence levels increase.

Sohn, Ki Hyeon↗
NASA NTRS2002

Prediction of Drag Reduction in Supersonic and Hypersonic Flows with Counterflow Jets

Computational fluid dynamics solutions of the flowfield of a truncated cone-cylinder with and without counterflow jets have been obtained for the short penetration mode (SPM) and long penetration mode (LPM) of the freestream-counterflow jet interaction flowfield. For the case without the counterflow jet, the comparison of the normalized surface pressures showed very good agreement with experimental data. For the case with the SPM jet, the predicted surface pressures did not compare as well with the experimental data upstream of the expansion corner, while aft of the expansion corner, the comparison of the solution and the data is seen to give much better agreement. The difference in the prediction and the data could be due to the transient character of the jet penetration modes, possible effects of the plasma physics that are not accounted for here, or even the less likely effect of flow turbulence, etc. For the LPM jet computations, one-dimensional isentropic relations were used to derived the jet exit conditions in order to obtain the LPM solutions. The solution for the jet exit Mach number of 3 shows a jet penetration several times longer than that of the SPM, and therefore much weaker bow shock, with an attendant reduction in wave drag. The LPM jet is, in essence, seen to be a "pencil" of fluid, with much higher dynamic pressure, embedded in the oncoming supersonic or hypersonic freestream. The methodology for determining the conditions for the LPM jet could enable a practical approach for the design and application of counterflow LPM jets for the reduction of wave drag and heat flux, thus significantly enhancing the aerodynamic characteristics and aerothermal performance of supersonic and hypersonic vehicles. The solutions show that the qualitative flow structure is very well captured. The obtained results, therefore, suggest that counterflowing jets are viable candidate technology concepts that can be employed to give significant reductions in wave drag, heat flux, and other attendant aerodynamic benefits.

Daso, Endwell O.↗
NASA NTRS2020

Model Design and Pre-Test CFD Analysis for a Supersonic Retropropulsion Wind Tunnel Test

Future Mars human lander missions will require using powered descent beginning at supersonic conditions, something which has never been done before at Mars. Significant aerosciences challenges exist due to interactions between the retrorocket exhaust plumes and the freestream flow that alter the aerodynamic behavior of the powered descent vehicle. Historically, wind tunnel tests have been used to study supersonic retropropulsion with inert gas exhaust simulants. Also, SpaceX has successfully decelerated the Falcon 9 first stage numerous times by using powered flight at supersonic conditions. On the computational side, Reynolds-Averaged Navier-Stokes flowfield simulations are regularly completed at full-scale conditions. However, the available ground and flight data do not provide a basis for calibrating the computational uncertainties for aerodynamic interference forces and moments on proposed Mars descent vehicles, either because of insufficient data or dissimilar vehicles geometries and/or conditions. Thus, additional ground testing is needed to continue addressing the aerosciences risks for large-scale human Mars landers. To that end, a retropropulsion wind tunnel test will be conducted in the NASA Langley Unitary Plan Wind Tunnel in 2020. The test is designed to improve upon similar past tests, both in terms of model design and measured data. The test campaign will use subscale model geometries derived from the two current NASA powered descent reference vehicles: a blunt low lift-to-drag vehicle and a more slender geometry. This paper covers the reference vehicles, test objectives, model design, scaling parameters, test matrix, and computational analysis to date.

Karl T Edquist↗
NASA NTRS1975

Results of aerodynamic heat transfer tests of a 0.0175-scale model of the Rockwell International Space Shuttle Orbiter 139 (model number 22-0) in the NASA/Ames 3.5-foot hypersonic wind tunnel (test OH6)

The results of a hypersonic wind tunnel test program conducted using a 0.0175 scale thin-skin thermocouple model of the Space Shuttle Orbiter to obtain aerodynamic heat transfer data on the Orbiter under simulated reentry conditions were presented. The test program was conducted at a Mach number of 7.3 and a freestream Reynolds number ranging between 1.0 and 6.0 million/foot. The model was tested for angles of attack ranging between 10 deg and 30 deg and a sideslip angle of 0 deg. The model was constructed of 15-5 PH stainless steel with the instrumented areas machined to a nominal skin thickness of 0.030 in. The model instrumentation consisted of 288 iron-constantan thermocouples spot welded to the skin inner surface, but only 75 of these were used in this test program. A high-speed, analog-to-digital data acquisition system was used to record data on magnetic tape.

Dye, W. H.↗
NASA NTRS1976

Correlations for determining thermodynamic properties of hydrogen-helium gas mixtures at temperatures from 7,000 to 35,000 K

Simple relations for determining the enthalpy and temperature of hydrogen-helium gas mixtures were developed for hydrogen volumetric compositions from 1.0 to 0.7. These relations are expressed as a function of pressure and density and are valid for a range of temperatures from 7,000 to 35,000 K and pressures from 0.10 to 3.14 MPa. The proportionality constant and exponents in the correlation equations were determined for each gas composition by applying a linear least squares curve fit to a large number of thermodynamic calculations obtained from a detailed computer code. Although these simple relations yielded thermodynamic properties suitable for many engineering applications, their accuracy was improved significantly by evaluating the proportionality constants at postshock conditions and correlating these values as a function of the gas composition and the product of freestream velocity and shock angle. The resulting equations for the proportionality constants in terms of velocity and gas composition and the corresponding simple realtions for enthalpy and temperature were incorporated into a flow field computational scheme. Comparison was good between the thermodynamic properties determined from these relations and those obtained by using a detailed computer code to determine the properties. Thus, an appreciable savings in computer time was realized with no significant loss in accuracy.

Zoby, E. V.↗
NASA NTRS1990

Assured Crew Return Vehicle flowfield and aerodynamic characteristics

A lifting body has been proposed as a candidate for the Assured Crew Return Vehicle which will serve as crew rescue vehicle for the Space Station. The focus of this work is on body surface definition, surface and volume grid definition, and the computation of inviscid flowfields about the vehicle at wind-tunnel conditions. Very good agreement is shown between the computed aerodynamic characteristics of the vehicle at a freestream Mach number of 10 and those measured in wind-tunnel tests.

Weilmuenster, K. James↗