Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Volume 1; Sharp Leading Edge; [conducted in the Langley National Transonic Facility (NTF)

NASA Technical Reports Server (NTRS)

Chu, Julio; Luckring, James M.

1996-01-01

An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 36 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at a Reynolds number of 6 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Steady and unsteady transonic pressure measurements on a clipped delta wing for pitching and control-surface oscillations

NASA Technical Reports Server (NTRS)

Hess, Robert W.; Cazier, F. W., Jr.; Wynne, Eleanor C.

1986-01-01

Steady and unsteady pressures were measured on a clipped delta wing with a 6-percent circular-arc airfoil section and a leading-edge sweep angle of 50.40 deg. The model was oscillated in pitch and had an oscillating trailing-edge control surface. Measurements were concentrated over a Mach number range from 0.88 to 0.94; less extensive measurements were made at Mach numbers of 0.40, 0.96, and 1.12. The Reynolds number based on mean chord was approximately 10 x 10 to the 6th power. The interaction of wing or control-surface deflection with the formation of shock waves and with a leading-edge vortex generated complex pressure distributions that were sensitive to frequency and to small changes in Mach number at transonic speeds.

A wind tunnel investigation of circular and straked cylinders in transonic cross flow

NASA Technical Reports Server (NTRS)

Macha, J.

1976-01-01

Pressure distributions around circular and circular/strake cylinders were measured in a wind tunnel at Mach numbers from 0.6 to 1.2 with Reynolds number independently variable from 10,000 to 100,000. The local pressures are integrated over the cylinder surface to determine the variation of drag coefficient with both Mach number and Reynolds number. Effects of tunnel blockage are evaluated by comparing results from circular cylinders of various diameters at common Mach and Reynolds number conditions. Compressibility effects are concluded to be responsible for a flight reduction of the drag coefficient near Mach 0.7. Drag increases with strake height, presumably approaching a maximum drag corresponding to a flat plate configuration.

Surface pressure and inviscid flow field properties McDonnell-Douglas booster nominal Mach number of 8, volume 3

NASA Technical Reports Server (NTRS)

Matthews, R. K.; Martindale, W. R.; Warmbrod, J. D.

1972-01-01

The results are presented of a wind tunnel test program to determine surface pressures and flow field properties on the space shuttle booster configuration. The tests were conducted in September 1971. Data were obtained at a nominal Mach number of 8 at angles of attack of 40 and 50 deg and at a free stream unit Reynolds number of 3.7 million per foot.

Experimental Investigation of Unsteady Shock Wave Turbulent Boundary Layer Interactions About a Blunt Fin

NASA Technical Reports Server (NTRS)

Barnhart, Paul J.; Greber, Isaac

1997-01-01

A series of experiments were performed to investigate the effects of Mach number variation on the characteristics of the unsteady shock wave/turbulent boundary layer interaction generated by a blunt fin. A single blunt fin hemicylindrical leading edge diameter size was used in all of the experiments which covered the Mach number range from 2.0 to 5.0. The measurements in this investigation included surface flow visualization, static and dynamic pressure measurements, both on centerline and off-centerline of the blunt fin axis. Surface flow visualization and static pressure measurements showed that the spatial extent of the shock wave/turbulent boundary layer interaction increased with increasing Mach number. The maximum static pressure, normalized by the incoming static pressure, measured at the peak location in the separated flow region ahead of the blunt fin was found to increase with increasing Mach number. The mean and standard deviations of the fluctuating pressure signals from the dynamic pressure transducers were found to collapse to self-similar distributions as a function of the distance perpendicular to the separation line. The standard deviation of the pressure signals showed initial peaked distribution, with the maximum standard deviation point corresponding to the location of the separation line at Mach number 3.0 to 5.0. At Mach 2.0 the maximum standard deviation point was found to occur significantly upstream of the separation line. The intermittency distributions of the separation shock wave motion were found to be self-similar profiles for all Mach numbers. The intermittent region length was found to increase with Mach number and decrease with interaction sweepback angle. For Mach numbers 3.0 to 5.0 the separation line was found to correspond to high intermittencies or equivalently to the downstream locus of the separation shock wave motion. The Mach 2.0 tests, however, showed that the intermittent region occurs significantly upstream of the separation line. Power spectral densities measured in the intermittent regions were found to have self-similar frequency distributions when compared as functions of a Strouhal number for all Mach numbers and interaction sweepback angles. The maximum zero-crossing frequencies were found to correspond with the peak frequencies in the power spectra measured in the intermittent region.

The Effects of Streamwise-Deflected Wing Tips on the Aerodynamic Characteristics of an Aspect Ratio-2 Triangular Wing, Body, and Tail Combination

NASA Technical Reports Server (NTRS)

Peterson, Victor L.

1959-01-01

An investigation has been conducted on a triangular wing and body combination to determine the effects on the aerodynamic characteristics resulting from deflecting portions of the wing near the tips 900 to the wing surface about streamwise hinge lines. Experimental data were obtained for Mach numbers of 0.70, 1.30, 1.70, and 2.22 and for angles of attack ranging from -5 deg to +18 deg at sideslip angles of 0 deg and 5 deg. The results showed that the aerodynamic center shift experienced by the triangular wing and body combination as the Mach number was increased from subsonic to supersonic could be reduced by about 40 percent by deflecting the outboard 4 percent of the total area of each wing panel. Deflection about the same hinge line of additional inboard surfaces consisting of 2 percent of the total area of each wing panel resulted in a further reduction of the aerodynamic center travel of 10 percent. The resulting reductions in the stability were accompanied by increases in the drag due to lift and, for the case of the configuration with all surfaces deflected, in the minimum drag. The combined effects of reduced stability and increased drag of the untrimmed configuration on the trimmed lift-drag ratios were estimated from an analysis of the cases in which the wing-body combination with or without tips deflected was assumed to be controlled by a canard. The configurations with deflected surfaces had higher trimmed lift-drag ratios than the model with undeflected surfaces at Mach numbers up to about 1.70. Deflecting either the outboard surfaces or all of the surfaces caused the directional stability to be increased by increments that were approximately constant with increasing angle of attack at each Mach number. The effective dihedral was decreased at all angles of attack and Mach numbers when the surfaces were deflected.

Transition of the Laminar Boundary Layer on a Delta Wing with 74 degree Sweep in Free Flight at Mach Numbers from 2.8 to 5.3

NASA Technical Reports Server (NTRS)

Chapman, Gary T.

1961-01-01

The tests were conducted at Mach numbers from 2.8 to 5.3, with model surface temperatures small compared to boundary-layer recovery temperature. The effects of Mach number, temperature ratio, unit Reynolds number, leading-edge diameter, and angle of attack were investigated in an exploratory fashion. The effect of heat-transfer condition (i.e., wall temperature to total temperature ratio) and Mach number can not be separated explicitly in free-flight tests. However, the data of the present report, as well as those of NACA TN 3473, were found to be more consistent when plotted versus temperature ratio. Decreasing temperature ratio increased the transition Reynolds number. The effect of unit Reynolds number was small as was the effect of leading-edge diameter within the range tested. At small values of angle of attack, transition moved forward on the windward surface and rearward on the leeward surface. This trend was reversed at high angles of attack (6 deg to 18 deg). Possible reasons for this are the reduction of crossflow on the windward side and the influence of the lifting vortices on the leeward surface. When the transition results on the 740 delta wing were compared to data at similar test conditions for an unswept leading edge, the results bore out the results of earlier research at nearly zero heat transfer; namely, sweep causes a large reduction in the transition Reynolds number.

A study of the effects of Reynolds number and Mach number on constant pressure coefficient jump for shock-induced trailing-edge separation

NASA Technical Reports Server (NTRS)

Cunningham, Atlee M., Jr.; Spragle, Gregory S.

1987-01-01

The influence of Mach and Reynolds numbers as well as airfoil and planform geometry on the phenomenon of constant shock jump pressure coefficient for conditions of shock induced trailing edge separation (SITES) was studied. It was demonstrated that the phenomenon does exist for a wide variety of two and three dimensional flow cases and that the influence of free stream Mach number was not significant. The influence of Reynolds number was found to be important but was not strong. Airfoil and planform geometric characteristics were found to be very important where the pressure coefficient jump was shown to vary with the sum of: (1) airfoil curvature at the upper surface crest, and (2) camber surface slope at the trailing edge. It was also determined that the onset of SITES could be defined as a function of airfoil geometric parameters and Mach number normal to the leading edge. This onset prediction was shown to predict the angle of onset to within + or - 1 deg accuracy or better for about 90% of the cases studied.

Some factors influencing radiation of sound from flow interaction with edges of finite surfaces

NASA Technical Reports Server (NTRS)

Hayden, R. E.; Fox, H. L.; Chanaud, R. C.

1976-01-01

Edges of surfaces which are exposed to unsteady flow cause both strictly acoustic effects and hydrodynamic effects, in the form of generation of new hydrodynamic sources in the immediate vicinity of the edge. An analytical model is presented which develops the explicit sound-generation role of the velocity and Mach number of the eddy convection past the edge, and the importance of relative scale lengths of the turbulence, as well as the relative intensity of pressure fluctuations. The Mach number (velocity) effects show that the important paramater is the convection Mach number of the eddies. The effects of turbulence scale lengths, isotropy, and spatial density (separation) are shown to be important in determining the level and spectrum of edge sound radiated for the edge dipole mechanism. Experimental data is presented which provides support for the dipole edge noise model in terms of Mach number (velocity) scaling, parametric dependence on flow field parameter, directivity, and edge diffraction effects.

Stability and control characteristics of a three-surface advanced fighter configuration at angles of attack up to 45 deg. [conducted in the Langley 16-foot transonic tunnel

NASA Technical Reports Server (NTRS)

Henderson, W. P.; Leavitt, L. D.

1981-01-01

The tests were conducted at Mach numbers from 0.40 to 0.90, at angles of attack up to 45 deg for the lower Mach numbers, and at angles of sideslip up to 15 deg. The model variations under study included adding a canard surface and deflecting horizontal tails, ailerons, and rudders.

Surface pressure and inviscid flow field properties of the North American Rockwell delta-wing orbiter for nominal Mach number of 8, volume 2

NASA Technical Reports Server (NTRS)

Matthews, R. K.; Martindale, W. R.; Warmbrod, J. D.

1972-01-01

The results are presented of a wind tunnel test program to determine surface pressures and flow field properties on the space shuttle orbiter configuration. The tests were conducted in September 1971. Data were obtained at a nominal Mach number of 8 and a free stream unit Reynolds number of 3.7 million per foot. Angle of attack was varied from 10 to 50 deg in 10-deg increments.

Investigation of an underslung half-cone inlet with compression surface mounted outboard from fuselage at Mach numbers of 1.5, 1.8, and 2.0

NASA Technical Reports Server (NTRS)

Yeager, Richard A; Gertsma, Laurence W

1958-01-01

An investigation was conducted to determine the performance of an underslung half-cone inlet mounted on a missile forebody model with the compression surface outboard from the fuselage. The inlet was designed for shock-on-lip operation at Mach number 2.0 with 25 degree half-angle spike. The cowling was attached to the fuselage through the boundary-layer plow and served as part of the fuselage boundary-layer diverter system. The performance of the half-cone inlet was compared with that of a scoop-type inlet and a normal-wedge inlet on a maximum-thrust-minus-drag basis. The increase in pressure recovery obtained with the half-cone inlet over that obtained with the reference inlets offset the slightly higher drags observed over the Mach number range for the half-cone so that the performance of this configuration was equal to that of the other inlets at Mach number 2.0 and was slightly superior at the lower Mach numbers. For a particular configuration, a peak pressure recovery of 0.879 was obtained at Mach number 2.0, zero angle of attack, and 4-percent throat bleed; the subcritical stability was 16 percent. Use of a fuselage-mounted boundary-layer splitter plate ahead of the inlet did not improve the stability. Subcritical distortion values were below 10 percent for all configurations. (author)

Wind-tunnel/flight correlation study of aerodynamic characteristics of a large flexible supersonic cruise airplane (XB-70-1). 3: A comparison between characteristics predicted from wind-tunnel measurements and those measured in flight

NASA Technical Reports Server (NTRS)

Arnaiz, H. H.; Peterson, J. B., Jr.; Daugherty, J. C.

1980-01-01

A program was undertaken by NASA to evaluate the accuracy of a method for predicting the aerodynamic characteristics of large supersonic cruise airplanes. This program compared predicted and flight-measured lift, drag, angle of attack, and control surface deflection for the XB-70-1 airplane for 14 flight conditions with a Mach number range from 0.76 to 2.56. The predictions were derived from the wind-tunnel test data of a 0.03-scale model of the XB-70-1 airplane fabricated to represent the aeroelastically deformed shape at a 2.5 Mach number cruise condition. Corrections for shape variations at the other Mach numbers were included in the prediction. For most cases, differences between predicted and measured values were within the accuracy of the comparison. However, there were significant differences at transonic Mach numbers. At a Mach number of 1.06 differences were as large as 27 percent in the drag coefficients and 20 deg in the elevator deflections. A brief analysis indicated that a significant part of the difference between drag coefficients was due to the incorrect prediction of the control surface deflection required to trim the airplane.

Wind-tunnel investigation of aerodynamic loading on a 0.237-scale model of a remotely piloted research vehicle with a thick, high-aspect-ratio supercritical wing

NASA Technical Reports Server (NTRS)

Byrdsong, T. A.; Brooks, C. W., Jr.

1983-01-01

Wind-tunnel measurements were made of the wing-surface static-pressure distributions on a 0.237 scale model of a remotely piloted research vehicle equipped with a thick, high-aspect-ratio supercritical wing. Data are presented for two model configurations (with and without a ventral pod) at Mach numbers from 0.70 to 0.92 at angles of attack from -4 deg to 8 deg. Large variations of wing-surface local pressure distributions were developed; however, the characteristic supercritical-wing pressure distribution occurred near the design condition of 0.80 Mach number and 2 deg angle of attack. The significant variations of the local pressure distributions indicated pronounced shock-wave movements that were highly sensitive to angle of attack and Mach number. The effect of the vertical pod varied with test conditions; however at the higher Mach numbers, the effects on wing flow characteristics were significant at semispan stations as far outboard as 0.815. There were large variations of the wing loading in the range of test conditions, both model configurations exhibited a well-defined peak value of normal-force coefficient at the cruise angle of attack (2 deg) and Mach number (0.80).

Periodic vortex shedding in the supersonic wake of a planar plate

NASA Technical Reports Server (NTRS)

Xing, W. F.; Marenbach, G.

1985-01-01

Vortex sheets in the wake have been mainly studied in incompressible flows and in the transonic region. Heinemann et al. (1976) have shown that for the subsonic region the Strouhal number is nearly independent of the Mach number. Motallebi and Norbury (1981) have observed an increase in the Strouhal number in transonic supersonic flow at Mach numbers up to 1.25. The present investigation is concerned with an extension of the studies of vortex shedding to higher supersonic Mach numbers, taking into account questions regarding the possibility of a generation of stable von Karman vortex paths in the considered Mach number range. It is found that the vortex sheet observed in a supersonic wake behind a rough plate is only stable and reproducible in cases involving a certain surface roughness and certain aspects of trailing edge geometry.

Aerodynamic performance of transonic and subsonic airfoils: Effects of surface roughness, turbulence intensity, Mach number, and streamline curvature-airfoil shape

NASA Astrophysics Data System (ADS)

Zhang, Qiang

The effects of surface roughness, turbulence intensity, Mach number, and streamline curvature-airfoil shape on the aerodynamic performance of turbine airfoils are investigated in compressible, high speed flows. The University of Utah Transonic Wind Tunnel is employed for the experimental part of the study. Two different test sections are designed to produce Mach numbers, Reynolds numbers, passage mass flow rates, and physical dimensions, which match values along turbine blades in operating engines: (i) a nonturning test section with a symmetric airfoil, and (ii) a cascade test section with a cambered turbine vane. The nonuniform, irregular, three-dimensional surface roughness is characterized using the equivalent sand grain roughness size. Changing the airfoil surface roughness condition has a substantial effect on wake profiles of total pressure loss coefficients, normalized Mach number, normalized kinetic energy, and on the normalized and dimensional magnitudes of Integrated Aerodynamic Losses produced by the airfoils. Comparisons with results for a symmetric airfoil and a cambered vane show that roughness has more substantial effects on losses produced by the symmetric airfoil than the cambered vane. Data are also provided that illustrate the larger loss magnitudes are generally present with flow turning and cambered airfoils, than with symmetric airfoils. Wake turbulence structure of symmetric airfoils and cambered vanes are also studied experimentally. The effects of surface roughness and freestream turbulence levels on wake distributions of mean velocity, turbulence intensity, and power spectral density profiles and vortex shedding frequencies are quantified one axial chord length downstream of the test airfoils. As the level of surface roughness increases, all wake profile quantities broaden significantly and nondimensional vortex shedding frequencies decrease. Wake profiles produced by the symmetric airfoil are more sensitive to variations of surface roughness and freestream turbulence, compared with data from the cambered vane airfoil. Stanton numbers, skin friction coefficients, aerodynamic losses, and Reynolds analogy behavior are numerically predicted for a turbine vane using the FLUENT with a k-epsilon RNG model to show the effects of Mach number, mainstream turbulence level, and surface roughness. Comparisons with wake aerodynamic loss experimental data are made. Numerically predicted skin friction coefficients and Stanton numbers are also used to deduce Reynolds analogy behavior on the vane suction and pressure sides.

Aerodynamics of a Transitioning Turbine Stator Over a Range of Reynolds Numbers

NASA Technical Reports Server (NTRS)

Boyle, R. J.; Lucci, B. L.; Verhoff, V. G.; Camperchioli, W. P.; La, H.

1998-01-01

Midspan aerodynamic measurements for a three vane-four passage linear turbine vane cascade are given. The vane axial chord was 4.45 cm. Surface pressures and loss coefficients were measured at exit Mach numbers of 0.3, 0.7, and 0.9. Reynolds number was varied by a factor of six at the two highest Mach numbers, and by a factor of ten at the lowest Mach number. Measurements were made with and without a turbulence grid. Inlet turbulence intensities were less than I% and greater than IO%. Length scales were also measured. Pressurized air fed the test section, and exited to a low pressure exhaust system. Maximum inlet pressure was two atmospheres. The minimum inlet pressure for an exit Mach number of 0.9 was one-third of an atmosphere, and at a Mach number of 0.3, the minimum pressure was half this value. The purpose of the test was to provide data for verification of turbine vane aerodynamic analyses, especially at low Reynolds numbers. Predictions obtained using a Navier-Stokes analysis with an algebraic turbulence model are also given.

Surface pressure and inviscid flow field properties of the McDonnell-Douglas delta-wing orbiter for nominal Mach number of 8, Volume 1

NASA Technical Reports Server (NTRS)

Warmbrod, J. D.; Martindale, M. R.; Matthews, R. K.

1972-01-01

The results of a wind tunnel test program to determine the surface pressures and flow distribution on the McDonnell Douglas Orbiter configuration are presented. Tests were conducted in hypersonic wind tunnel at Mach 8. The freestream unit Reynolds number was 3.7 time one million per foot. Angle of attack was varied from 10 degrees to 60 degrees in 10 degree increments.

Dynamic Stall Data for 2-D and 3-D Test Cases

DTIC Science & Technology

2000-10-01

CASES Professor R A McD Galbraith Dr F N Coton Dr R B Green Dr M Vezza University of Glasgow INTRODUCTION Background Although substantial work has...aerofoil shape, aspect ratio, surface finish , data reduction software and Mach number, all but the Mach number had no effect on the observed trends

Surface pressure fluctuations in hypersonic turbulent boundary layers

NASA Technical Reports Server (NTRS)

Raman, K. R.

1974-01-01

The surface pressure fluctuations on a flat plate model at hypersonic Mach numbers of 5.2, 7.4 and 10.4 with an attached turbulent boundary layer were measured using flush mounted small piezoelectric sensors. A high frequency resolution of the pressure field was achieved using specially designed small piezoelectric sensors that had a good frequency response well above 300 KHz. The RMS pressures and non-dimensional energy spectra for all above Mach numbers are presented. The convective velocities, obtained from space time correlation considerations are equal to 0.7 U sub infinity. The results indicate the RMS pressures vary from 5 to 25 percent of the mean static pressures. The ratios of RMS pressure to dynamic pressure are less than the universally accepted subsonic value of 6 x 10/3. The ratio decreases in value as the Mach number or the dynamic pressure is increased. The ratio of RMS pressure to wall shear for Mach number 7.4 satisfies one smaller than or equal to p/tau sub w smaller than or equal to three.

The NASA Langley Laminar-Flow-Control Experiment on a Swept Supercritical Airfoil: Basic Results for Slotted Configuration

NASA Technical Reports Server (NTRS)

Harris, Charles D.; Brooks, Cuyler W., Jr.; Clukey, Patricia G.; Stack, John P.

1989-01-01

The effects of Mach number and Reynolds number on the experimental surface pressure distributions and transition patterns for a large chord, swept supercritical airfoil incorporating an active Laminar Flow Control suction system with spanwise slots are presented. The experiment was conducted in the Langley 8 foot Transonic Pressure Tunnel. Also included is a discussion of the influence of model/tunnel liner interactions on the airfoil pressure distribution. Mach number was varied from 0.40 to 0.82 at two chord Reynolds numbers, 10 and 20 x 1,000,000, and Reynolds number was varied from 10 to 20 x 1,000,000 at the design Mach number.

Pressure and heating-rate distributions on a corrugated surface in a supersonic turbulent boundary layer

NASA Technical Reports Server (NTRS)

Sawyer, J. W.

1977-01-01

Drag and heating rates on wavy surfaces typical of current corrugated plate designs for thermal protection systems were determined experimentally. Pressure-distribution, heating-rate, and oil-flow tests were conducted in the Langley Unitary Plan wind tunnel at Mach numbers of 2.4 and 4.5 with the corrugated surface exposed to both thick and thin turbulent boundary layers. Tests were conducted with the corrugations at cross-flow angles from 0 deg to 90 deg to the flow. Results show that for cross-flow angles of 30 deg or less, the pressure drag coefficients are less than the local flat-plate skin-friction coefficients and are not significantly affected by Mach number, Reynolds number, or boundary-layer thickness over the ranges investigated. For cross-flow angles greater than 30 deg, the drag coefficients increase significantly with cross-flow angle and moderately with Reynolds number. Increasing the Mach number causes a significant reduction in the pressure drag. The average and peak heating penalties due to the corrugated surface are small for cross-flow angles of 10 deg or less but are significantly higher for the larger cross-flow angles.

Flight Wing Surface Pressure and Boundary-Layer Data Report from the F-111 Smooth Variable-Camber Supercritical Mission Adaptive Wing

NASA Technical Reports Server (NTRS)

Powers, Sheryll Goecke; Webb, Lannie D.

1997-01-01

Flight tests were conducted using the advanced fighter technology integration F-111 (AFTI/F-111) aircraft modified with a variable-sweep supercritical mission adaptive wing (MAW). The MAW leading- and trailing-edge variable-camber surfaces were deflected in flight to provide a near-ideal wing camber shape for the flight condition. The MAW features smooth, flexible upper surfaces and fully enclosed lower surfaces, which distinguishes it from conventional flaps that have discontinuous surfaces and exposed or semi-exposed mechanisms. Upper and lower surface wing pressure distributions were measured along four streamwise rows on the right wing for cruise, maneuvering, and landing configurations. Boundary-layer measurements were obtained near the trailing edge for one of the rows. Cruise and maneuvering wing leading-edge sweeps were 26 deg for Mach numbers less than 1 and 45 deg or 58 deg for Mach numbers greater than 1. The landing wing sweep was 9 deg or 16 deg. Mach numbers ranged from 0.27 to 1.41, angles of attack from 2 deg to 13 deg, and Reynolds number per unit foot from 1.4 x 10(exp 6) to 6.5 x 10(exp 6). Leading-edge cambers ranged from O deg to 20 deg down, and trailing-edge cambers ranged from 1 deg up to 19 deg down. Wing deflection data for a Mach number of 0.85 are shown for three cambers. Wing pressure and boundary-layer data are given. Selected data comparisons are shown. Measured wing coordinates are given for three streamwise semispan locations for cruise camber and one spanwise location for maneuver camber.

Experimental Investigation of Roughness Effects on Transition Onset and Turbulent Heating Augmentation on a Hemisphere at Mach 6 and Mach 10

NASA Technical Reports Server (NTRS)

Hollis, Brian R.

2017-01-01

An experimental investigation of the effects of distributed surface roughness on boundary-layer transition and turbulent heating has been conducted. Hypersonic wind tunnel testing was performed using hemispherical models with surface roughness patterns simulating those produced by heat shield ablation. Global aeroheating and transition onset data were obtained using phosphor thermography at Mach 6 and Mach 10 over a range of roughness heights and free stream Reynolds numbers sufficient to produce laminar, transitional and turbulent flow. Upstream movement of the transition onset location and increasing heating augmentation over predicted smooth-wall levels were observed with both increasing roughness heights and increasing free stream Reynolds numbers. The experimental heating data are presented herein, as are comparisons to smooth-wall heat transfer distributions from computational flow-field simulations. The transition onset data are also tabulated, and correlations of these data are presented.

High-Speed Wind-Tunnel Tests of a Model of the Lockheed YP-80A Airplane Including Correlation with Flight Tests and Tests of Dive-Recovery Flaps

NASA Technical Reports Server (NTRS)

Cleary, Joseph W.; Gray, Lyle J.

1947-01-01

This report contains the results of tests of a 1/3-scale model of the Lockheed YP-90A "Shooting Star" airplane and a comparison of drag, maximum lift coefficient, and elevator angle required for level flight as measured in the wind tunnel and in flight. Included in the report are the general aerodynamic characteristics of the model and of two types of dive-recovery flaps, one at several positions along the chord on the lower surface of the wing and the other on the lower surface of the fuselage. The results show good agreement between the flight and wind-tunnel measurements at all Mach numbers. The results indicate that the YP-80A is controllable in pitch by the elevators to a Mach number of at least 0.85. The fuselage dive-recovery flaps are effective for producing a climbing moment and increasing the drag at Mach numbers up to at least 0.8. The wing dive-recovery flaps are most effective for producing a climbing moment at 0.75 Mach number. At 0.85 Mach number, their effectiveness is approximately 50 percent of the maximum. The optimum position for the wing dive-recovery flaps to produce a climbing moment is at approximately 35 percent of the chord.

Wind tunnel investigation of three axisymmetric cowls of different lengths at Mach numbers from 0.60 to 0.92

NASA Technical Reports Server (NTRS)

Re, Richard J.; Abeyounis, William K.

1993-01-01

Pressure distributions on three inlets having different cowl lengths were obtained in the Langley 16-Foot Transonic Tunnel. The cowl diameter ratio (highlight diameter to maximum diameter) was 0.85 and the cowl length ratios (cowl length to maximum diameter) were 0.337, 0.439, and 0.547. The cowls had identical nondimensionalized (with respect to cowl length) external geometry and identical internal geometry. The internal contraction ratio (highlight area to throat area) was 1.250. The inlets had longitudinal rows of static pressure orifices on the top and bottom (external) surfaces and on the contraction (internal) and diffuser surfaces. The afterbody was cylindrical in shape, and its diameter was equal to the maximum diameter of the cowl. Depending on the cowl configuration and free-stream Mach number, the mass-flow ratio varied between 0.27 and 0.87 during the tests. Angle of attack varied from 0 to 4.1 deg at selected Mach numbers and mass-flow ratios, and the Reynolds number varied with the Mach number from 3.2x10(exp 6) to 4.2x10(exp 6) per foot.

Effect of geometry variations on lee-surface vortex-induced heating for flat-bottom three-dimensional bodies at Mach 6

NASA Technical Reports Server (NTRS)

Hefner, J. N.

1973-01-01

Studies have shown that vortices can produce relatively severe heating on the leeward surfaces of conceptual hypersonic vehicles and that surface geometry can strongly influence this vortex-induced heating. Results which show the effects of systematic geometry variations on the vortex-induced lee-surface heating on simple flat-bottom three-dimensional bodies at angles of attack of 20 deg and 40 deg are presented. The tests were conducted at a free-stream Mach number of 6 and at a Reynolds number of 1.71 x 10 to the 7th power per meter.

Effect of Tail Surfaces on the Base Drag of a Body of Revolution at Mach Numbers of 1.5 and 2.0

NASA Technical Reports Server (NTRS)

Spahr, J Richard; Dickey, Robert R

1951-01-01

Wind-tunnel tests were performed at Mach numbers of 1.5 and 2.0 to investigate the influence of tail surfaces on the base drag of a body of revolution without boattailing and having a turbulent boundary layer. The tail surfaces were of rectangular plan form of aspect ratio 2.33 and has symmetrical, circular-arc airfoil section. The results of the investigation showed that the addition of these tail surfaces with the trailing edges at or near the body base incurred a large increase in the base-drag coefficient. For a cruciform tail having a 10-percent-thick airfoil section, this increase was about 70 percent at a Mach number of 1.5 and 35 percent at a Mach number of 2.0. As the trailing edge of the tail was moved forward or rearward of the base by about one tail-chord length, the base-drag increment was reduced to nearly zero. The increments in base-drag coefficient due to the presence of 10-percent-thick tail surfaces were generally twice those for 5-percent-thick surfaces. The base-drag increments due to the presence of a cruciform tail were less than twice those for a plane tail. An estimate of the change in base pressure due to the tail surfaces was made, based on a simple superposition of the airfoil-pressure field onto the base-pressure field behind the body. A comparison of the results with the experimental values indicated that in most cases the trend in the variation of the base-drag increment with changes in tail position could be predicted by this approximate method but that the quantitative agreement at most tail locations was poor.

Unsteady blade-surface pressures on a large-scale advanced propeller: Prediction and data

NASA Technical Reports Server (NTRS)

Nallasamy, M.; Groeneweg, J. F.

1990-01-01

An unsteady 3-D Euler analysis technique is employed to compute the flow field of an advanced propeller operating at an angle of attack. The predicted blade pressure waveforms are compared with wind tunnel data at two Mach numbers, 0.5 and 0.2. The inflow angle is three degrees. For an inflow Mach number of 0.5, the predicted pressure response is in fair agreement with data: the predicted phases of the waveforms are in close agreement with data while the magnitudes are underpredicted. At the low Mach number of 0.2 (takeoff), the numerical solution shows the formation of a leading edge vortex which is in qualitative agreement with measurements. However, the highly nonlinear pressure response measured on the blade suction surface is not captured in the present inviscid analysis.

Unsteady blade surface pressures on a large-scale advanced propeller - Prediction and data

NASA Technical Reports Server (NTRS)

Nallasamy, M.; Groeneweg, J. F.

1990-01-01

An unsteady three dimensional Euler analysis technique is employed to compute the flowfield of an advanced propeller operating at an angle of attack. The predicted blade pressure waveforms are compared with wind tunnel data at two Mach numbers, 0.5 and 0.2. The inflow angle is three degrees. For an inflow Mach number of 0.5, the predicted pressure response is in fair agreement with data: the predicted phases of the waveforms are in close agreement with data while the magnitudes are underpredicted. At the low Mach number of 0.2 (take-off) the numerical solution shows the formation of a leading edge vortex which is in qualitative agreement with measurements. However, the highly nonlinear pressure response measured on the blade suction surface is not captured in the present inviscid analysis.

Aerodynamic characteristics of wings designed with a combined-theory method to cruise at a Mach number of 4.5

NASA Technical Reports Server (NTRS)

Mack, Robert J.

1988-01-01

A wind-tunnel study was conducted to determine the capability of a method combining linear theory and shock-expansion theory to design optimum camber surfaces for wings that will fly at high-supersonic/low-hypersonic speeds. Three force models (a flat-plate reference wing and two cambered and twisted wings) were used to obtain aerodynamic lift, drag, and pitching-moment data. A fourth pressure-orifice model was used to obtain surface-pressure data. All four wing models had the same planform, airfoil section, and centerbody area distribution. The design Mach number was 4.5, but data were also obtained at Mach numbers of 3.5 and 4.0. Results of these tests indicated that the use of airfoil thickness as a theoretical optimum, camber-surface design constraint did not improve the aerodynamic efficiency or performance of a wing as compared with a wing that was designed with a zero-thickness airfoil (linear-theory) constraint.

Heat transfer characteristics of hypersonic waveriders with an emphasis on the leading edge effects. M.S. Thesis, 1991

NASA Technical Reports Server (NTRS)

Vanmol, Denis O.; Anderson, John D., Jr.

1992-01-01

The heat transfer characteristics in surface radiative equilibrium and the aerodynamic performance of blunted hypersonic waveriders are studied along two constant dynamic pressure trajectories for four different Mach numbers. The inviscid leading edge drag was found to be a small (4 to 8 percent) but not negligible fraction of the inviscid drag of the vehicle. Although the viscous drag at the leading edge can be neglected, the presence of the leading edge will influence the transition pattern of the upper and the lower surfaces and therefore affect the viscous drag of the entire vehicle. For an application similar to the National Aerospace Plane (NASP), the present study demonstrates that the waverider remains a valuable concept at high Mach numbers if a state-of-the-art active cooling device is used along the leading edge. At low Mach number (less than 5), the study shows the surface radiative cooling might be sufficient. In all cases, radiative cooling is sufficient for the upper and lower surfaces of the vehicle if ceramic composites are used as thermal protection.

Surface pressure data for a supersonic-cruise airplane configuration at Mach numbers of 2.30, 2.96, 3.30

NASA Technical Reports Server (NTRS)

Shrout, B. L.; Corlett, W. A.; Collins, I. K.

1979-01-01

The tabulated results of surface pressure tests conducted on the wing and fuselage of an airplane model in the Langley Unitary Plan wind tunnel are presented without analysis. The model tested was that of a supersonic-cruise airplane with a highly swept arrow-wing planform, two engine nacelles mounted beneath the wing, and outboard vertical tails. Data were obtained at Mach numbers of 2.30, 2.96, and 3.30 for angles of attack from -4 deg to 12 deg. The Reynolds number for these tests was 6,560,000 per meter.

X-33 Computational Aeroheating/Aerodynamic Predictions and Comparisons With Experimental Data

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Thompson, Richard A.; Berry, Scott A.; Horvath, Thomas J.; Murphy, Kelly J.; Nowak, Robert J.; Alter, Stephen J.

2003-01-01

This report details a computational fluid dynamics study conducted in support of the phase II development of the X-33 vehicle. Aerodynamic and aeroheating predictions were generated for the X-33 vehicle at both flight and wind-tunnel test conditions using two finite-volume, Navier-Stokes solvers. Aerodynamic computations were performed at Mach 6 and Mach 10 wind-tunnel conditions for angles of attack from 10 to 50 with body-flap deflections of 0 to 20. Additional aerodynamic computations were performed over a parametric range of free-stream conditions at Mach numbers of 4 to 10 and angles of attack from 10 to 50. Laminar and turbulent wind-tunnel aeroheating computations were performed at Mach 6 for angles of attack of 20 to 40 with body-flap deflections of 0 to 20. Aeroheating computations were performed at four flight conditions with Mach numbers of 6.6 to 8.9 and angles of attack of 10 to 40. Surface heating and pressure distributions, surface streamlines, flow field information, and aerodynamic coefficients from these computations are presented, and comparisons are made with wind-tunnel data.

F-16XL Wing Pressure Distributions and Shock Fence Results from Mach 1.4 to Mach 2.0

NASA Technical Reports Server (NTRS)

Landers, Stephen F.; Saltzman, John A.; Bjarke, Lisa J.

1997-01-01

Chordwise pressure distributions were obtained in-flight on the upper and lower surfaces of the F-16XL ship 2 aircraft wing between Mach 1.4 and Mach 2.0. This experiment was conducted to determine the location of shock waves which could compromise or invalidate a follow-on test of a large chord laminar flow control suction panel. On the upper surface, the canopy closure shock crossed an area which would be covered by a proposed laminar flow suction panel. At the laminar flow experiment design Mach number of 1.9, 91 percent of the suction panel area would be forward of the shock. At Mach 1.4, that value reduces to 65 percent. On the lower surface, a shock from the inlet diverter would impinge on the proposed suction panel leading edge. A chordwise plate mounted vertically to deflect shock waves, called a shock fence, was installed between the inlet diverter and the leading edge. This plate was effective in reducing the pressure gradients caused by the inlet shock system.

An experimental/computational study of sharp fin induced shock wave/turbulent boundary layer interactions at Mach 5 - Experimental results

NASA Technical Reports Server (NTRS)

Rodi, Patrick E.; Dolling, David S.

1992-01-01

A combined experimental/computational study has been performed of sharp fin induced shock wave/turbulent boundary layer interactions at Mach 5. The current paper focuses on the experiments and analysis of the results. The experimental data include mean surface heat transfer, mean surface pressure distributions and surface flow visualization for fin angles of attack of 6, 8, 10, 12, 14 and 16-degrees at Mach 5 under a moderately cooled wall condition. Comparisons between the results and correlations developed earlier show that Scuderi's correlation for the upstream influence angle (recast in a conical form) is superior to other such correlations in predicting the current results, that normal Mach number based correlations for peak pressure heat transfer are adequate and that the initial heat transfer peak can be predicted using pressure-interaction theory.

Aerodynamic Characteristics of a Revised Target Drone Vehicle at Mach Numbers from 1.60 to 2.86

NASA Technical Reports Server (NTRS)

Blair, A. B., Jr.; Babb, C. Donald

1968-01-01

An investigation has been conducted in the Langley Unitary Plan wind tunnel to determine the aerodynamic characteristics of a revised target drone vehicle through a Mach number range from 1.60 to 2.86. The vehicle had canard surfaces and a swept clipped-delta wing with twin tip-mounted vertical tails.

Flight Investigation at High Speeds of Profile Drag of Wing of a P-47D Airplane Having Production Surfaces Covered with Camouflage Paint

NASA Technical Reports Server (NTRS)

Daum, Fred L.; Zalovcik, John A.

1946-01-01

Wing section outboard of flap was tested by wake surveys in Mach range of 0.25 - 0.78 and lift coefficient range 0.06 - 0.69. Results indicated that minimum profile-drag coefficient of 0.0097 was attained for lift coefficients from 0.16 to 0.25 at Mach less than 0.67. Below Mach number at which compressibility shock occurred, variations in Mach of 0.2 had negligible effect on profile drag coefficient. Shock was not evident until critical Mach was exceeded by 0.025.

Aerothermodynamic Analysis of Commercial Experiment Transporter (COMET) Reentry Capsule

NASA Technical Reports Server (NTRS)

Wood, William A.; Gnoffo, Peter A.; Rault, Didier F. G.

1996-01-01

An aerothermodynamic analysis of the Commercial Experiment Transporter (COMET) reentry capsule has been performed using the laminar thin-layer Navier-Stokes solver Langley Aerothermodynamic Upwind Relaxation Algorithm. Flowfield solutions were obtained at Mach numbers 1.5, 2, 5, 10, 15, 20, 25, and 27.5. Axisymmetric and 5, 10, and 20 degree angles of attack were considered across the Mach-number range, with the Mach 25 conditions taken to 90 degrees angle of attack and the Mach 27.5 cases taken to 60 degrees angle of attack. Detailed surface heat-transfer rates were computed at Mach 20 and 25, revealing that heating rates on the heat-shield shoulder ,can exceed the stagnation-point heating by 230 percent. Finite-rate chemistry solutions were performed above Mach 10, otherwise perfect gas computations were made. Drag, lift, and pitching moment coefficients are computed and details of a wake flow are presented. The effect of including the wake in the solution domain was investigated and base pressure corrections to forebody drag coefficients were numerically determined for the lower Mach numbers. Pitching moment comparisons are made with direct simulation Monte Carlo results in the more rarefied flow at the highest Mach numbers, showing agreement within two-percent. Thin-layer Navier-Stokes computations of the axial force are found to be 15 percent higher across the speed range than the empirical/Newtonian based results used during the initial trajectory analyses.

A simple analytical aerodynamic model of Langley Winged-Cone Aerospace Plane concept

NASA Technical Reports Server (NTRS)

Pamadi, Bandu N.

1994-01-01

A simple three DOF analytical aerodynamic model of the Langley Winged-Coned Aerospace Plane concept is presented in a form suitable for simulation, trajectory optimization, and guidance and control studies. The analytical model is especially suitable for methods based on variational calculus. Analytical expressions are presented for lift, drag, and pitching moment coefficients from subsonic to hypersonic Mach numbers and angles of attack up to +/- 20 deg. This analytical model has break points at Mach numbers of 1.0, 1.4, 4.0, and 6.0. Across these Mach number break points, the lift, drag, and pitching moment coefficients are made continuous but their derivatives are not. There are no break points in angle of attack. The effect of control surface deflection is not considered. The present analytical model compares well with the APAS calculations and wind tunnel test data for most angles of attack and Mach numbers.

Turbulent-Spot Growth Characteristics: Wind-Tunnel and Flight Measurements of Natural Transition at High Reynolds and Mach Numbers

NASA Technical Reports Server (NTRS)

Clark, J. P.; Jones, T. V.; LaGraff, J. E.

2007-01-01

A series of experiments are described which examine the growth of turbulent spots on a flat plate at Reynolds and Mach numbers typical of gas-turbine blading. A short-duration piston tunnel is employed and rapid-response miniature surface-heat-transfer gauges are used to asses the state of the boundary layer. The leading- and trailing-edge velocities of spots are reported for different external pressure gradients and Mach numbers. Also, the lateral spreading angle is determined from the heat-transfer signals which demonstrate dramatically the reduction in spot growth associated with favorable pressure gradients. An associated experiment on the development of turbulent wedges is also reported where liquid-crystal heat-transfer techniques are employed in low-speed wind tunnel to visualize and measure the wedge characteristics. Finally, both liquid crystal techniques and hot-film measurements from flight tests at Mach number of 0.6 are presented.

Measurements of surface-pressure fluctuations on the XB-70 airplane at local Mach numbers up to 2.45

NASA Technical Reports Server (NTRS)

Lewis, T. L.; Dods, J. B., Jr.; Hanly, R. D.

1973-01-01

Measurements of surface-pressure fluctuations were made at two locations on the XB-70 airplane for nine flight-test conditions encompassing a local Mach number range from 0.35 to 2.45. These measurements are presented in the form of estimated power spectral densities, coherence functions, and narrow-band-convection velocities. The estimated power spectral densities compared favorably with wind-tunnel data obtained by other experimenters. The coherence function and convection velocity data supported conclusions by other experimenters that low-frequency surface-pressure fluctuations consist of small-scale turbulence components with low convection velocity.

Heat transfer distributions induced by elevon deflections on swept wings and adjacent surfaces at Mach 6

NASA Technical Reports Server (NTRS)

Johnson, C. B.; Kaufman, L. G., II

1978-01-01

Surface heat transfer distributions are presented for swept wing semispan models having trailing edge elevon ramp angles of 0, 10, 20, and 30 degrees. The wing sweepback angles are 0, 50, and 70 degrees. The models have attachable cylindrical and flat plate center bodies and various attachable wing-tip fins. The data, obtained for a 0 degree angle of attack, a free stream Mach number of 6, and a wing root chord Reynolds number of about 17,000,000, reveal considerably larger regions of elevon induced thermal loads on adjacent surfaces than would be suggested by fully attached flow analyses.

In-flight boundary-layer measurements on a hollow cylinder at a Mach number of 3.0

NASA Technical Reports Server (NTRS)

Quinn, R. D.; Gong, L.

1980-01-01

Skin temperatures, shear forces, surface static pressures, boundary layer pitot pressures, and boundary layer total temperatures were measured on the external surface of a hollow cylinder that was 3.04 meters long and 0.437 meter in diameter and was mounted beneath the fuselage of the YF-12A airplane. The data were obtained at a nominal free stream Mach number of 3.0 (a local Mach number of 2.9) and at wall to recovery temperature ratios of 0.66 to 0.91. The local Reynolds number had a nominal value of 4,300,000 per meter. Heat transfer coefficients and skin friction coefficients were derived from skin temperature time histories and shear force measurements, respectively. In addition, boundary layer velocity profiles were derived from pitot pressure measurements, and a Reynolds analogy factor was obtained from the heat transfer and skin friction measurements. The measured data are compared with several boundary layer prediction methods.

Assessment of Potential Aerodynamic Benefits from Spanwise Blowing at the Wing Tip. Ph.D. Thesis - George Washington Univ.

NASA Technical Reports Server (NTRS)

Mineck, Raymond Edward

1992-01-01

A comprehensive set of experimental and analytical investigations have been conducted to assess the potential aerodynamic benefits from spanwise blowing at the tip of a moderate aspect ratio, swept wing. An analytical model has been developed to simulate a jet exhausting from the wing tip. An experimental study of a subsonic jet exhausting from the wing tip was conducted to investigate the effect of spanwise blowing from the tip on the aerodynamic characteristics of a moderate aspect ratio, swept wing. Wing force and moment data and surface pressure data were measured at Mach numbers up to 0.72. Results indicate that small amounts of blowing from small jets increase the lift curve slope a small amount, but have no effect on drag. Larger amounts of blowing from longer jets blowing increases lift near the tip and reduce drag at low Mach numbers. These benefits decrease with increasing Mach number, and vanish at Mach 0.5. A Navier-Stokes solver with modified boundary conditions at the tip was used to extrapolate the results to a Mach number of 0.72. With current technology and conventional wing shapes, spanwise blowing at the wing tip does not appear to be a practical means of reducing drag of moderate aspect ratio wings at high subsonic Mach numbers.

Assessment of CFD Hypersonic Turbulent Heating Rates for Space Shuttle Orbiter

NASA Technical Reports Server (NTRS)

Wood, William A.; Oliver, A. Brandon

2011-01-01

Turbulent CFD codes are assessed for the prediction of convective heat transfer rates at turbulent, hypersonic conditions. Algebraic turbulence models are used within the DPLR and LAURA CFD codes. The benchmark heat transfer rates are derived from thermocouple measurements of the Space Shuttle orbiter Discovery windward tiles during the STS-119 and STS-128 entries. The thermocouples were located underneath the reaction-cured glass coating on the thermal protection tiles. Boundary layer transition flight experiments conducted during both of those entries promoted turbulent flow at unusually high Mach numbers, with the present analysis considering Mach 10{15. Similar prior comparisons of CFD predictions directly to the flight temperature measurements were unsatisfactory, showing diverging trends between prediction and measurement for Mach numbers greater than 11. In the prior work, surface temperatures and convective heat transfer rates had been assumed to be in radiative equilibrium. The present work employs a one-dimensional time-accurate conduction analysis to relate measured temperatures to surface heat transfer rates, removing heat soak lag from the flight data, in order to better assess the predictive accuracy of the numerical models. The turbulent CFD shows good agreement for turbulent fuselage flow up to Mach 13. But on the wing in the wake of the boundary layer trip, the inclusion of tile conduction effects does not explain the prior observed discrepancy in trends between simulation and experiment; the flight heat transfer measurements are roughly constant over Mach 11-15, versus an increasing trend with Mach number from the CFD.

Acoustic Radiation From a Mach 14 Turbulent Boundary Layer

NASA Technical Reports Server (NTRS)

Zhang, Chao; Duan, Lian; Choudhari, Meelan M.

2016-01-01

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.

Preliminary results of unsteady blade surface pressure measurements for the SR-3 propeller

NASA Technical Reports Server (NTRS)

Heidelberg, L. J.; Clark, B. J.

1986-01-01

Unsteady blade surface pressures were measured on an advanced, highly swept propeller known as SR-3. These measurements were obtained because the unsteady aerodynamics of these highly loaded transonic blades is important to noise generation and aeroelastic response. Specifically, the response to periodic angle-of-attack change was measured for both two- and eight-bladed configurations over a range of flight Mach numbers from 0.4 to 0.85. The periodic angle-of-attack change was obtained by placing the propeller axis at angles up to 4 deg to the flow. Most of the results are presented in terms of the unsteady pressure coefficient variation with Mach number. Both cascade and Mach number effects were largest on the suction surface near the leading edge. The results of a three-dimensional Euler code applied in a quasi-steady fashion were compared to measured data at the reduced frequency of 0.1 and showed relatively poor agreement. Pressure waveforms are shown that suggest shock phenomena may play an important part in the unsteady pressure response at some blade locations.

Two-dimensional converging-diverging rippled nozzles at transonic speeds. [performed in the Langley 16-Foot Transonic Tunnel

NASA Technical Reports Server (NTRS)

Carlson, John R.; Asbury, Scott C.

1994-01-01

An experimental investigation was performed in the Langley 16-Foot Transonic tunnel to determine the effects of external and internal flap rippling on the aerodynamics of a nonaxisymmetric nozzle. Data were obtained at several Mach numbers from static conditions to 1.2 over a range of nozzle pressure ratios. Nozzles with chordal boattail angles of 10, 20, and 30 degrees, with and without surface rippling, were tested. No effect on discharge coefficient due to surface rippling was observed. Internal thrust losses due to surface rippling were measured and attributed to a combination of additional internal skin friction and shock losses. External nozzle drag for the baseline configurations were generally less than that for the rippled configurations at all free-stream Mach numbers tested. The difference between the baseline and rippled nozzle drag levels generally increased with increasing boat tail angle. The thrust-minus-drag level for each rippled nozzle configuration was less than the equivalent baseline configuration for each Mach number at the design nozzle pressure ratio.

Computer program for calculating aerodynamic characteristics of upper-surface-blowing and over-wing-blowing configurations

NASA Technical Reports Server (NTRS)

Lan, C. E.; Fillman, G. L.; Fox, C. H., Jr.

1977-01-01

The program is based on the inviscid wing-jet interaction theory of Lan and Campbell, and the jet entrainment theory of Lan. In the interaction theory, the flow perturbations are computed both inside and outside the jet, separately, and then matched on the jet surface to satisfy the jet boundary conditions. The jet Mach number is allowed to be different from the free stream value (Mach number nonuniformity). These jet boundary conditions require that the static pressure be continuous across the jet surface which must always remain as a stream surface. These conditions, as well as the wing-surface tangency condition, are satisified only in the linearized sense. The detailed formulation of these boundary conditions is based on the quasi-vortex-lattice method of Lan.

Comparison of analytical and experimental steadyand unsteady-pressure distributions at Mach number 0.78 for a high-aspect-ratio supercritical wing model with oscillating control surfaces

NASA Technical Reports Server (NTRS)

Mccain, W. E.

1984-01-01

The unsteady aerodynamic lifting surface theory, the Doublet Lattice method, with experimental steady and unsteady pressure measurements of a high aspect ratio supercritical wing model at a Mach number of 0.78 were compared. The steady pressure data comparisons were made for incremental changes in angle of attack and control surface deflection. The unsteady pressure data comparisons were made at set angle of attack positions with oscillating control surface deflections. Significant viscous and transonic effects in the experimental aerodynamics which cannot be predicted by the Doublet Lattice method are shown. This study should assist development of empirical correction methods that may be applied to improve Doublet Lattice calculations of lifting surface aerodynamics.

In-Flight Boundary-Layer Transition of a Large Flat Plate at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Banks, D. W.; Frederick, M. A.; Tracy, R. R.; Matisheck, J. R.; Vanecek, N. D.

2012-01-01

A flight experiment was conducted to investigate the pressure distribution, local-flow conditions, and boundary-layer transition characteristics on a large flat plate in flight at supersonic speeds up to Mach 2.00. The tests used a NASA testbed aircraft with a bottom centerline mounted test fixture. The primary objective of the test was to characterize the local flow field in preparation for future tests of a high Reynolds number natural laminar flow test article. A second objective was to determine the boundary-layer transition characteristics on the flat plate and the effectiveness of using a simplified surface coating. Boundary-layer transition was captured in both analog and digital formats using an onboard infrared imaging system. Surface pressures were measured on the surface of the flat plate. Flow field measurements near the leading edge of the test fixture revealed the local flow characteristics including downwash, sidewash, and local Mach number. Results also indicated that the simplified surface coating did not provide sufficient insulation from the metallic structure, which likely had a substantial effect on boundary-layer transition compared with that of an adiabatic surface. Cold wall conditions were predominant during the acceleration to maximum Mach number, and warm wall conditions were evident during the subsequent deceleration.

Measurement of Unsteady Blade Surface Pressure on a Single Rotation Large Scale Advanced Prop-fan with Angular and Wake Inflow at Mach Numbers from 0.02 to 0.70

NASA Technical Reports Server (NTRS)

Bushnell, P.; Gruber, M.; Parzych, D.

1988-01-01

Unsteady blade surface pressure data for the Large-Scale Advanced Prop-Fan (LAP) blade operation with angular inflow, wake inflow and uniform flow over a range of inflow Mach numbers of 0.02 to 0.70 is provided. The data are presented as Fourier coefficients for the first 35 harmonics of shaft rotational frequency. Also presented is a brief discussion of the unsteady blade response observed at takeoff and cruise conditions with angular and wake inflow.

Heat transfer investigation of two Langley Research Center delta wing configurations at a Mach number of 10.5, volume 1

NASA Technical Reports Server (NTRS)

Eaves, R. H.; Buchanan, T. D.; Warmbrod, J. D.; Johnson, C. B.

1972-01-01

Heat transfer tests for two delta wing configurations were conducted in the hypervelocity wind tunnel. The 24-inch long models were tested at a Mach number of approximately 10.5 and at angles of attack of 20, 40, and 60 degrees over a length Reynolds number range from 5 million to 23 million on 4 May to 4 June 1971. Heat transfer results were obtained from model surface heat gage measurements and thermographic phosphor paint.

Surface flow and heating distributions on a cylinder in near wake of Aeroassist Flight Experiment (AFE) configuration at incidence in Mach 10 Air

NASA Technical Reports Server (NTRS)

Wells, William L.

1990-01-01

Experimental heat transfer distributions and surface streamline directions are presented for a cylinder in the near wake of the Aeroassist Flight Experiment forebody configuration. Tests were conducted in air at a nominal free stream Mach number of 10, with post shock Reynolds numbers based on model base height of 6,450 to 50,770, and angles of attack of 5, 0, -5, and -10 degrees. Heat transfer data were obtained with thin film resistance gage and surface streamline directions by the oil flow technique. Comparisons between measured values and predicted values were made by using a Navier-Stokes computer code.

Pressure distribution for the wing of the YAV-8B airplane; with and without pylons

NASA Technical Reports Server (NTRS)

Saltzman, Edwin J.; Delfrate, John H.; Sabsay, Catherine M.; Yarger, Jill M.

1992-01-01

Pressure distribution data have been obtained in flight at four span stations on the wing panel of the YAV-8B airplane. Data obtained for the supercritical profiled wing, with and without pylons installed, ranged from Mach 0.46 to 0.88. The altitude ranged from approximately 20,000 to 40,000 ft and the resultant Reynolds numbers varied from approximately 7.2 million to 28.7 million based on the mean aerodynamic chord. Pressure distribution data and flow visualization results show that the full-scale flight wing performance is compromised because the lower surface cusp region experiences flow separation for some important transonic flight conditions. This condition is aggravated when local shocks occur on the lower surface of the wing (mostly between 20 and 35 percent chord) when the pylons are installed for Mach 0.8 and above. There is evidence that convex fairings, which cover the pylon attachment flanges, cause these local shocks. Pressure coefficients significantly more negative than those for sonic flow also occur farther aft on the lower surface (near 60 percent chord) whether or not the pylons are installed for Mach numbers greater than or equal to 0.8. These negative pressure coefficient peaks and associated local shocks would be expected to cause increasing wave and separation drag at transonic Mach number increases.

Selected results of the F-15 propulsion interactions program

NASA Technical Reports Server (NTRS)

Webb, L. D.; Nugent, J.

1982-01-01

A better understanding of propulsion system/airframe flow interactions could aid in the reduction of aircraft drag. For this purpose, NASA and the United States Air Force have conducted a series of wind-tunnel and flight tests on the F-15 airplane. This paper presents a correlation of flight test data from tests conducted at the NASA Dryden Flight Research Facility of the Ames Research Center, with data obtained from wind-tunnel tests. Flights were made at stabilized Mach numbers around 0.6, 0.9, 1.2, and 1.5 with accelerations up to near Mach number 2. Wind-tunnel tests used a 7.5 percent-scale F-15 inlet/airframe model. Flight and wind-tunnel pressure coefficients showed good agreement in most cases. Correlation of interaction effects caused by changes in cowl angle, angle-of-attack, and Mach number are presented. For the afterbody region, the pressure coefficients on the nozzle surfaces were influenced by boattail angles and Mach number. Boundary-layer thickness decreased as angle of attack increased above 4 deg.

Unsteady pressure and structural response measurements of an elastic supercritical wing

NASA Technical Reports Server (NTRS)

Eckstrom, Clinton V.; Seidel, David A.; Sandford, Maynard C.

1988-01-01

Results are presented which define unsteady flow conditions associated with high dynamic response experienced on a high aspect ratio elastic supercritical wing at transonic test conditions while being tested in the NASA Langley Transonic Dynamics Tunnel. The supercritical wing, designed for a cruise Mach number of 0.80, experienced the high dynamic response in the Mach number range from 0.90 to 0.94 with the maximum response occurring at a Mach number of approximately 0.92. At the maximum wing response condition the forcing function appears to be the oscillatory chordwise movement of strong shocks located on both the wing upper and lower surfaces in conjunction with the flow separating and reattaching in the trailing edge region.

Unsteady pressure and structural response measurements on an elastic supercritical wing

NASA Technical Reports Server (NTRS)

Eckstrom, Clinton V.; Seidel, David A.; Sandford, Maynard C.

1988-01-01

Results are presented which define unsteady flow conditions associated with high dynamic response experienced on a high aspect ratio elastic supercritical wing at transonic test conditions while being tested in the NASA Langley Transonic Dynamics Tunnel. The supercritical wing, designed for a cruise Mach number of 0.80, experienced the high dynamic response in the Mach number range from 0.90 to 0.94 with the maximum response occurring at a Mach number of approximately 0.92. At the maximum wing response condition the forcing function appears to be the oscillatory chordwise movement of strong shocks located on both the wing upper and lower surfaces in conjuction with the flow separating and reattaching in the trailing edge region.

Aerodynamic characteristics of a canard-controlled missile at Mach numbers of 0.8, 1.3, and 1.75. [in the Ames 6 by 6 foot wind tunnel

NASA Technical Reports Server (NTRS)

Kassner, D. L.; Wettlaufer, B.

1977-01-01

A typical missile model with nose-mounted canards and cruciform tail surfaces was tested in the Ames 6- by 6-Foot Wind Tunnel to determine the contributions of the component aerodynamic surfaces to the static aerodynamic characteristics at Mach numbers of 0.8, 1.3, and 1.75 and Reynolds number of 625,000 based on body diameter. Data were obtained at angles of attack ranging from 0 deg to 24 deg for various stages of model build-up (i.e., with and without canard and/or tail surfaces). In addition, two different sets of canards and tail surfaces were investigated. For the canard and tail arrangements investigated, the model was trimmable at angles of attack up to about 7 deg with canard deflections of about 10 deg. Also, the tail arrangements studied provided ample pitch stability.

Combustion-Powered Actuation for Dynamic Stall Suppression - Simulations and Low-Mach Experiments

NASA Technical Reports Server (NTRS)

Matalanis, Claude G.; Min, Byung-Young; Bowles, Patrick O.; Jee, Solkeun; Wake, Brian E.; Crittenden, Tom; Woo, George; Glezer, Ari

2014-01-01

An investigation on dynamic-stall suppression capabilities of combustion-powered actuation (COMPACT) applied to a tabbed VR-12 airfoil is presented. In the first section, results from computational fluid dynamics (CFD) simulations carried out at Mach numbers from 0.3 to 0.5 are presented. Several geometric parameters are varied including the slot chordwise location and angle. Actuation pulse amplitude, frequency, and timing are also varied. The simulations suggest that cycle-averaged lift increases of approximately 4% and 8% with respect to the baseline airfoil are possible at Mach numbers of 0.4 and 0.3 for deep and near-deep dynamic-stall conditions. In the second section, static-stall results from low-speed wind-tunnel experiments are presented. Low-speed experiments and high-speed CFD suggest that slots oriented tangential to the airfoil surface produce stronger benefits than slots oriented normal to the chordline. Low-speed experiments confirm that chordwise slot locations suitable for Mach 0.3-0.4 stall suppression (based on CFD) will also be effective at lower Mach numbers.

Testing of the Crew Exploration Vehicle in NASA Langley's Unitary Plan Wind Tunnel

NASA Technical Reports Server (NTRS)

Murphy, Kelly J.; Borg, Stephen E.; Watkins, Anthony N.; Cole, Daniel R.; Schwartz, Richard J.

2007-01-01

As part of a strategic, multi-facility test program, subscale testing of NASA s Crew Exploration Vehicle was conducted in both legs of NASA Langley s Unitary Plan Wind Tunnel. The objectives of these tests were to generate aerodynamic and surface pressure data over a range of supersonic Mach numbers and reentry angles of attack for experimental and computational validation and aerodynamic database development. To provide initial information on boundary layer transition at supersonic test conditions, transition studies were conducted using temperature sensitive paint and infrared thermography optical techniques. To support implementation of these optical diagnostics in the Unitary Wind Tunnel, the experiment was first modeled using the Virtual Diagnostics Interface software. For reentry orientations of 140 to 170 degrees (heat shield forward), windward surface flow was entirely laminar for freestream unit Reynolds numbers equal to or less than 3 million per foot. Optical techniques showed qualitative evidence of forced transition on the windward heat shield with application of both distributed grit and discreet trip dots. Longitudinal static force and moment data showed the largest differences with Mach number and angle of attack variations. Differences associated with Reynolds number variation and/or laminar versus turbulent flow on the heat shield were very small. Static surface pressure data supported the aforementioned trends with Mach number, Reynolds number, and angle of attack.

Pressure distributions induced by elevon deflections on swept wings and adjacent end-plate surfaces at Mach 6

NASA Technical Reports Server (NTRS)

Kaufman, L. G., II; Johnson, C. B.

1977-01-01

Surface pressure distributions are presented for regions where three-dimensional separated flow effects are prominent on swept-wing-elevon-end-plate models of 0 degree, 50 degree, and 70 degree sweepback, and with 0 degree, 10 degree, 20 degree, and 30 degree elevon deflections. Surface-oil-flow photographs and pressure distributions on the flat-plate wing, elevon, and end-plate surfaces are presented for numerous geometric variations, including various spacings between the elevon and the end plate, with and without a tip fin. The data, for a free-stream Mach number of 6 and a wing-root-chord Reynolds number of 20 x 10 to the sixth power, reveal considerably larger regions of elevon induced loads on the adjacent end-plate surface than would be anticipated by using inviscid flow analyses.

Prediction of static aerodynamic characteristics for slender bodies alone and with lifting surfaces to very high angles of attack

NASA Technical Reports Server (NTRS)

Jorgensen, L. H.

1977-01-01

An engineering-type method is presented for computing normal-force and pitching-moment coefficients for slender bodies of circular and noncircular cross section alone and with lifting surfaces. In this method, a semi-empirical term representing viscous-separation crossflow is added to a term representing potential-theory crossflow. For many bodies of revolution, computed aerodynamic characteristics are shown to agree with measured results for investigated free-stream Mach numbers from 0.6 to 2.9. The angles of attack extend from 0 deg to 180 deg for M = 2.9 from 0 deg to 60 deg for M = 0.6 to 2.0. For several bodies of elliptic cross section, measured results are also predicted reasonably well over the investigated Mach number range from 0.6 to 2.0 and at angles of attack from 0 deg to 60 deg. As for the bodies of revolution, the predictions are best for supersonic Mach numbers. For body-wing and body-wing-tail configurations with wings of aspect ratios 3 and 4, measured normal-force coefficients and centers are predicted reasonably well at the upper test Mach number of 2.0. Vapor-screen and oil-flow pictures are shown for many body, body-wing and body-wing-tail configurations. When spearation and vortex patterns are asymmetric, undesirable side forces are measured for the models even at zero sideslip angle. Generally, the side-force coefficients decrease or vanish with the following: increase in Mach number, decrease in nose fineness ratio, change from sharp to blunt nose, and flattening of body cross section (particularly the body nose).

Fuselage and nozzle pressure distributions of a 1/12-scale F-15 propulsion model at transonic speeds. Effect of fuselage modifications and nozzle variables

NASA Technical Reports Server (NTRS)

Pendergraft, O. C., Jr.; Carson, G. T., Jr.

1984-01-01

Static pressure coefficient distributions on the forebody, afterbody, and nozzles of a 1/12 scale F-15 propulsion model was determined in the 16 foot transonic tunnel for Mach numbers from 0.60 to 1.20, angles of attack from -2 deg to 7 deg and ratio of jet total pressure to free stream static pressure from 1 up to about 7, depending on Mach number. The effects of nozzle geometry and horizontal tail deflection on the pressure distributions were investigated. Boundary layer total pressure profiles were determined at two locations ahead of the nozzles on the top nacelle surface. Reynolds number varied from about 1.0 x 10 to the 7th power per meter, depending on Mach number.

DSMC Simulations of High Mach Number Taylor-Couette Flow

NASA Astrophysics Data System (ADS)

Pradhan, Sahadev

2017-11-01

The main focus of this work is to characterise the Taylor-Couette flow of an ideal gas between two coaxial cylinders at Mach number Ma =(Uw /√{ kbTw / m }) in the range 0.01

Leeward flow over delta wings at supersonic speeds

NASA Technical Reports Server (NTRS)

Szodruch, J. G.

1980-01-01

A survey was made of the parameters affecting the development of the leeward symmetric separated flow over slender delta wings immersed in a supersonic stream. The parameters included Mach number, Reynolds number, angle of attack, leading-edge sweep angle, and body cross-sectional shape, such that subsonic and supersonic leading-edge flows are encountered. It was seen that the boundaries between the various flow regimes existing about the leeward surface may conveniently be represented on a diagram with the components of angle of attack and Mach number normal to the leading edge as governing parameters.

Surface pressure data on a series of conical forebodies at Mach numbers from 1.70 to 4.50 and combined angles of attack and sideslip

NASA Technical Reports Server (NTRS)

Townsend, J. C.; Collins, I. K.; Howell, D. T.; Hayes, C.

1979-01-01

Tabulated surface pressure data for a series of forebodies which have analytically defined cross sections and are based on a 20 degs half-angle cone are presented without analysis. Five of the cross sections were ellipses having axis ratios of 3/1, 2/1, 1/1, 1/2, and 1/3. The sixth cross section was defined by a curve having a single lobe. The data generally cover angles of attack from -5 degs to 20 degs at angles of sideslip from 0 degs to 5 degs for Mach numbers of 1.70, 2.50, 3.95, and 4.50 at a constant Reynolds number.

Mach stem formation in reflection and focusing of weak shock acoustic pulses.

PubMed

Karzova, Maria M; Khokhlova, Vera A; Salze, Edouard; Ollivier, Sébastien; Blanc-Benon, Philippe

2015-06-01

The aim of this study is to show the evidence of Mach stem formation for very weak shock waves with acoustic Mach numbers on the order of 10(-3) to 10(-2). Two representative cases are considered: reflection of shock pulses from a rigid surface and focusing of nonlinear acoustic beams. Reflection experiments are performed in air using spark-generated shock pulses. Shock fronts are visualized using a schlieren system. Both regular and irregular types of reflection are observed. Numerical simulations are performed to demonstrate the Mach stem formation in the focal region of periodic and pulsed nonlinear beams in water.

Linear stability of compressible Taylor-Couette flow

NASA Technical Reports Server (NTRS)

Kao, Kai-Hsiung; Chow, Chuen-Yen

1992-01-01

A temporal stability analysis of compressible Taylor-Couette flow is presented. The viscous flow studied in this paper is contained between two concentric cylinders of infinite length, which are rotating with different angular velocities and are kept at different surface temperatures. The effects of differential rotation and temperature difference on the stability of Taylor-Couette flow are contrasted for a range of Mach numbers ranging from incompressible to Mach 3.0. The relative motion of the cylinders dramatically affects the characteristics of the Couette flow at the onset of instability. The flow is stabilized or destabilized depending upon the temperature ratio and speeds of the two cylinders. Independent of Mach number and temperature ratio, increasing Reynolds number generally promotes a destabilizing effect, indicating the inviscid nature of the Taylor-Couette flow.

Free-Spinning-Tunnel Investigation of a 1/20-Scale Model of an Unswept-Wing Jet-Propelled Trainer Airplane

NASA Technical Reports Server (NTRS)

Bowman, James S., Jr.; Healy, Frederick M.

1960-01-01

A flutter analysis employing the kernel function for three- dimensional, subsonic, compressible flow is applied to a flutter-tested tail surface which has an aspect ratio of 3.5, a taper ratio of 0.15, and a leading-edge sweep of 30 deg. Theoretical and experimental results are compared at Mach numbers from 0.75 to 0.98. Good agreement between theoretical and experimental flutter dynamic pressures and frequencies is achieved at Mach numbers to 0.92. At Mach numbers from 0.92 to 0.98, however, a second solution to the flutter determinant results in a spurious theoretical flutter boundary which is at a much lower dynamic pressure and at a much higher frequency than the experimental boundary.

Flight-measured pressure characteristics of aft-facing steps in high Reynolds number flow at Mach numbers of 2.20, 2.50, and 2.80 and comparison with other data

NASA Technical Reports Server (NTRS)

Powers, S. G.

1978-01-01

The YF-12 airplane was studied to determine the pressure characteristics associated with an aft-facing step in high Reynolds number flow for nominal Mach numbers of 2.20, 2.50, and 2.80. Base pressure coefficients were obtained for three step heights. The surface static pressures ahead of and behind the step were measured for the no-step condition and for each of the step heights. A boundary layer rake was used to determine the local boundary layer conditions. The Reynolds number based on the length of flow ahead of the step was approximately 10 to the 8th power and the ratios of momentum thickness to step height ranged from 0.2 to 1.0. Base pressure coefficients were compared with other available data at similar Mach numbers and at ratios of momentum thickness to step height near 1.0. In addition, the data were compared with base pressure coefficients calculated by a semiempirical prediction method. The base pressure ratios are shown to be a function of Reynolds number based on momentum thickness. Profiles of the surface pressures ahead of and behind the step and the local boundary layer conditions are also presented.

High-Speed Experiments on Combustion-Powered Actuation for Dynamic Stall Suppression

NASA Technical Reports Server (NTRS)

Matalanis, Claude; Bowles, Patrick; Lorber, Peter; Crittenden, Thomas; Glezer, Ari; Schaeffler, Norman; Min, Byung-Young; Jee, Solkeun; Kuczek, Andrzej; Wake, Brian

2016-01-01

This work documents high-speed wind tunnel experiments conducted on a pitching airfoil equipped with an array of combustion-powered actuators (COMPACT). The main objective of these experiments was to demonstrate the stall-suppression capability of COMPACT on a high-lift rotorcraft airfoil, the VR-12, at relevant Mach numbers. Through dynamic pressure measurements at the airfoil surface it was shown that COMPACT can positively affect the stall behavior of the VR-12 at Mach numbers up to 0.4. Static airfoil results demonstrated 25% and 50% increases in post-stall lift at Mach numbers of 0.4 and 0.3, respectively. Deep dynamic stall results showed cycle-averaged lift coefficient increases up to 11% at Mach 0.4. Furthermore, it was shown that these benefits could be achieved with relatively few pulses during down-stroke and with no need to pre-anticipate the stall event. The flow mechanisms responsible for stall suppression were investigated using particle image velocimetry.

Recent CFD Simulations of Shuttle Orbiter Contingency Abort Aerodynamics

NASA Technical Reports Server (NTRS)

Papadopoulos, Periklis; Prabhu, Dinesh; Wright, Michael; Davies, Carol; McDaniel, Ryan; Venkatapathy, Ethiraj; Wersinski, Paul; Gomez, Reynaldo; Arnold, Jim (Technical Monitor)

2001-01-01

Modern Computational Fluid Dynamics (CFD) techniques were used to compute aerodynamic forces and moments of the Space Shuttle Orbiter in specific portions of contingency abort trajectory space. The trajectory space covers a Mach number range of 3.5-15, an angle-of-attack range of 20-60 degrees, an altitude range of 100-190 kft, and several different settings of the control surfaces (elevons, body flap, and speed brake). While approximately 40 cases have been computed, only a sampling of the results is presented here. The computed results, in general, are in good agreement with the Orbiter Operational Aerodynamic Data Book (OADB) data (i.e., within the uncertainty bands) for almost all the cases. However, in a limited number of high angle-of-attack cases (at Mach 15), there are significant differences between the computed results, especially the vehicle pitching moment, and the OADB data. A preliminary analysis of the data from the CFD simulations at Mach 15 shows that these differences can be attributed to real-gas/Mach number effects.

Flight and wind-tunnel correlation of boundary-layer transition on the AEDC transition cone

NASA Technical Reports Server (NTRS)

Fisher, D. L.; Dougherty, N. S., Jr.

1982-01-01

Transition and fluctuating surface pressure data were acquired on a 10 deg included angle cone, using the same instrumentation and technique over a wide range of Mach and Reynolds numbers in 23 wind tunnels and in flight. Transition was detected with a traversing pitot-pressure probe in contact with the surface. The surface pressure fluctuations were measured with microphones set flush in the cone surface. Good correlation of end of transition Reynolds number RE(T) was obtained between data from the lower disturbance wind tunnels and flight up to a boundary layer edge Mach number, M(e) = 1.2. Above M(e) = 1.2, however, this correlation deteriorates, with the flight Re(T) being 25 to 30% higher than the wind tunnel Re(T) at M(e) = 1.6. The end of transition Reynolds number correlated within + or - 20% with the surface pressure fluctuations, according to the equation used. Broad peaks in the power spectral density distributions indicated that Tollmien-Schlichting waves were the probable cause of transition in flight and in some of the wind tunnels.

The Total-Pressure Recovery and Drag Characteristics of Several Auxiliary Inlets at Transonic Speeds

NASA Technical Reports Server (NTRS)

Dennard, John S.

1959-01-01

Several flush and scoop-type auxiliary inlets have been tested for a range of Mach numbers from 0.55 to 1.3 to determine their transonic total-pressure recovery and drag characteristics. The inlet dimensions were comparable with the thickness of the boundary layer in which they were tested. Results indicate that flush inlets should be inclined at very shallow angles with respect to the surface for optimum total-pressure recovery and drag characteristics. Deep, narrow inlets have lower drag than wide shallow ones at Mach numbers greater than 0.9 but at lower Mach numbers the wider inlets proved superior. Inlets with a shallow approach ramp, 7 deg, and diverging ramp walls which incorporated boundary-layer bypass had lower drag than any other inlet tested for Mach numbers up to 1.2 and had the highest pressure recovery of all of the flush inlets. The scoop inlets, which operated in a higher velocity flow than the flush inlets, had higher drag coefficients. Several of these auxiliary inlets projected multiple, periodic shock waves into the stream when they were operated at low mass-flow ratios.

Experimental evaluation of shockless supercritical airfoils in cascade

NASA Technical Reports Server (NTRS)

Boldman, D. R.; Buggele, A. E.; Shaw, L. M.

1983-01-01

Surface Mach number distributions, total pressure loss coefficients, and schlieren images of the flow are presented over a range of inlet Mach numbers and air angles. Several different trailing edge geometries were tested. At design conditions a leading edge separation bubble was observed resulting in higher losses than anticipated. The minimum losses were obtained at a negative incidence condition in which the flow was accelerating over most of the supercritical region. Relatively minor differences in losses were measured with the different trailing edge geometries studied.

Aerodynamic characteristics of a canard-controlled missile at Mach numbers of 1.5 and 2.0.

NASA Technical Reports Server (NTRS)

Kassner, D. L.; Wettlaufer, B.

1977-01-01

A typical missile model with nose mounted canards and cruciform tail surfaces was tested in the Ames 6- by 6-Foot Wind Tunnel to determine the contributions of the component aerodynamic surfaces to the static aerodynamic characteristics at Mach numbers of 1.5 and 2.0 and Reynolds number of 1 million based on body diameter. Data were obtained at angles of attack ranging from -3 deg to 12 deg for various stages of model build-up (i.e., with and without canard and/or tail surfaces). Results were obtained both with the model unrolled and rolled 45 deg. For the canard and tail arrangements investigated, the model was trimmable at angles of attack up to about 10 deg with canard deflections of 9 deg. Also, the tail arrangements studied provided ample pitch stability. there were no appreciable effects of model roll orientation.

Wind-tunnel tests on a 3-dimensional fixed-geometry scramjet inlet at M = 2.30 to 4.60

NASA Technical Reports Server (NTRS)

Mueller, J. N.; Trexler, C. A.; Souders, S. W.

1977-01-01

Wind-tunnel tests were conducted on a baseline scramjet inlet model having fixed geometry and swept leading edges at M = 2.30, 2.96, 3.95, and 4.60 in the Langley unitary plan wind tunnel. The unit Reynolds number of the tests was held constant at 6.56 million per meter (2 million per foot). The objectives of the tests were to establish inlet performance and starting characteristics in the lower Mach number range of operation (less than M = 5). Surface pressures obtained on the inlet components are presented, along with the results of the internal flow surveys made at the throat and capture stations of the inlet. Contour plots of the inlet-flow-field parameters such as Mach numbers, pressure recovery, flow capture, local static and total pressure ratios at the survey stations are shown for the test Mach numbers.

Wind tunnel and analytical investigation of over-the-wing propulsion/air frame interferences for a short-haul aircraft at Mach numbers from 0.6 to 0.78. [conducted in the Lewis 8 by 6 foot tunnel

NASA Technical Reports Server (NTRS)

Wells, O. D.; Lopez, M. L.; Welge, H. R.; Henne, P. A.; Sewell, A. E.

1977-01-01

Results of analytical calculations and wind tunnel tests at cruise speeds of a representative four engine short haul aircraft employing upper surface blowing (USB) with a supercritical wing are discussed. Wind tunnel tests covered a range of Mach number M from 0.6 to 0.78. Tests explored the use of three USB nozzle configurations. Results are shown for the isolated wing body and for each of the three nozzle types installed. Experimental results indicate that a low angle nacelle and streamline contoured nacelle yielded the same interference drag at the design Mach number. A high angle powered lift nacelle had higher interference drag primarily because of nacelle boattail low pressures and flow separation. Results of varying the spacing between the nacelles and the use of trailing edge flap deflections, wing upper surface contouring, and a convergent-divergent nozzle to reduce potential adverse jet effects were also discussed. Analytical comparisons with experimental data, made for selected cases, indicate favorable agreement.

Assessment of Turbulent CFD Against STS-128 Hypersonic Flight Data

NASA Technical Reports Server (NTRS)

Wood, William A.; Kleb, William L.; Hyatt, Andrew J.

2010-01-01

Turbulent CFD simulations are compared against surface temperature measurements of the space shuttle orbiter windward tiles at reentry flight conditions. Algebraic turbulence models are used within both the LAURA and DPLR CFD codes. The flight data are from temperature measurements obtained by seven thermocouples during the STS-128 mission (September 2009). The flight data indicate boundary layer transition onset over the Mach number range 13.5{15.5, depending upon the location on the vehicle. But the boundary layer flow appeared to be transitional down through Mach 12, based upon the flight data and CFD trends. At Mach 9 the simulations match the flight data on average within 20 F/11 C, where typical surface temperatures were approximately 1600 F/870 C.

Computation of nonstationary strong shock diffraction by curved surfaces

NASA Technical Reports Server (NTRS)

Yang, J. Y.; Lombard, C. K.; Bershader, D.

1986-01-01

A two-dimensional, high resolution shock-capturing algorithm was used on a supercomputer to solve Eulerian gasdynamic equations in order to simulate nonstationary strong shock diffraction by a circular arc model in a shock tube. The hypersonic Mach shock wave was assumed to arrive at a high angle of incidence, and attention was given to the effect of varying values of the ratio of specific heats on the shock diffraction process. Details of the conservation equations of the numerical algorithm, written in curvilinear coordinates, are provided, and model output is illustrated with the results generated for a Mach shock encountering a 15 deg circular arc. The sample graphics include isopycnics, a shock surface density profile, and pressure and Mach number contours.

Wind tunnel results for a high-speed, natural laminar-flow airfoil designed for general aviation aircraft

NASA Technical Reports Server (NTRS)

Sewall, William G.; Mcghee, Robert J.; Viken, Jeffery K.; Waggoner, Edgar G.; Walker, Betty S.; Millard, Betty F.

1985-01-01

Two dimensional wind tunnel tests were conducted on a high speed natural laminar flow airfoil in both the Langley 6 x 28 inch Transonic Tunnel and the Langley Low Turbulence Pressure Tunnel. The test conditions consisted of Mach numbers ranging from 0.10 to 0.77 and Reynolds numbers ranging from 3 x 1 million to 11 x 1 million. The airfoil was designed for a lift coefficient of 0.20 at a Mach number of 0.70 and Reynolds number of 11 x 1 million. At these conditions, laminar flow would extend back to 50 percent chord of the upper surface and 70 percent chord of the lower surface. Low speed results were also obtained with a 0.20 chord trailing edge split flap deflected 60 deg.

Goertler instability in compressible boundary layers along curved surfaces with suction and cooling

NASA Technical Reports Server (NTRS)

El-Hady, N.; Verma, A. K.

1982-01-01

The Goertler instability of the laminar compressible boundary layer flows along concave surfaces is investigated. The linearized disturbance equations for the three-dimensional, counter-rotating streamwise vortices in two-dimensional boundary layers are presented in an orthogonal curvilinear coordinate. The basic approximation of the disturbance equations, that includes the effect of the growth of the boundary layer, is considered and solved numerically. The effect of compressibility on critical stability limits, growth rates, and amplitude ratios of the vortices is evaluated for a range of Mach numbers for 0 to 5. The effect of wall cooling and suction of the boundary layer on the development of Goertler vortices is investigated for different Mach numbers.

Pressure and thermal distributions on wings and adjacent surfaces induced by elevon deflections at Mach 6

NASA Technical Reports Server (NTRS)

Kaufman, L. G., II; Johnson, C. B.

1979-01-01

Surface pressure distributions and heat transfer distributions were obtained on wing half-models in regions where three dimensional separated flow effects are prominent. Unswept and 50 deg and 70 deg swept semispan wings were tested, for trailing-edge-elevon ramp angles of 0 deg, 10 deg, 20 deg, and 30 deg, with and without cylindrical and flat plate center bodies and with and without various wing-tip plates and fins. The data, obtained for a free stream Mach number of 6 and a wing-root-chord Reynolds number of 18.5 million, reveal considerably larger regions of increased pressure and thermal loads than would be anticipated using non-separated flow analyses.

Wind-Tunnel Tests of Seven Static-Pressure Probes at Transonic Speeds

NASA Technical Reports Server (NTRS)

Capone, Francis J.

1961-01-01

Wind-tunnel tests have been conducted to determine the errors of 3 seven static-pressure probes mounted very close to the nose of a body of revolution simulating a missile forebody. The tests were conducted at Mach numbers from 0.80 to 1.08 and at angles of attack from -1.7 deg to 8.4 deg. The test Reynolds number per foot varied from 3.35 x 10(exp 6) to 4.05 x 10(exp 6). For three 4-vane, gimbaled probes, the static-pressure errors remained constant throughout the test angle-of-attack range for all Mach numbers except 1.02. For two single-vane, self-rotating probes having two orifices at +/-37.5 deg. from the plane of symmetry on the lower surface of the probe body, the static-pressure error varied as much as 1.5 percent of free-stream static pressure through the test angle-of- attack range for all Mach numbers. For two fixed, cone-cylinder probes of short length and large diameter, the static-pressure error varied over the test angle-of-attack range at constant Mach numbers as much as 8 to 10 percent of free-stream static pressure.

Trailing Edge Blowing on a Two-Dimensional Six-Percent Thick Elliptical Circulation Control Airfoil Up to Transonic Conditions

NASA Technical Reports Server (NTRS)

Alexander, Michael G.; Anders, Scott G.; Johnson, Stuart K.; Florance, Jennifer P.; Keller, Donald F.

2005-01-01

A wind tunnel test was conducted in the NASA Langley Transonic Dynamics Tunnel (TDT) on a six percent thick slightly cambered elliptical circulation control airfoil with both upper and lower surface blowing capability. Parametric evaluations of jet slot heights and Coanda surface shapes were conducted at momentum coefficients (Cm) from 0.0 to 0.12. Test data were acquired at Mach numbers of 0.3, 0.5, 0.7, 0.8, and 0.84 at Reynolds numbers per foot of 2.43 x 105 to 1.05 x 106. For a transonic condition, (Mach = 0.8 at alpha = 3 degrees), it was generally found the smaller slot and larger Coanda surface combination was overall more effective than other slot/Coanda surface combinations. Lower surface blowing was not as effective as the upper surface blowing over the same range of momentum coefficients. No appreciable Coanda surface, slot height, or slot blowing position preference was indicated transonically with the dual slot blowing.

An experimental study of the flow field surrounding a subsonic jet in a cross flow. M.S. Thesis

NASA Technical Reports Server (NTRS)

Dennis, Robert Foster

1993-01-01

An experimental investigation of the flow interaction of a 5.08 cm (2.00 in.) diameter round subsonic jet exhausting perpendicularly to a flat plate in a subsonic cross flow was conducted in the NASA Ames 7x1O ft. Wind Tunnel Number One. Flat plate surface pressures were measured at 400 locations in a 30.48 cm (12.0 in.) concentric circular array surrounding the jet exit. Results from these measurements are provided in tabular and graphical form for jet-to-crossflow velocity ratios ranging from 4 to 12, and for jet exit Mach numbers ranging from 0.50 to 0.93. Laser doppler velocimeter (LDV) three component velocity measurements were made in selected regions in the developed jet plume and near the flat plate surface, at a jet Mach number of 0.50 and jet-to-crossflow velocity ratios of 6 and 8. The results of both pressure and LDV measurements are compared with the results of previous experiments. In addition, pictures of the jet plume shape at jet velocity ratios ranging from 4 to 12 were obtained using schleiren photography. The LDV measurements are consistent with previous work, but more extensive measurements will be necessary to provide a detailed picture of the flow field. The surface pressure results compare closely with previous work and provide a useful characterization of jet induced surface pressures. The results demonstrate the primary influence of jet velocity ratio and the secondary influence of jet Mach number in determining such surface pressures.

An experimental investigation of an oblique-wing and body combination at Mach numbers between 0.60 and 1.40

NASA Technical Reports Server (NTRS)

Graham, L. A.; Jones, R. T.; Boltz, F. W.

1972-01-01

An experimental investigation was conducted in an 11- by 11-foot wind tunnel to determine the aerodynamic characteristics of an oblique high aspect ratio wing in combination with a high fineness-ratio Sears-Haack body. Longitudinal and lateral-directional stability data were obtained at wing yaw angles from 0 deg to 60 deg over a test Mach number range from 0.6 to 1.4 for angles of attack between minus 6 deg and 9 deg. The effects of changes in Reynolds number, dihedral, and trailing-edge angle were studied along with the effects of a roughness strip on the upper and lower surfaces of the wing. Flow-visualization studies were made to determine the nature of the flow on the wing surfaces.

Hypersonic blunt body computations including real gas effects

NASA Technical Reports Server (NTRS)

Montagne, J.-L.; Yee, H. C.; Klopfer, G. H.; Vinokur, M.

1989-01-01

Various second-order explicit and implicit TVD shock-capturing methods, a generalization of Roe's approximate Riemann solver, and a generalized flux-vector splitting scheme are used to study two-dimensional hypersonic real-gas flows. Special attention is given to the identification of some of the elements and parameters which can affect the convergence rate for high Mach numbers or real gases, but have negligible effect for low Mach numbers, for cases involving steady-state inviscid blunt flows. Blunt body calculations at Mach numbers of greater than 15 are performed to treat real-gas effects, and impinging shock results are obtained to test the treatment of slip surfaces and complex structures. Even with the addition of improvements, the convergence rate of algorithms in the hypersonic flow regime is found to be generally slower for a real gas than for a perfect gas.

Effect of wall cooling on the stability of compressible subsonic flows over smooth humps and backward-facing steps

NASA Technical Reports Server (NTRS)

Al-Maaitah, Ayman A.; Nayfeh, Ali, H.; Ragab, Saad A.

1989-01-01

The effect of wall cooling on the two-dimensional linear stability of subsonic flows over two-dimensional surface imperfections is investigated. Results are presented for flows over smooth humps and backward-facing steps with Mach numbers up to 0.8. The results show that, whereas cooling decreases the viscous instability, it increases the shear-layer instability and hence it increases the growth rates in the separation region. The coexistence of more than one instability mechanism makes a certain degree of wall cooling most effective. For the Mach numbers 0.5 and 0.8, the optimum wall temperatures are about 80 pct and 60 pct of the adiabatic wall temperature, respectively. Increasing the Mach number decreases the effectiveness of cooling slightly and reduces the optimum wall temperature.

Streamlines behind curved shock waves in axisymmetric flow fields

NASA Astrophysics Data System (ADS)

Filippi, A. A.; Skews, B. W.

2018-07-01

Streamlines behind axisymmetric curved shock waves were used to predict the internal surfaces that produced them. Axisymmetric ring wedge models with varying internal radii of curvature and leading-edge angles were used to produce numerical results. Said numerical simulations were validated using experimental shadowgraph results for a series of ring wedge test pieces. The streamlines behind curved shock waves for lower leading-edge angles are examined at Mach 3.4, whereas the highest leading-edge angle cases are explored at Mach 2.8 and 3.4. Numerical and theoretical streamlines are compared for the highest leading-edge angle cases at Mach 3.6. It was found that wall-bounding theoretical streamlines did not match the internal curved surface. This was due to extreme streamline curvature curving the streamlines when the shock angle approached the Mach angle at lower leading-edge angles. Increased Mach number and internal radius of curvature produced more reasonable results. Very good agreement was found between the theoretical and numerical streamlines at lower curvatures before the influence of the trailing edge expansion fan.

Space shuttle: Heat transfer investigation of the McDonnell-Douglas delta wing orbiter at a nominal Mach number of 10.5

NASA Technical Reports Server (NTRS)

Eaves, R. H.; Buchanan, T. D.

1972-01-01

Heat transfer tests for the delta wing orbiter were conducted in a hypervelocity wind tunnel. A 1.1 percent scale model was tested at a Mach number of approximately 10.5 over an angle of attack range from 10 to 60 degrees over a length Reynolds number range from 5 times 10 to the 6th power to 24 times 10 to the 6th power. Heat transfer results were obtained from model surface heat gage measurements and thermographic phosphor paint. Limited pressure measurements were obtained.

Heat-Transfer Measurements on a 5.5- Inch-Diameter Hemispherical Concave Nose in Free Flight at Mach Numbers up to 6.6

NASA Technical Reports Server (NTRS)

Levine, Jack; Rumsey, Charles B.

1958-01-01

The aerodynamic heat transfer to a hemispherical concave nose has been measured in free flight at Mach numbers from 3.5 to 6.6 with corresponding Reynolds numbers based on nose diameter from 7.4 x 10(exp 6) to 14 x 10(exp 6). Over the test Mach number range the heating on the cup nose, expressed as a ratio to the theoretical stagnation-point heating on a hemisphere nose of the same diameter, varied from 0.05 to 0.13 at the stagnation point of the cup, was approximately 0.1 at other locations within 40 deg of the stagnation point, and varied from 0.6 to 0.8 just inside the lip where the highest heating rates occurred. At a Mach number of 5 the total heat input integrated over the surface of the cup nose including the lip was 0.55 times the theoretical value for a hemisphere nose with laminar boundary layer and 0.76 times that for a flat face. The heating at the stagnation point was approximately 1/5 as great as steady-flow tunnel results. Extremely high heating rates at the stagnation point (on the order of 30 times the stagnation-point values of the present test), which have occurred in conjunction with unsteady oscillatory flow around cup noses in wind-tunnel tests at Mach and Reynolds numbers within the present test range, were not observed.

Large-scale Advanced Prop-fan (LAP) high speed wind tunnel test report

NASA Technical Reports Server (NTRS)

Campbell, William A.; Wainauski, Harold S.; Arseneaux, Peter J.

1988-01-01

High Speed Wind Tunnel testing of the SR-7L Large Scale Advanced Prop-Fan (LAP) is reported. The LAP is a 2.74 meter (9.0 ft) diameter, 8-bladed tractor type rated for 4475 KW (6000 SHP) at 1698 rpm. It was designated and built by Hamilton Standard under contract to the NASA Lewis Research Center. The LAP employs thin swept blades to provide efficient propulsion at flight speeds up to Mach .85. Testing was conducted in the ONERA S1-MA Atmospheric Wind Tunnel in Modane, France. The test objectives were to confirm that the LAP is free from high speed classical flutter, determine the structural and aerodynamic response to angular inflow, measure blade surface pressures (static and dynamic) and evaluate the aerodynamic performance at various blade angles, rotational speeds and Mach numbers. The measured structural and aerodynamic performance of the LAP correlated well with analytical predictions thereby providing confidence in the computer prediction codes used for the design. There were no signs of classical flutter throughout all phases of the test up to and including the 0.84 maximum Mach number achieved. Steady and unsteady blade surface pressures were successfully measured for a wide range of Mach numbers, inflow angles, rotational speeds and blade angles. No barriers were discovered that would prevent proceeding with the PTA (Prop-Fan Test Assessment) Flight Test Program scheduled for early 1987.

CFD Simulations in Support of Shuttle Orbiter Contingency Abort Aerodynamic Database Enhancement

NASA Technical Reports Server (NTRS)

Papadopoulos, Periklis E.; Prabhu, Dinesh; Wright, Michael; Davies, Carol; McDaniel, Ryan; Venkatapathy, E.; Wercinski, Paul; Gomez, R. J.

2001-01-01

Modern Computational Fluid Dynamics (CFD) techniques were used to compute aerodynamic forces and moments of the Space Shuttle Orbiter in specific portions of contingency abort trajectory space. The trajectory space covers a Mach number range of 3.5-15, an angle-of-attack range of 20deg-60deg, an altitude range of 100-190 kft, and several different settings of the control surfaces (elevons, body flap, and speed brake). Presented here are details of the methodology and comparisons of computed aerodynamic coefficients against the values in the current Orbiter Operational Aerodynamic Data Book (OADB). While approximately 40 cases have been computed, only a sampling of the results is provided here. The computed results, in general, are in good agreement with the OADB data (i.e., within the uncertainty bands) for almost all the cases. However, in a limited number of high angle-of-attack cases (at Mach 15), there are significant differences between the computed results, especially the vehicle pitching moment, and the OADB data. A preliminary analysis of the data from the CFD simulations at Mach 15 shows that these differences can be attributed to real-gas/Mach number effects. The aerodynamic coefficients and detailed surface pressure distributions of the present simulations are being used by the Shuttle Program in the evaluation of the capabilities of the Orbiter in contingency abort scenarios.

Nonlinear reflection of a spherically divergent N-wave from a plane surface: Optical interferometry measurements in air

DOE Office of Scientific and Technical Information (OSTI.GOV)

Karzova, M., E-mail: masha@acs366.phys.msu.ru; Physics Faculty, Moscow State University, Leninskie Gory, 119991 Moscow; Yuldashev, P.

2015-10-28

Mach stem is a well-known structure typically observed in the process of strong (acoustic Mach numbers greater than 0.4) step-shock waves reflection from a rigid boundary. However, this phenomenon has been much less studied for weak shocks in nonlinear acoustic fields where Mach numbers are in the range from 0.001 to 0.01 and pressure waveforms have more complicated waveforms than step shocks. The goal of this work was to demonstrate experimentally how nonlinear reflection occurs in air for very weak spherically divergent acoustic spark-generated pulses resembling an N-wave. Measurements of reflection patterns were performed using a Mach-Zehnder interferometer. A thinmore » laser beam with sub-millimeter cross-section was used to obtain the time resolution of 0.4 µs, which is 6 times higher than the time resolution of the condenser microphones. Pressure waveforms were reconstructed using the inverse Abel transform applied to the phase of the signal measured by the interferometer. The Mach stem formation was observed experimentally as a result of collision of the incident and reflected shock pulses. It was shown that irregular reflection of the pulse occurred in a dynamic way and the length of the Mach stem increased linearly while the pulse propagated along the surface. Since the front shock of the spark-generated pulse was steeper than the rear shock, irregular type of reflection was observed only for the front shock of the pulse while the rear shock reflection occurred in a regular regime.« less

Transonic Flow Past Cone Cylinders

NASA Technical Reports Server (NTRS)

Solomon, George E

1955-01-01

Experimental results are presented for transonic flow post cone-cylinder, axially symmetric bodies. The drag coefficient and surface Mach number are studied as the free-stream Mach number is varied and, wherever possible, the experimental results are compared with theoretical predictions. Interferometric results for several typical flow configurations are shown and an example of shock-free supersonic-to-subsonic compression is experimentally demonstrated. The theoretical problem of transonic flow past finite cones is discussed briefly and an approximate solution of the axially symmetric transonic equations, valid for a semi-infinite cone, is presented.

Fundamental aerodynamic characteristics of delta wings with leading-edge vortex flows

NASA Technical Reports Server (NTRS)

Wood, R. M.; Miller, D. S.

1985-01-01

An investigation of the aerodynamics of sharp leading-edge delta wings at supersonic speeds has been conducted. The supporting experimental data for this investigation were taken from published force, pressure, and flow-visualization data in which the Mach number normal to the wing leading edge is always less than 1.0. The individual upper- and lower-surface nonlinear characteristics for uncambered delta wings are determined and presented in three charts. The upper-surface data show that both the normal-force coefficient and minimum pressure coefficient increase nonlinearly with a decreasing slope with increasing angle of attack. The lower-surface normal-force coefficient was shown to be independent of Mach number and to increase nonlinearly, with an increasing slope, with increasing angle of attack. These charts are then used to define a wing-design space for sharp leading-edge delta wings.

Local Flow Conditions for Propulsion Experiments on the NASA F-15B Propulsion Flight Test Fixture

NASA Technical Reports Server (NTRS)

Vachon, Michael J.; Moes, Timothy R.; Corda, Stephen

2005-01-01

Local flow conditions were measured underneath the National Aeronautics and Space Administration F-15B airplane to support development of future experiments on the Propulsion Flight Test Fixture (PFTF). The local Mach number and flow angles were measured using a conventional air data boom on a cone-cylinder mounted under the PFTF and compared with the airplane air data nose boom measurements. At subsonic flight speeds, the airplane and PFTF Mach numbers were approximately equal. Transonic Mach number values were up to 0.1 greater at the PFTF than the airplane, which is a counterintuitive result. The PFTF local supersonic Mach numbers were as much as 0.46 less than the airplane values. The maximum local Mach number at the PFTF was approximately 1.6 at an airplane Mach number near 2.0. The PFTF local angle of attack was negative at all Mach numbers, ranging from -3 to -8 degrees. When the airplane angle of sideslip was zero, the PFTF local value was zero between Mach 0.8 and Mach 1.1, -2 degrees between Mach 1.1 and Mach 1.5, and increased from zero to 1 degree from Mach 1.5 to Mach 2.0. Airplane inlet shock waves crossed the aerodynamic interface plane between Mach 1.85 and Mach 1.90.

Propagation of propeller tone noise through a fuselage boundary layer

NASA Technical Reports Server (NTRS)

Hanson, D. B.; Magliozzi, B.

1984-01-01

In earlier experimental and analytical studies, it was found that the boundary layer on an aircraft could provide significant shielding from propeller noise at typical transport airplane cruise Mach numbers. In this paper a new three-dimensional theory is described that treats the combined effects of refraction and scattering by the fuselage and boundary layer. The complete wave field is solved by matching analytical expressions for the incident and scattered waves in the outer flow to a numerical solution in the boundary layer flow. The model for the incident waves is a near-field frequency-domain propeller source theory developed previously for free field studies. Calculations for an advanced turboprop (Prop-Fan) model flight test at 0.8 Mach number show a much smaller than expected pressure amplification at the noise directivity peak, strong boundary layer shielding in the forward quadrant, and shadowing around the fuselage. Results are presented showing the difference between fuselage surface and free-space noise predictions as a function of frequency and Mach number. Comparison of calculated and measured effects obtained in a Prop-Fan model flight test show good agreement, particularly near and aft of the plane of rotation at high cruise Mach number.

Forebody and Inlet Design for the HIFiRE 2 Flight Test

NASA Technical Reports Server (NTRS)

Ferlemann, Paul G.

2008-01-01

A forebody and inlet have been designed for the HIFiRE 2 scramjet flight test. The test will explore the operating, performance, and stability characteristics of a simple hydrocarbon-fueled scramjet combustor as it transitions from dual-mode to scramjet-mode operation and during supersonic combustion at Mach 8+ flight conditions. Requirements for the compression system were derived from inlet starting and combustor inflow requirements as well as physical size constraints. The design process is described. A planar, fixed geometry, mixed compression concept was used to produce laterally uniform flow at the inlet entrance and a conservative amount of internal contraction with respect to inlet starting. A grid sensitivity study was performed so that important flow physics caused by three-dimensional shock boundary layer interactions could be captured with confidence. Results from low Mach number operability studies, nominal trajectory cases, and high dynamic pressure heat load cases are discussed. The forebody and inlet solutions provide information for on-going combustor calculations, mass capture across the trajectory for fuel system design, and surface heating rates for thermal/structural analysis. The design has a one freestream Mach number margin for inlet starting, exceeds the high Mach number combustor entrance pressure requirement, produces high quality flow at the inlet exit for all Mach numbers and vehicle attitudes in the design space, and fits inside the booster shroud.

Comparison of concurrent strain gage- and pressure transducer-measured flight loads on a lifting reentry vehicle and correlation with wind tunnel predictions

NASA Technical Reports Server (NTRS)

Tang, M. H.; Sefic, W. J.; Sheldon, R. G.

1978-01-01

Concurrent strain gage and pressure transducer measured flight loads on a lifting reentry vehicle are compared and correlated with wind tunnel-predicted loads. Subsonic, transonic, and supersonic aerodynamic loads are presented for the left fin and control surfaces of the X-24B lifting reentry vehicle. Typical left fin pressure distributions are shown. The effects of variations in angle of attack, angle of sideslip, and Mach number on the left fin loads and rudder hinge moments are presented in coefficient form. Also presented are the effects of variations in angle of attack and Mach number on the upper flap, lower flap, and aileron hinge-moment coefficients. The effects of variations in lower flap hinge moments due to changes in lower flap deflection and Mach number are presented in terms of coefficient slopes.

Steady and unsteady aerodynamic forces from the SOUSSA surface-panel method for a fighter wing with tip missile and comparison with experiment and PANAIR

NASA Technical Reports Server (NTRS)

Cunningham, Herbert J.

1987-01-01

The body surface-panel method SOUSSA is applied to calculate steady and unsteady lift and pitching moment coefficients on a thin fighter-type wing model with and without a tip-mounted missile. Comparisons are presented with experimental results and with PANAIR and PANAIR-related calculations for Mach numbers from 0.6 to 0.9. In general the SOUSSA program, the experiments, and the PANAIR (and related) programs give lift and pitching-moment results which agree at least fairly well, except for the unsteady clean-wing experimental moment and the unsteady moment on the wing tip body calculated by a PANAIR-predecessor program at a Mach number of 0.8.

Numerical experiment on the flow field properties of a blunted body with a counterflowing jet in supersonic flows

NASA Astrophysics Data System (ADS)

Huang, Wei; Zhang, Rui-Rui; Yan, Li; Ou, Min; Moradi, R.

2018-06-01

The prediction of the drag and heat flux reduction characteristics is a very important issue in the conceptual design phase of the hypersonic vehicle. In this paper, the flow field properties around a blunted body with a counterflowing jet in the supersonic flow with the freestream Mach number being 3.98 were investigated numerically, and they are obtained by means of the two-dimensional axisymmetric Reynolds-averaged Navier-Stokes (RANS) equations coupled with the two equation standard k-ε turbulence model. The surface Stanton number distributions, as well as the surface static pressures, were extracted from the flow field structures in order to evaluate the drag and heat flux reduction characteristics. Further, the influences of the jet pressure ratio and the jet Mach number on the drag and heat flux reduction were analyzed based on the detailed code validation and grid independency analysis process. The obtained results show that the flow cell Reynolds number has a great impact on the heat flux prediction, and its best value is 5.0 for the case studied in the current study. However, the flow cell Reynolds number and the grid scale both have only a slight impact on the prediction of the surface static pressure distribution, as well as the turbulence model. The larger jet pressure ratio is beneficial for the drag and heat flux reduction, and the smaller jet Mach number is beneficial for the heat flux reduction. Further, the long penetration mode is beneficial for the drag reduction, but it is not beneficial for the heat flux reduction.

Subsonic Aerodynamic Characteristics of an Airplane Configuration with a 63 deg Sweptback Wing and Twin-Boom Tails

NASA Technical Reports Server (NTRS)

Savage, Howard F.; Edwards, George G.

1959-01-01

A wind-tunnel investigation has been conducted to determine the effects of an unconventional tail arrangement on the subsonic static longitudinal and lateral stability characteristics of a model having a 63 deg sweptback wing of aspect ratio 3.5 and a fuselage. Tail booms, extending rearward from approximately the midsemispan of each wing panel, supported independent tail assemblies well outboard of the usual position at the rear of the fuselage. The horizontal-tail surfaces had the leading edge swept back 45 deg and an aspect ratio of 2.4. The vertical tail surfaces were geometrically similar to one panel of the horizontal tail. For comparative purposes, the wing-body combination was also tested with conventional fuselage-mounted tail surfaces. The wind-tunnel tests were conducted at Mach numbers from 0.25 to 0.95 with a Reynolds number of 2,000,000, at a Mach number of 0.46 with a Reynolds number of 3,500,000, and at a Mach number of 0.20 with a Reynolds number of 7,000,000. The results of the investigation indicate that longitudinal stability existed to considerably higher lift coefficients for the outboard tail configuration than for the configuration with conventional tail. Wing fences were necessary with both configurations for the elimination of sudden changes in longitudinal stability at lift coefficients between 0.3 and 0.5. Sideslip angles up to 15 deg had only small effects upon the pitching-moment characteristics of the outboard tail configuration. There was an increase in the directional stability for the outboard tail configuration at the higher angles of attack as opposed to a decrease for the conventional tail configuration at most of the Mach numbers and Reynolds numbers of this investigation. The dihedral effect increased rapidly with increasing angle of attack for both the outboard and the conventional tail configurations but the increase was greater for the outboard tail configuration. The data indicate that the outboard tail is an effective roll control.

Flow-separation patterns on symmetric forebodies

NASA Technical Reports Server (NTRS)

Keener, Earl R.

1986-01-01

Flow-visualization studies of ogival, parabolic, and conical forebodies were made in a comprehensive investigation of the various types of flow patterns. Schlieren, vapor-screen, oil-flow, and sublimation flow-visualization tests were conducted over an angle-of-attack range from 0 deg. to 88 deg., over a Reynolds-number range from 0.3X10(6) to 2.0X10(6) (based on base diameter), and over a Mach number range from 0.1 to 2. The principal effects of angle of attack, Reynolds number, and Mach number on the occurrence of vortices, the position of vortex shedding, the principal surface-flow-separation patterns, the magnitude of surface-flow angles, and the extent of laminar and turbulent flow for symmetric, asymmetric, and wake-like flow-separation regimes are presented. It was found that the two-dimensional cylinder analogy was helpful in a qualitative sense in analyzing both the surface-flow patterns and the external flow field. The oil-flow studies showed three types of primary separation patterns at the higher Reynolds numbers owing to the influence of boundary-layer transition. The effect of angle of attack and Reynolds number is to change the axial location of the onset and extent of the primary transitional and turbulent separation regions. Crossflow inflectional-instability vortices were observed on the windward surface at angles of attack from 5 deg. to 55 deg. Their effect is to promote early transition. At low angles of attack, near 10 deg., an unexpected laminar-separation bubble occurs over the forward half of the forebody. At high angles of attack, at which vortex asymmetry occurs, the results support the proposition that the principal cause of vortex asymmetry is the hydrodynamic instability of the inviscid flow field. On the other hand, boundary-layer asymmetries also occur, especially at transitional Reynolds numbers. The position of asymmetric vortex shedding moves forward with increasing angle of attack and with increasing Reynolds number, and moves rearward with increasing Mach number.

Experimental investigation of a Mach 6 fixed-geometry inlet featuring a swept external-internal compression flow field

NASA Technical Reports Server (NTRS)

Torrence, M. G.

1975-01-01

An investigation of a fixed-geometry, swept external-internal compression inlet was conducted at a Mach number of 6.0 and a test-section Reynolds number of 1.55 x 10 to the 7th power per meter. The test conditions was constant for all runs with stagnation pressure and temperature at 20 atmospheres and 500 K, respectively. Tests were made at angles of attack of -5 deg, 0 deg, 3 deg, and 5 deg. Measurements consisted of pitot- and static-pressure surveys in inlet throat, wall static pressures, and surface temperatures. Boundary-layer bleed was provided on the centerbody and on the cowl internal surface. The inlet performance was consistently high over the range of the angle of attack tested, with an overall average total pressure recovery of 78 percent and corresponding adiabatic kinetic-energy efficiency of 99 percent. The inlet throat flow distribution was uniform and the Mach number and pressure level were of the correct magnitude for efficient combustor design. The utilization of a swept compression field to meet the starting requirements of a fixed-geometry inlet produced neither flow instability nor a tendency to unstart.

Direct measurements and analysis of skin friction and cooling downstream of multiple flush-slot injection into a turbulent Mach 6 boundary layer

NASA Technical Reports Server (NTRS)

Howard, F. G.; Strokowski, A. J.

1978-01-01

Experiments were conducted to determine the reduction in surface skin friction and the effectiveness of surface cooling downstream of one to four successive flush slots injecting cold air at an angle of 10 deg into a turbulent Mach 6 boundary layer. Data were obtained by direct measurement of surface shear and equilibrium temperatures, respectively. Increasing the number of slots decreased the skin friction, but the incremental improvement in skin-friction reduction decreased as the number of slots was increased. Cooling effectiveness was found to improve, for a given total mass injection, as the number of slots was increased from one to four. Comparison with previously reported step-slot data, however, indicated that step slots with tangential injection are more effective for both reducing skin friction and cooling than the present flush-slot configuration. Finite-difference predictions are in reasonable agreement with skin-friction data and with boundary-layer profile data.

Euler Calculations at Off-Design Conditions for an Inlet of Inward Turning RBCC-SSTO Vehicle

NASA Technical Reports Server (NTRS)

Takashima, N.; Kothari, A. P.

1998-01-01

The inviscid performance of an inward turning inlet design is calculated computationally for the first time. Hypersonic vehicle designs based on the inward turning inlets have been shown analytically to have increased effective specific impulse and lower heat load than comparably designed vehicles with two-dimensional inlets. The inward turning inlets are designed inversely from inviscid stream surfaces of known flow fields. The computational study is performed on a Mach 12 inlet design to validate the performance predicted by the design code (HAVDAC) and calculate its off-design Mach number performance. The three-dimensional Euler equations are solved for Mach 4, 8, and 12 using a software package called SAM, which consists of an unstructured mesh generator (SAMmesh), a three-dimensional unstructured mesh flow solver (SAMcfd), and a CAD-based software (SAMcad). The computed momentum averaged inlet throat pressure is within 6% of the design inlet throat pressure. The mass-flux at the inlet throat is also within 7 % of the value predicted by the design code thereby validating the accuracy of the design code. The off-design Mach number results show that flow spillage is minimal, and the variation in the mass capture ratio with Mach number is comparable to an ideal 2-D inlet. The results from the inviscid flow calculations of a Mach 12 inward turning inlet indicate that the inlet design has very good on and off-design performance which makes it a promising design candidate for future air-breathing hypersonic vehicles.

An experimental study of a three-dimensional shock wave/turbulent boundary-layer interaction at a hypersonic Mach number

NASA Technical Reports Server (NTRS)

Kussoy, M. I.; Horstman, K. C.; Kim, K.-S.

1991-01-01

Experimental data for a series of three-dimensional shock-wave/turbulent-boundary-layer interaction flows at Mach 8.2 are presented. The test bodies, composed of sharp fins fastened to a flat-plate test surface, were designed to generate flows with varying degrees of pressure gradient, boundary-layer separation, and turning angle. The data include surface-pressure, heat-transfer, and skin-friction distributions, as well as limited mean flowfield surveys both in the undisturbed and interaction regimes. The data were obtained for the purpose of validating computational models of these hypersonic interactions.

Flight test results of riblets at supersonic speeds

NASA Technical Reports Server (NTRS)

Zuniga, Fanny A.; Anderson, Bianca T.; Bertelrud, Arild

1992-01-01

A flight experiment to test and evaluate the skin friction drag characteristics of a riblet surface in turbulent flow at supersonic speeds was conducted at NASA Dryden. Riblets of groove sizes 0.0030 and 0.0013 in. were mounted on the F-104G flight test fixture. The test surfaces were surveyed with boundary layer rakes and pressure orifices to examine the boundary layer profiles and pressure distributions of the flow. Skin friction reductions caused by the riblet surface were reported based on measured differences of momentum thickness between the smooth and riblet surfaces obtained from the boundary layer data. Flight test results for the 0.0030 in. riblet show skin friction reductions of 4 to 8 % for Mach numbers ranging from 1.2 to 1.6 and Reynolds numbers ranging from 2 to 3.4 million per unit foot. The results from the 0.0013 in. riblets show skin friction reductions of 4 to 15 % for Mach 1.2 to 1.4 and Reynolds numbers ranging from 3.6 to 6 million per unit foot.

Measurements of unsteady pressure and structural response for an elastic supercritical wing

NASA Technical Reports Server (NTRS)

Eckstrom, Clinton V.; Seidel, David A.; Sandford, Maynard C.

1994-01-01

Results are presented which define unsteady flow conditions associated with the high-dynamic structural response of a high-aspect-ratio, elastic, supercritical wing at transonic speeds. The wing was tested in the Langley Transonic Dynamics Tunnel with a heavy gas test medium. The supercritical wing, designed for a cruise lift coefficient of 0.53 at a Mach number of 0.80, experienced the high-dynamic structural response from Mach 0.90 to 0.94 with the maximum response occurring at about Mach 0.92. At the maximum response conditions of the wing, the forcing function appears to be the oscillatory chordwise movement of strong shocks located on the upper and lower surfaces of the wing in conjunction with the flow separation on the lower surface of the wing in the trailing-edge cove region.

Investigation of transonic region of high dynamic response encountered on an elastic supercritical wing

NASA Technical Reports Server (NTRS)

Seidel, David A.; Eckstrom, Clinton V.; Sandford, Maynard C.

1987-01-01

Unsteady aerodynamic data were measured on an aspect ratio 10.3 elastic supercritical wing while undergoing high dynamic response above Mach number of 0.90. These tests were conducted in the NASA Langley Transonic Dynamics Tunnel. A previous test of this wing predicted an unusual instability boundary based upon subcritical response data. During the present test no instability was found, but an angle of attack dependent narrow Mach number region of high dynamic wing response was observed over a wide range of dynamic pressures. The effect on dynamic wing response of wing angle of attack, static outboard control surface deflection and a lower surface spanwise fence located near the 60 percent local chordline was investigated. The driving mechanism of the dynamic wing response appears to be related to chordwise shock movement in conjunction with flow separation and reattachment on both the upper and lower surfaces.

Investigation of transonic region of high dynamic response encountered on an elastic supercritical wing

NASA Technical Reports Server (NTRS)

Seidel, David A.; Eckstrom, Clinton V.; Sandford, Maynard C.

1987-01-01

Unsteady aerodynamic data were measured on an aspect ratio 10.3 elastic supercritical wing while undergoing high dynamic response above a Mach number of 0.90. These tests were conducted in the NASA Langley Transonic Dynamics Tunnel. A previous test of this wing predicted an unusual instability boundary based on subcritical response data. During the present test no instability was found, but an angle of attack dependent narrow Mach number region of high dynamic wing response was observed over a wide range of dynamic pressures. The effect on dynamic wing response of wing angle of attack, static outbound control surface deflection and a lower surface spanwise fence located near the 60 percent local chordline was investigated. The driving mechanism of the dynamic wing response appears to be related to chordwise shock movement in conjunction with flow separation and reattachment on both the upper and lower surfaces.

Flow-field surveys on the windward side of the NASA 040A space shuttle orbiter at 31 deg angle of attack and Mach 20 in helium

NASA Technical Reports Server (NTRS)

Ashby, G. C., Jr.; Helms, V. T., III

1977-01-01

Pitot pressure and flow angle distributions in the windward flow field of the NASA 040A space shuttle orbiter configuration and surface pressures were measured, at a Mach number of 20 and an angle of attack of 31 deg. The free stream Reynolds number, based on model length, was 5.39 x 10 to the 6th power. Results show that cores of high pitot pressure, which are related to the body-shock-wing-shock intersections, occur on the windward plane of symmetry in the vicinity of the wing-body junction and near midspan on the wing. Theoretical estimates of the flow field pitot pressures show that conical flow values for the windward plane of symmetry surface are representative of the average level over the entire lower surface.

Variable Mach number design approach for a parallel waverider with a wide-speed range based on the osculating cone theory

NASA Astrophysics Data System (ADS)

Zhao, Zhen-tao; Huang, Wei; Li, Shi-Bin; Zhang, Tian-Tian; Yan, Li

2018-06-01

In the current study, a variable Mach number waverider design approach has been proposed based on the osculating cone theory. The design Mach number of the osculating cone constant Mach number waverider with the same volumetric efficiency of the osculating cone variable Mach number waverider has been determined by writing a program for calculating the volumetric efficiencies of waveriders. The CFD approach has been utilized to verify the effectiveness of the proposed approach. At the same time, through the comparative analysis of the aerodynamic performance, the performance advantage of the osculating cone variable Mach number waverider is studied. The obtained results show that the osculating cone variable Mach number waverider owns higher lift-to-drag ratio throughout the flight profile when compared with the osculating cone constant Mach number waverider, and it has superior low-speed aerodynamic performance while maintaining nearly the same high-speed aerodynamic performance.

Results of investigations (OA77 and OA78) on an 0.015-scale 140A/B configuration space shuttle vehicle orbiter model 49-0 in the AEDC VKF B and C wind tunnels, revision A

NASA Technical Reports Server (NTRS)

Gillins, R. L.

1975-01-01

Aerodynamic data obtained from wind tunnel tests of an 0.015-scale 140A/B configuration SSV Orbiter model in the AEDC VKF B and C wind tunnels are presented. Tests were conducted at Mach numbers of 6 and 8 in the B tunnel and at a Mach number of 10 to in the C tunnel to verify hypersonic stability and control characteristics, determine control surface effectiveness, and investigate Reynolds number effects of the 140A/B configuration. Force data were obtained for various control surface settings and Reynolds numbers in the angle-of-attack range of 15 deg to 45 deg and at angles of sideslip of -5 deg to +10 deg. Data were obtained for a few configurations at angles of attack from -27 deg to 45 deg. Control surface variables included elevon, rudder, speedbrake and bodyflap deflections. The effects of an alternate wing leading edge shape were investigated to determine its hypersonic stability and control characteristics.

In-Flight Boundary-Layer Transition on a Large Flat Plate at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Banks, Daniel W.; Fredericks, Michael Alan; Tracy, Richard R.; Matisheck, Jason R.; Vanecek, Neal D.

2012-01-01

A flight experiment was conducted to investigate the pressure distribution, local flow conditions, and boundary-layer transition characteristics on a large flat plate in flight at supersonic speeds up to Mach 2.0. The primary objective of the test was to characterize the local flow field in preparation for future tests of a high Reynolds number natural laminar flow test article. The tests used a F-15B testbed aircraft with a bottom centerline mounted test fixture. A second objective was to determine the boundary-layer transition characteristics on the flat plate and the effectiveness of using a simplified surface coating for future laminar flow flight tests employing infrared thermography. Boundary-layer transition was captured using an onboard infrared imaging system. The infrared imagery was captured in both analog and digital formats. Surface pressures were measured with electronically scanned pressure modules connected to 60 surface-mounted pressure orifices. The local flow field was measured with five 5-hole conical probes mounted near the leading edge of the test fixture. Flow field measurements revealed the local flow characteristics including downwash, sidewash, and local Mach number. Results also indicated that the simplified surface coating did not provide sufficient insulation from the metallic structure, which likely had a substantial effect on boundary-layer transition compared with that of an adiabatic surface. Cold wall conditions were predominant during the acceleration to maximum Mach number, and warm wall conditions were evident during the subsequent deceleration. The infrared imaging system was able to capture shock wave impingement on the surface of the flat plate in addition to indicating laminar-to-turbulent boundary-layer transition.

Knowledge-based system for detailed blade design of turbines

NASA Astrophysics Data System (ADS)

Goel, Sanjay; Lamson, Scott

1994-03-01

A design optimization methodology that couples optimization techniques to CFD analysis for design of airfoils is presented. This technique optimizes 2D airfoil sections of a blade by minimizing the deviation of the actual Mach number distribution on the blade surface from a smooth fit of the distribution. The airfoil is not reverse engineered by specification of a precise distribution of the desired Mach number plot, only general desired characteristics of the distribution are specified for the design. Since the Mach number distribution is very complex, and cannot be conveniently represented by a single polynomial, it is partitioned into segments, each of which is characterized by a different order polynomial. The sum of the deviation of all the segments is minimized during optimization. To make intelligent changes to the airfoil geometry, it needs to be associated with features observed in the Mach number distribution. Associating the geometry parameters with independent features of the distribution is a fairly complex task. Also, for different optimization techniques to work efficiently the airfoil geometry needs to be parameterized into independent parameters, with enough degrees of freedom for adequate geometry manipulation. A high-pressure, low reaction steam turbine blade section was optimized using this methodology. The Mach number distribution was partitioned into pressure and suction surfaces and the suction surface distribution was further subdivided into leading edge, mid section and trailing edge sections. Two different airfoil representation schemes were used for defining the design variables of the optimization problem. The optimization was performed by using a combination of heuristic search and numerical optimization. The optimization results for the two schemes are discussed in the paper. The results are also compared to a manual design improvement study conducted independently by an experienced airfoil designer. The turbine blade optimization system (TBOS) is developed using the described methodology of coupling knowledge engineering with multiple search techniques for blade shape optimization. TBOS removes a major bottleneck in the design cycle by performing multiple design optimizations in parallel, and improves design quality at the same time. TBOS not only improves the design but also the designers' quality of work by taking the mundane repetitive task of design iterations away and leaving them more time for innovative design.

Effect of empennage arrangement on single-engine nozzle/afterbody static pressures at transonic speeds

NASA Technical Reports Server (NTRS)

Henderson, William P.; Burley, James R., II

1987-01-01

An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to determine the effects on empennage arrangement on single-engine nozzle/afterbody static pressures. Tests were done at Mach numbers from 0.60 to 1.20, nozzle pressure ratios from 1.0 (jet off) to 8.0. and angles of attack from -3 to 9 deg (at jet off conditions), depending on Mach number. Three empennage arrangements (aft, staggered, and forward) were investigated. Extensive measurements were made of static pressure on the nozzle/afterbody in the vicinity of the tail surfaces.

Investigations at Supersonic Speeds of 22 Triangular Wings Representing Two Airfoil Sections for Each of 11 Apex Angles

NASA Technical Reports Server (NTRS)

Love, Eugene S

1955-01-01

The results of tests of 22 triangular wings, representing two leading-edge shapes for each of 11 apex angles, at Mach numbers 1.62, 1.92, and 1.40 are presented and compared with theory. All wings have a common thickness ratio of 8 percent and a common maximum-thickness point at 18 percent chord. Lift, drag, and pitching moment are given for all wings at each Mach number. The relation of transition in the boundary layer, shocks on the wing surfaces, and characteristics of the pressure distributions is discussed for several wings.

Linearized compressible-flow theory for sonic flight speeds

NASA Technical Reports Server (NTRS)

Heaslet, Max A; Lomax, Harvard; Spreiter, John R

1950-01-01

The partial differential equation for the perturbation velocity potential is examined for free-stream Mach numbers close to and equal to one. It is found that, under the assumptions of linearized theory, solutions can be found consistent with the theory for lifting-surface problems both in stationary three-dimensional flow and in unsteady two-dimensional flow. Several examples are solved including a three dimensional swept-back wing and two dimensional harmonically-oscillating wing, both for a free stream Mach number equal to one. Momentum relations for the evaluation of wave and vortex drag are also discussed. (author)

Aerodynamic load distributions at transonic speeds for a close-coupled wing-canard configuration: Tabulated pressure data

NASA Technical Reports Server (NTRS)

Washburn, K. E.; Gloss, B. B.

1978-01-01

Wind tunnel studies are reported on both the canard and wing surfaces of a model that is geometrically identical to one used in several force and moment tests to provide insight into the various aerodynamic interference effects. In addition to detailed pressures measurements, the pressures were integrated to illustrate the effects of Mach number, canard location, and canard-wing interference on various aerodynamic parameters. Transonic pressure tunnel Mach numbers ranged from 0.70 to 1.20 for data taken from 0 deg to approximately 16 deg angle-of-attack at 0 deg sideslip.

Pressure measurements on a thick cambered and twisted 58 deg delta wing at high subsonic speeds

NASA Technical Reports Server (NTRS)

Chu, Julio; Lamar, John E.

1987-01-01

A pressure experiment at high subsonic speeds was conducted by a cambered and twisted thick delta wing at the design condition (Mach number 0.80), as well as at nearby Mach numbers (0.75 and 0.83) and over an angle-of-attack range. Effects of twin vertical tails on the wing pressure measurements were also assessed. Comparisons of detailed theoretical and experimental surface pressures and sectional characteristics for the wing alone are presented. The theoretical codes employed are FLO-57, FLO-28, PAN AIR, and the Vortex Lattice Method-Suction Analogy.

Effect of nozzle and vertical-tail variables on the performance of a 3-surface F-15 model at transonic Mach numbers. [Langley 16 foot transonic tunnel

NASA Technical Reports Server (NTRS)

Pendergraft, O. C., Jr.; Bare, E. A.

1982-01-01

An investigation was conducted in the Langley 16 foot transonic tunnel to determine the longitudinal aerodynamic characteristics of twin two dimensional nozzles and twin baseline axisymmetric nozzles installed on a fully metric 0.047 scale model of the F-15 three surface configuration (canards, wing, horizontal tails). The effects on performance of two dimensional nozzle in flight thrust reversing, locations and orientation of the vertical tails, and deflections of the horizontal tails were also determined. Test data were obtained at static conditions and at Mach numbers from 0.60 to 1.20 over an angle of attack range from -2 deg to 15 deg. Nozzle pressure ratio was varied from jet off to about 6.5.

Inlet flow field investigation. Part 1: Transonic flow field survey

NASA Technical Reports Server (NTRS)

Yetter, J. A.; Salemann, V.; Sussman, M. B.

1984-01-01

A wind tunnel investigation was conducted to determine the local inlet flow field characteristics of an advanced tactical supersonic cruise airplane. A data base for the development and validation of analytical codes directed at the analysis of inlet flow fields for advanced supersonic airplanes was established. Testing was conducted at the NASA-Langley 16-foot Transonic Tunnel at freestream Mach numbers of 0.6 to 1.20 and angles of attack from 0.0 to 10.0 degrees. Inlet flow field surveys were made at locations representative of wing (upper and lower surface) and forebody mounted inlet concepts. Results are presented in the form of local inlet flow field angle of attack, sideflow angle, and Mach number contours. Wing surface pressure distributions supplement the flow field data.

Effects of spoiler surfaces on the aeroelastic behavior of a low-aspect-ratio rectangular wing

NASA Technical Reports Server (NTRS)

Cole, Stanley R.

1990-01-01

An experimental research study to determine the effectiveness of spoiler surfaces in suppressing flutter onset for a low-aspect-ratio, rectangular wing was conducted in the Langley Transonic Dynamics Tunnel (TDT). The wing model used in this flutter test consisted of a rigid wing mounted to the wind-tunnel wall by a flexible, rectangular beam. The flexible beam was connected to the wing root and cantilever mounted to the wind-tunnel wall. The wing had a 1.5 aspect ratio based on wing semispan and a NACA 64A010 airfoil shape. The spoiler surfaces consisted of thin, rectangular aluminum plates that were vertically mounted to the wing surface. The spoiler surface geometry and location on the wing surface were varied to determine the effects of these parameters on the classical flutter of the wing model. Subsonically, the experiment showed that spoiler surfaces increased the flutter dynamic pressure with each successive increase in spoiler height or width. This subsonic increase in flutter dynamic pressure was approximately 15 percent for the maximum height spoiler configuration and for the maximum width spoiler configuration. At transonic Mach numbers, the flutter dynamic pressure conditions were increased even more substantially than at subsonic Mach numbers for some of the smaller spoiler surfaces. But greater than a certain spoiler size (in terms of either height or width) the spoilers forced a torsional instability in the transonic regime that was highly Mach number dependent. This detrimental torsional instability was found at dynamic pressures well below the expected flutter conditions. Variations in the spanwise location of the spoiler surfaces on the wing showed little effect on flutter. Flutter analysis was conducted for the basic configuration (clean wing with all spoiler surface mass properties included). The analysis correlated well with the clean wing experimental flutter results.

Bifurcation parameters of a reflected shock wave in cylindrical channels of different roughnesses

NASA Astrophysics Data System (ADS)

Penyazkov, O.; Skilandz, A.

2018-03-01

To investigate the effect of bifurcation on the induction time in cylindrical shock tubes used for chemical kinetic experiments, one should know the parameters of the bifurcation structure of a reflected shock wave. The dynamics and parameters of the shock wave bifurcation, which are caused by reflected shock wave-boundary layer interactions, are studied experimentally in argon, in air, and in a hydrogen-nitrogen mixture for Mach numbers M = 1.3-3.5 in a 76-mm-diameter shock tube without any ramp. Measurements were taken at a constant gas density behind the reflected shock wave. Over a wide range of experimental conditions, we studied the axial projection of the oblique shock wave and the pressure distribution in the vicinity of the triple Mach configuration at 50, 150, and 250 mm from the endwall, using side-wall schlieren and pressure measurements. Experiments on a polished shock tube and a shock tube with a surface roughness of 20 {μ }m Ra were carried out. The surface roughness was used for initiating small-scale turbulence in the boundary layer behind the incident shock wave. The effect of small-scale turbulence on the homogenization of the transition zone from the laminar to turbulent boundary layer along the shock tube perimeter was assessed, assuming its influence on a subsequent stabilization of the bifurcation structure size versus incident shock wave Mach number, as well as local flow parameters behind the reflected shock wave. The influence of surface roughness on the bifurcation development and pressure fluctuations near the wall, as well as on the Mach number, at which the bifurcation first develops, was analyzed. It was found that even small additional surface roughness can lead to an overshoot in pressure growth by a factor of two, but it can stabilize the bifurcation structure along the shock tube perimeter.

On the calculation of dynamic and heat loads on a three-dimensional body in a hypersonic flow

NASA Astrophysics Data System (ADS)

Bocharov, A. N.; Bityurin, V. A.; Evstigneev, N. M.; Fortov, V. E.; Golovin, N. N.; Petrovskiy, V. P.; Ryabkov, O. I.; Teplyakov, I. O.; Shustov, A. A.; Solomonov, Yu S.

2018-01-01

We consider a three-dimensional body in a hypersonic flow at zero angle of attack. Our aim is to estimate heat and aerodynamic loads on specific body elements. We are considering a previously developed code to solve coupled heat- and mass-transfer problem. The change of the surface shape is taken into account by formation of the iterative process for the wall material ablation. The solution is conducted on the multi-graphics-processing-unit (multi-GPU) cluster. Five Mach number points are considered, namely for M = 20-28. For each point we estimate body shape after surface ablation, heat loads on the surface and aerodynamic loads on the whole body and its elements. The latter is done using Gauss-type quadrature on the surface of the body. The comparison of the results for different Mach numbers is performed. We also estimate the efficiency of the Navier-Stokes code on multi-GPU and central processing unit architecture for the coupled heat and mass transfer problem.

Inviscid/Boundary-Layer Aeroheating Approach for Integrated Vehicle Design

NASA Technical Reports Server (NTRS)

Lee, Esther; Wurster, Kathryn E.

2017-01-01

A typical entry vehicle design depends on the synthesis of many essential subsystems, including thermal protection system (TPS), structures, payload, avionics, and propulsion, among others. The ability to incorporate aerothermodynamic considerations and TPS design into the early design phase is crucial, as both are closely coupled to the vehicle's aerodynamics, shape and mass. In the preliminary design stage, reasonably accurate results with rapid turn-representative entry envelope was explored. Initial results suggest that for Mach numbers ranging from 9-20, a few inviscid solutions could reasonably sup- port surface heating predictions at Mach numbers variation of +/-2, altitudes variation of +/-10 to 20 kft, and angle-of-attack variation of +/- 5. Agreement with Navier-Stokes solutions was generally found to be within 10-15% for Mach number and altitude, and 20% for angle of attack. A smaller angle-of-attack increment than the 5 deg around times for parametric studies and quickly evolving configurations are necessary to steer design decisions. This investigation considers the use of an unstructured 3D inviscid code in conjunction with an integral boundary-layer method; the former providing the flow field solution and the latter the surface heating. Sensitivity studies for Mach number, angle of attack, and altitude, examine the feasibility of using this approach to populate a representative entry flight envelope based on a limited set of inviscid solutions. Each inviscid solution is used to generate surface heating over the nearby trajectory space. A subset of a considered in this study is recommended. Results of the angle-of-attack sensitivity studies show that smaller increments may be needed for better heating predictions. The approach is well suited for application to conceptual multidisciplinary design and analysis studies where transient aeroheating environments are critical for vehicle TPS and thermal design. Concurrent prediction of aeroheating environments, coupled with the use of unstructured methods, is considered enabling for TPS material selection and design in conceptual studies where vehicle mission, shape, and entry strategies evolve rapidly.

Cone-Probe Rake Design and Calibration for Supersonic Wind Tunnel Models

NASA Technical Reports Server (NTRS)

Won, Mark J.

1999-01-01

A series of experimental investigations were conducted at the NASA Langley Unitary Plan Wind Tunnel (UPWT) to calibrate cone-probe rakes designed to measure the flow field on 1-2% scale, high-speed wind tunnel models from Mach 2.15 to 2.4. The rakes were developed from a previous design that exhibited unfavorable measurement characteristics caused by a high probe spatial density and flow blockage from the rake body. Calibration parameters included Mach number, total pressure recovery, and flow angularity. Reference conditions were determined from a localized UPWT test section flow survey using a 10deg supersonic wedge probe. Test section Mach number and total pressure were determined using a novel iterative technique that accounted for boundary layer effects on the wedge surface. Cone-probe measurements were correlated to the surveyed flow conditions using analytical functions and recursive algorithms that resolved Mach number, pressure recovery, and flow angle to within +/-0.01, +/-1% and +/-0.1deg , respectively, for angles of attack and sideslip between +/-8deg. Uncertainty estimates indicated the overall cone-probe calibration accuracy was strongly influenced by the propagation of measurement error into the calculated results.

Assessment of dual-point drag reduction for an executive-jet modified airfoil section

NASA Technical Reports Server (NTRS)

Allison, Dennis O.; Mineck, Raymond E.

1996-01-01

This paper presents aerodynamic characteristics and pressure distributions for an executive-jet modified airfoil and discusses drag reduction relative to a baseline airfoil for two cruise design points. A modified airfoil was tested in the adaptive-wall test section of the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT) for Mach numbers ranging from 0.250 to 0.780 and chord Reynolds numbers ranging from 3.0 x 10(exp 6) to 18.0 x 10(exp 6). The angle of attack was varied from minus 2 degrees to almost 10 degrees. Boundary-layer transition was fixed at 5 percent of chord on both the upper and lower surfaces of the model for most of the test. The two design Mach numbers were 0.654 and 0.735, chord Reynolds numbers were 4.5 x 10(exp 6) and 8.9 x 10(exp 6), and normal-force coefficients were 0.98 and 0.51. Test data are presented graphically as integrated force and moment coefficients and chordwise pressure distributions. The maximum normal-force coefficient decreases with increasing Mach number. At a constant normal-force coefficient in the linear region, as Mach number increases an increase occurs in the slope of normal-force coefficient versus angle of attack, negative pitching-moment coefficient, and drag coefficient. With increasing Reynolds number at a constant normal-force coefficient, the pitching-moment coefficient becomes more negative and the drag coefficient decreases. The pressure distributions reveal that when present, separation begins at the trailing edge as angle of attack is increased. The modified airfoil, which is designed with pitching moment and geometric constraints relative to the baseline airfoil, achieved drag reductions for both design points (12 and 22 counts). The drag reductions are associated with stronger suction pressures in the first 10 percent of the upper surface and weakened shock waves.

Characteristics of Mach 10 transitional and turbulent boundary layers

NASA Technical Reports Server (NTRS)

Watson, R. D.

1978-01-01

Measurements of the mean flow properties of transitional and turbulent boundary layers in helium on 4 deg and 5 deg wedges were made for flows with edge Mach numbers from 9.5 to 11.3, ratios of wall temperature to total temperature of 0.4 to 0.95, and maximum length Reynolds numbers of one hundred million. The data include pitot and total temperature surveys and measurements of heat transfer and surface shear. In addition, with the assumption of local similarity, turbulence quantities such as the mixing length were derived from the mean flow profiles. Low Reynolds number and precursor transition effects were significant factors at these test conditions and were included in finite difference boundary layer predictions.

An investigation of several NACA 1-series nose inlets with and without protruding central bodies at high-subsonic Mach numbers and at a Mach number of 1.2

NASA Technical Reports Server (NTRS)

Pendley, Robert E; Robinson, Harold L

1950-01-01

An investigation of three NACA 1-series nose inlets, two of which were fitted with protruded central bodies, was conducted in the Langley 8-foot high-speed tunnel. An elliptical-nose body, which had a critical Mach number approximately equal to that of one of the nose inlets, was also tested. Tests were made near zero angle of attack for a Mach number range from 0.4 to 0.925 and for the supersonic Mach number of 1.2. The inlet-velocity-ratio range extended from zero to a maximum value of 1.34. Measurements included pressure distribution, external drag, and total-pressure loss of the internal flow near the inlet. Drag was not measured for the tests at the supersonic Mach number. Over the range of inlet-velocity ratio investigated, the calculated external pressure-drag coefficient at a Mach number of 1.2 was consecutively lower for the nose inlets of higher critical Mach number, and the pressure-drag coefficient of the longest nose inlet was in the range of pressure-drag coefficient for two solid noses of fineness ratio 2.4 and 6.0. For Mach numbers below the Mach number of the supercritical drag rise, extrapolation of the test data indicated that the external drag of the nose inlets was little affected by the addition of central bodies at or slightly below the minimum inlet-velocity ratio for unseparated central-body flow. The addition of central bodies to the nose inlets also led to no appreciable effects on either the Mach number of the supercritical drag rise, or, for inlet-velocity ratios high enough to avoid a pressure peak at the inlet lip, on the critical Mach number. The total-pressure recovery of the inlets tested, which were of a subsonic type, was sensibly unimpaired at the supersonic Mach number of 1.2 Low-speed measurements of the minimum inlet-velocity ratio for unseparated central-body flow appear to be applicable for Mach numbers extending to 1.2.

Rocket-Model Investigation of the Longitudinal Stability, Drag, and Duct Performance Characteristics of the North American MX-770 (X-10) Missile at Mach Numbers from 0.80 to 1.70

NASA Technical Reports Server (NTRS)

Bond, Aleck C.; Swanson, Andrew G.

1953-01-01

A free-flight 0.12-scale rocket-boosted model of the North American MX-770 (X-10) missile has been tested in flight by the Pilotless Aircraft Research Division of the Langley Aeronautical Laboratory. Drag, longitudinal stability, and duct performance data were obtained at Mach numbers from 0.8 to 1.7 covering a Reynolds number range of about 9 x 10(exp 6) to 24 x 10(exp 6) based on wing mean aerodynamic chord. The lift-curve slope, static stability, and damping-in-pitch derivatives showed similar variations with Mach number, the parameters increasing from subsonic values in the transonic region and decreasing in the supersonic region. The variations were for the most part fairly smooth. The aerodynamic center of the configuration shifted rearward in the transonic region and moved forward gradually in the supersonic region. The pitching effectiveness of the canard control surfaces was maintained throughout the flight speed range, the supersonic values being somewhat greater than the subsonic. Trim values of angle of attack and lift coefficient changed abruptly in the transonic region, the change being associated with variations in the out-of-trim pitching moment, control effectiveness, and aerodynamic-center travel in this speed range. Duct total-pressure recovery decreased with increase in free-stream Mach number and the values were somewhat less than normal-shock recovery. Minimum drag data indicated a supersonic drag coefficient about twice the subsonic drag coefficient and a drag-rise Mach number of approximately 0.90. Base drag was small subsonically but was about 25 percent of the minimum drag of the configuration supersonically.

Supersonic aerodynamic characteristics of a circular body Earth-to-Orbit vehicle

NASA Technical Reports Server (NTRS)

Ware, George M.; Engelund, Walter C.; Macconochie, Ian O.

1994-01-01

The circular body configuration is a generic single- or multi-stage reusable Earth-to-orbit transport. A thick clipped-delta wing is the major lifting surface. For directional control, three different vertical fin arrangements were investigated: a conventional aft-mounted center fin, wingtip fins, and a nose-mounted fin. The tests were conducted in the Langley Unitary Plan Wind Tunnel. The configuration is longitudinally stable about the estimated center of gravity of 0.72 body length up to a Mach number of about 3.0. Above Mach 3.0, the model is longitudinally unstable at low angles of attack but has a stable secondary trim point at angles of attack above 30 deg. The model has sufficient pitch control authority with elevator and body flap to produce stable trim over the test range. The model with the center fin is directionally stable at low angles of attack up to a Mach number of 3.90. The rudder-like surfaces on the tip fins and the all-movable nose fin are designed as active controls to produce artificial directional stability and are effective in producing yawing moment. The wing trailing-edge aileron surfaces are effective in producing rolling moment, but they also produce large adverse yawing moment.

Mach number dependence of electron heating at high Mach number interplanetary shocks in the inner heliospere

NASA Astrophysics Data System (ADS)

Matsukiyo, Shuichi

In the inner heliosphere a variety of interplanetary shocks with different Mach numbers are expected to be present. A possible maximum Mach number at 0.3AU from the sun is esti-mated to be about 40. Efficiency of electron heating in such high Mach number shocks is one of the outstanding issues of space plasma physics as well as astrophysics. Here, from this aspect, electron heating rate through microinstabilities generated in the transition region of a quasi-perpendicular shock for wide range of Mach numbers is investigated. Saturation levels of effective electron temperature as a result of modified two-stream instability (MTSI) are es-timated by using a semianalytic approach which we call an extended quasilinear analysis here. The results are compared with one-dimensional full particle-in-cell simulations. It is revealed that Mach number dependence of the effective electron temperature is weak when a Mach num-ber is below a certain critical value. Above the critical value, electron temperature increases being proportional to an upstream flow energy because of that a dominant microinstability in the foot changes from the MTSI to Buneman instability. The critical Mach number is roughly estimated to be a few tens.

X-33 (Rev-F) Aeroheating Results of Test 6770 in NASA Langley 20-Inch Mach 6 Air Tunnel

NASA Technical Reports Server (NTRS)

Berry, Scott A.; Horvath, Thomas J.; Kowalkowski, Matthew K.; Liechty, Derek S.

1999-01-01

Aeroheating characteristics of the X-33 Rev-F configuration have been experimentally examined in the Langley 20-Inch Mach 6 Air Tunnel (Test 6770). Global surface heat transfer distributions, surface streamline patterns, and shock shapes were measured on a 0.013-scale model at Mach 6 in air. Parametric variations include angles-of-attack of 20-deg, 30-deg, and 40-deg; Reynolds numbers based on model length of 0.9 to 4.9 million; and body-flap deflections of 10-deg and 20-deg. The effects of discrete roughness elements on boundary layer transition, which included trip height, size, and location, both on and off the windward centerline, were investigated. This document is intended to serve as a quick release of preliminary data to the X-33 program; analysis is limited to observations of the experimental trends in order to expedite dissemination.

The NCOREL computer program for 3D nonlinear supersonic potential flow computations

NASA Technical Reports Server (NTRS)

Siclari, M. J.

1983-01-01

An innovative computational technique (NCOREL) was established for the treatment of three dimensional supersonic flows. The method is nonlinear in that it solves the nonconservative finite difference analog of the full potential equation and can predict the formation of supercritical cross flow regions, embedded and bow shocks. The method implicitly computes a conical flow at the apex (R = 0) of a spherical coordinate system and uses a fully implicit marching technique to obtain three dimensional cross flow solutions. This implies that the radial Mach number must remain supersonic. The cross flow solutions are obtained by using type dependent transonic relaxation techniques with the type dependency linked to the character of the cross flow velocity (i.e., subsonic/supersonic). The spherical coordinate system and marching on spherical surfaces is ideally suited to the computation of wing flows at low supersonic Mach numbers due to the elimination of the subsonic axial Mach number problems that exist in other marching codes that utilize Cartesian transverse marching planes.

Vibration and flutter characteristics of the SR7L large-scale propfan

NASA Technical Reports Server (NTRS)

August, Richard; Kaza, Krishna Rao V.

1988-01-01

An investigation of the vibration characteristics and aeroelastic stability of the SR7L Large-Scale Advanced Propfan was performed using a finite element blade model and an improved aeroelasticity code. Analyses were conducted for different blade pitch angles, blade support conditions, number of blades, rotational speeds, and freestream Mach numbers. A finite element model of the blade was used to determine the blade's vibration behavior and sensitivity to support stiffness. The calculated frequencies and mode shape obtained with this model agreed well with the published experimental data. A computer code recently developed at NASA Lewis Research Center and based on three-dimensional, unsteady, lifting surface aerodynamic theory was used for the aeroelastic analysis to examine the blade's stability at a cruise condition of Mach 0.8 at 1700 rpm. The results showed that the blade is stable for that operating point. However, a flutter condition was predicted if the cruise Mach number was increased to 0.9.

Wind tunnel investigation of the aerodynamic characteristics of five forebody models at high angles of attack at Mach numbers from 0.25 to 2

NASA Technical Reports Server (NTRS)

Keener, E. R.; Taleghani, J.

1975-01-01

Five forebody models of various shapes were tested in the Ames 6- by 6-Foot Wind Tunnel to determine the aerodynamic characteristics at Mach numbers from 0.25 to 2 at a Reynolds number of 800000. At a Mach number of 0.6 the Reynolds number was varied from 0.4 to 1.8 mil. Angle of attack was varied from -2 deg to 88 deg at zero sideslip. The purpose of the investigation was to determine the effect of Mach number of the side force that develops at low speeds and zero sideslip for all of these forebody models when the nose is pointed. Test results show that with increasing Mach number the maximum side forces decrease to zero between Mach numbers of 0.8 and 1.5, depending on the nose angle; the smaller the nose angle of the higher the Mach number at which the side force exists. At a Mach number of 0.6 there is some variation of side force with Reynolds number, the variation being the largest for the more slender tangent ogive.

Feasibility study of three-dimensional holographic interferometry for aerodynamics

NASA Technical Reports Server (NTRS)

Craig, J. E.

1983-01-01

Laser holographic interferometry was tried as a nonintrusive diagnostic tool for studying unsteady two dimensional flows. A NACA 0012 airfoil was tested, while undergoing dynamic stall, over a range of Mach numbers of 0.3 to 0.5, Reynolds number of 0.5 x 10 to the 6th power to 2 x 10 to the 6th power, and at reduced frequencies of 0.015 to 0.15. It was found that both quantitative and qualitative data could be obtained by the technique. Surface pressures on the airfoil can be measured to within 1% of those measured with orifices and pressure transducers when the flow is attached. Velocity profiles were measured near the wake region, and they compared very well with laser velocimeter data for attached flows. For separated flows with large scale vortices, densities can be measured, but pressures and velocities cannot be deduced with the assumption of constant pressure gradient in the normal direction. The sensitivity of the interferograms was good at a Mach number of 0.4 and a Reynolds number of 4 x 10 to the 6th power/ft; the sensitivity worsened at smaller Mach numbers and Reynolds numbers, and improved at larger ones.

Airfoil shape for flight at subsonic speeds

DOEpatents

Whitcomb, Richard T.

1976-01-01

An airfoil having an upper surface shaped to control flow accelerations and pressure distribution over the upper surface and to prevent separation of the boundary layer due to shock wave formulation at high subsonic speeds well above the critical Mach number. A highly cambered trailing edge section improves overall airfoil lifting efficiency.

Exploratory study of in-plane streamline curvature effects on a turbulent boundary layer at a Mach number of 3

NASA Technical Reports Server (NTRS)

Bogdonoff, Seymour M.

1991-01-01

This report on a program to study in-plane streamline curvature effects in a turbulent boundary layer at a Mach number of 3. The original proposal, for a 3-year program to explore in-plane streamline curvature effects on a supersonic turbulent boundary layer using three-dimensional pressure fields generated by fins and wall geometry, ended after one year. The purpose of these tests was to compare these streamline curvature effects to the more classical two-dimensional curvature generated by wall shape and imposed pressure gradients, previously considered primarily in a plane normal to the floor. The studies were carried out in the Mach number of 3, 8 x 8 inch High Reynolds Number Supersonic Tunnel. The usual surface visualization and mean wall static pressures were supplemented by the use of many small high frequency wall static pressure gauges (Kulites) to get some indication of the amplification of boundary layer disturbances by the in-plane streamline curvature caused by the three-dimensional pressure fields imposed on the boundary layer.

Numerical Study of Pressure Fluctuations due to a Mach 6 Turbulent Boundary Layer

NASA Technical Reports Server (NTRS)

Duan, Lian; Choudhari, Meelan M.

2013-01-01

Direct numerical simulations (DNS) are used to examine the pressure fluctuations generated by a Mach 6 turbulent boundary layer with nominal freestream Mach number of 6 and Reynolds number of Re(sub t) approx. =. 464. The emphasis is on comparing the primarily vortical pressure signal at the wall with the acoustic freestream signal under higher Mach number conditions. Moreover, the Mach-number dependence of pressure signals is demonstrated by comparing the current results with those of a supersonic boundary layer at Mach 2.5 and Re(sub t) approx. = 510. It is found that the freestream pressure intensity exhibits a strong Mach number dependence, irrespective of whether it is normalized by the mean wall shear stress or by the mean pressure, with the normalized fluctuation amplitude being significantly larger for the Mach 6 case. Spectral analysis shows that both the wall and freestream pressure fluctuations of the Mach 6 boundary layer have enhanced energy content at high frequencies, with the peak of the premultiplied frequency spectrum of freestream pressure fluctuations being at a frequency of omega(delta)/U(sub infinity) approx. = 3.1, which is more than twice the corresponding frequency in the Mach 2.5 case. The space-time correlations indicate that the pressure-carrying eddies for the higher Mach number case are of smaller size, less elongated in the spanwise direction, and convect with higher convection speeds relative to the Mach 2.5 case. The demonstrated Mach-number dependence of the pressure field, including radiation intensity, directionality, and convection speed, is consistent with the trend exhibited in experimental data and can be qualitatively explained by the notion of "eddy Mach wave" radiation.

Flight investigation of an underwing nacelle installation of an auxiliary-inlet ejector nozzle with a clamshell flow diverter from Mach 0.6 to 1.3

NASA Technical Reports Server (NTRS)

Head, V. L.

1972-01-01

A nozzle installation of general interest is a podded engine mounted near the aft lower surface of the wing. The effect of this installation on the performance of an auxiliary-inlet ejector nozzle with a clamshell flow diverter was investigated over a Mach number range of 0.6 to 1.3 by using a modified F-106B aircraft. The clamshell flow diverter was tested in a 17 deg position with double-hinged synchronized floating doors. The ejector nozzle trailing-edge flaps were simulated in the closed position with a rigid structure which provided a boattail angle of 10 deg. Primary nozzle area was varied as exhaust gas temperature was varied between 975 and 1561 K. With the nozzle in a subsonic cruise position, the nozzle gross thrust coefficient was 0.918 at a flight Mach number of 0.9.

Testing continuum descriptions of low-Mach-number shock structures

NASA Technical Reports Server (NTRS)

Pham-Van-diep, Gerald C.; Erwin, Daniel A.; Muntz, E. P.

1991-01-01

Numerical experiments have been performed on normal shock waves with Monte Carlo Direct Simulations (MCDS's) to investigate the validity of continuum theories at very low Mach numbers. Results from the Navier-Stokes and the Burnett equations are compared to MCDS's for both hard-sphere and Maxwell gases. It is found that the maximum-slope shock thicknesses are described equally well (within the MCDS computational scatter) by either of the continuum formulations for Mach numbers smaller than about 1.2. For Mach numbers greater that 1.2, the Burnett predictions are more accurate than the Navier-Stokes results. Temperature-density profile separations are best described by the Burnett equations for Mach numbers greater than about 1.3. At lower Mach numbers the MCDS scatter is too great to differentiate between the two continuum theories. For all Mach numbers above one, the shock shapes are more accurately described by the Burnett equations.

Effects of canard location on the aerodynamic characteristics of a blunt-nosed missile at Mach numbers of 1.5 and 2.0. [in the Ames 6x6 wind tunnel

NASA Technical Reports Server (NTRS)

Kassner, D. L.; Wettlaufer, B.

1977-01-01

A blunt-nosed missile model with nose-mounted canards and cruciform tail surfaces was tested in the Ames 6 by 6-Foot Wind Tunnel to determine the contributions of the component aerodynamic surfaces to the static aerodynamic characteristics at Mach numbers of 1.5 and 2.0 and Reynolds number of 1 million based on body diameter. Data were obtained at angles of attack ranging from -3 deg to 12 deg and canard-deflection angles from -3 deg to 15 deg for various stages of model build-up (i.e., with and without canard and/or tail surfaces). Results were obtained with the canards at two different nose locations. For the canard and tail arrangements investigated, the model was trimmable at angles of attack up to about 4 deg or 5 deg with canard deflections of 9 deg. For this blunt-nosed model, there was little effect of canard location on trim angle of attack. The tail arrangements studied provided ample pitch stability.

Space shuttle: Static aerodynamic characteristics characteristics and control effectiveness for McDonnell-Douglas orbiter configuration for Mach number range of 0.4 to 5.0

NASA Technical Reports Server (NTRS)

Ellis, R. R.

1971-01-01

An experimental aerodynamic wind tunnel investigation was conducted employing a 0.00325 scale model of the McDonnell-Douglas space shuttle orbiter configuration. This investigation was conducted in the NASA/Marshall Space Flight Center 14- by 14- inch trisonic wind tunnel. The investigation was to determine the aerodynamic characteristics of the orbiter over the Mach number range of 0.4 to 5.0, an angle of attack variation from -4 degrees to 50 degrees, and -6 degrees to 9 degrees angle of sideslip. Control surface effectiveness was investigated for elevator, aileron, and rudder deflections.

radEq Add-On Module for CFD Solver Loci-CHEM

NASA Technical Reports Server (NTRS)

McCloud, Peter

2013-01-01

Loci-CHEM to be applied to flow velocities where surface radiation due to heating from compression and friction becomes significant. The module adds a radiation equilibrium boundary condition to the computational fluid dynamics (CFD) code to produce accurate results. The module expanded the upper limit for accurate CFD solutions of Loci-CHEM from Mach 4 to Mach 10 based on Space Shuttle Orbiter Re-Entry trajectories. Loci-CHEM already has a very promising architecture and performance, but absence of radiation equilibrium boundary condition limited the application of Loci-CHEM to below Mach 4. The immediate advantage of the add-on module is that it allows Loci-CHEM to work with supersonic flows up to Mach 10. This transformed Loci-CHEM from a rocket engine- heritage CFD code with general subsonic and low-supersonic applications, to an aeroheating code with hypersonic applications. The follow-on advantage of the module is that it is a building block for additional add-on modules that will solve for the heating generated at Mach numbers higher than 10.

An Analytical Explanation for the X-43A Flush Air Data Sensing System Pressure Mismatch Between Flight and Theory

NASA Technical Reports Server (NTRS)

Ellsworth, Joel C.

2010-01-01

Following the successful Mach 7 flight test of the X-43A, unexpectedly low pressures were measured by the aft set of the onboard Flush Air Data Sensing System s pressure ports. These in-flight aft port readings were significantly lower below Mach 3.5 than was predicted by theory. The same lower readings were also seen in the Mach 10 flight of the X-43A and in wind-tunnel data. The pre-flight predictions were developed based on 2-dimensional wedge flow, which fails to predict some of the significant 3-dimensional flow features in this geometry at lower Mach numbers. Using Volterra s solution to the wave equation as a starting point, a three-dimensional finite wedge approximation to flow over the X-43A forebody is presented. The surface pressures from this approximation compare favorably with the measured wind tunnel and flight data at speeds of Mach 2.5 and 3.

Parametric Analysis of a Hypersonic Inlet using Computational Fluid Dynamics

NASA Astrophysics Data System (ADS)

Oliden, Daniel

For CFD validation, hypersonic flow fields are simulated and compared with experimental data specifically designed to recreate conditions found by hypersonic vehicles. Simulated flow fields on a cone-ogive with flare at Mach 7.2 are compared with experimental data from NASA Ames Research Center 3.5" hypersonic wind tunnel. A parametric study of turbulence models is presented and concludes that the k-kl-omega transition and SST transition turbulence model have the best correlation. Downstream of the flare's shockwave, good correlation is found for all boundary layer profiles, with some slight discrepancies of the static temperature near the surface. Simulated flow fields on a blunt cone with flare above Mach 10 are compared with experimental data from CUBRC LENS hypervelocity shock tunnel. Lack of vibrational non-equilibrium calculations causes discrepancies in heat flux near the leading edge. Temperature profiles, where non-equilibrium effects are dominant, are compared with the dissociation of molecules to show the effects of dissociation on static temperature. Following the validation studies is a parametric analysis of a hypersonic inlet from Mach 6 to 20. Compressor performance is investigated for numerous cowl leading edge locations up to speeds of Mach 10. The variable cowl study showed positive trends in compressor performance parameters for a range of Mach numbers that arise from maximizing the intake of compressed flow. An interesting phenomenon due to the change in shock wave formation for different Mach numbers developed inside the cowl that had a negative influence on the total pressure recovery. Investigation of the hypersonic inlet at different altitudes is performed to study the effects of Reynolds number, and consequently, turbulent viscous effects on compressor performance. Turbulent boundary layer separation was noted as the cause for a change in compressor performance parameters due to a change in Reynolds number. This effect would not be noticeable if laminar flow was assumed. Mach numbers up to 20 are investigated to study the effects of vibrational and chemical non-equilibrium on compressor performance. A direct impact on the trends on the kinetic energy efficiency and compressor efficiency was found due to dissociation.

Properties of Merger Shocks in Merging Galaxy Clusters

NASA Astrophysics Data System (ADS)

Ha, Ji-Hoon; Ryu, Dongsu; Kang, Hyesung

2018-04-01

X-ray shocks and radio relics detected in the cluster outskirts are commonly interpreted as shocks induced by mergers of subclumps. We study the properties of merger shocks in merging galaxy clusters, using a set of cosmological simulations for the large-scale structure formation of the universe. As a representative case, we focus on the simulated clusters that undergo almost head-on collisions with mass ratio ∼2. Due to the turbulent nature of the intracluster medium, shock surfaces are not smooth, but composed of shocks with different Mach numbers. As the merger shocks expand outward from the core to the outskirts, the average Mach number, < {M}s> , increases in time. We suggest that the shocks propagating along the merger axis could be manifested as X-ray shocks and/or radio relics. The kinetic energy through the shocks, F ϕ , peaks at ∼1 Gyr after their initial launching, or at ∼1–2 Mpc from the core. Because of the Mach number dependent model adopted here for the cosmic-ray (CR) acceleration efficiency, their CR-energy-weighted Mach number is higher with < {M}s{> }CR}∼ 3{--}4, compared to the kinetic-energy-weighted Mach number, < {M}s{> }φ ∼ 2{--}3. Most energetic shocks are to be found ahead of the lighter dark matter (DM) clump, while the heavier DM clump is located on the opposite side of clusters. Although our study is limited to the merger case considered, the results such as the means and variations of shock properties and their time evolution could be compared with the observed characteristics of merger shocks, constraining interpretations of relevant observations.

Experimental Results for a Flapped Natural-laminar-flow Airfoil with High Lift/drag Ratio

NASA Technical Reports Server (NTRS)

Mcghee, R. J.; Viken, J. K.; Pfenninger, W.; Beasley, W. D.; Harvey, W. D.

1984-01-01

Experimental results have been obtained for a flapped natural-laminar-flow airfoil, NLF(1)-0414F, in the Langley Low-Turbulence Pressure Tunnel. The tests were conducted over a Mach number range from 0.05 to 0.40 and a chord Reynolds number range from about 3.0 x 10(6) to 22.0 x 10(6). The airfoil was designed for 0.70 chord laminar flow on both surfaces at a lift coefficient of 0.40, a Reynolds number of 10.0 x 10(6), and a Mach number of 0.40. A 0.125 chord simple flap was incorporated in the design to increase the low-drag, lift-coefficient range. Results were also obtained for a 0.20 chord split-flap deflected 60 deg.

An experimental study of transonic flow about a supercritical airfoil. Static pressure and drag data obtained from tests of a supercritical airfoil and an NACA 0012 airfoil at transonic speeds, supplement

NASA Technical Reports Server (NTRS)

Spaid, F. W.; Dahlin, J. A.; Roos, F. W.; Stivers, L. S., Jr.

1983-01-01

Surface static-pressure and drag data obtained from tests of two slightly modified versions of the original NASA Whitcomb airfoil and a model of the NACA 0012 airfoil section are presented. Data for the supercritical airfoil were obtained for a free-stream Mach number range of 0.5 to 0.9, and a chord Reynolds number range of 2 x 10 to the 6th power to 4 x 10 to the 6th power. The NACA 0012 airfoil was tested at a constant chord Reynolds number of 2 x 10 to the 6th power and a free-stream Mach number range of 0.6 to 0.8.

Results from investigations in three NASA/LaRC hypersonic wind tunnels on a .004 scale model space shuttle orbiter (model 13P-O) to determine real gas effects (LA78, LA87, LA88)

NASA Technical Reports Server (NTRS)

Ball, J. W.; Klug, G. W.

1976-01-01

Results from tests in the NASA/CF4, 20 inch Mach 6 and the 22 inch Helium Tunnel consist of pressure measurements on the lower surfaces of the Rockwell Space Shuttle Orbiter. All data are in absolute pressures. Data were recorded with the model at a Mach number of 6 and 20 at angles of 10 deg to 30 deg.

Recovery Temperature, Transition, and Heat Transfer Measurements at Mach 5

NASA Technical Reports Server (NTRS)

Brinich, Paul F.

1961-01-01

Schlieren, recovery temperature, and heat-transfer measurements were made on a hollow cylinder and a cone with axes alined parallel to the stream. Both the cone and cylinder were equipped with various bluntnesses, and the tests covered a Reynolds number range up to 20 x 10(exp 6) at a free-stream Mach number of 4.95 and wall to free-stream temperature ratios from 1.8 to 5.2 (adiabatic). A substantial transition delay due to bluntness was found for both the cylinder and the cone. For the present tests (Mach 4.95), transition was delayed by a factor of 3 on the cylinder and about 2 on the cone, these delays being somewhat larger than those observed in earlier tests at Mach 3.1. Heat-transfer tests on the cylinder showed only slight effects of wall temperature level on transition location; this is to be contrasted to the large transition delays observed on conical-type bodies at low surface temperatures at Mach 3.1. The schlieren and the peak-recovery-temperature methods of detecting transition were compared with the heat-transfer results. The comparison showed that the first two methods identified a transition point which occurred just beyond the end of the laminar run as seen in the heat-transfer data.

An experimental investigation of the NASA space shuttle external tank at hypersonic Mach numbers

NASA Technical Reports Server (NTRS)

Wittliff, C. E.

1975-01-01

Pressure and heat transfer tests were conducted simulating flight conditions which the space shuttle external tank will experience prior to break-up. The tests were conducted in the Calspan 48-inch Hypersonic Shock Tunnel and simulated entry conditions for nominal, abort-once-around (AOA), and return to launch site (RTLS) launch occurrences. Surface pressure and heat-transfer-rate distributions were obtained with and without various protuberences (or exterior hardware) on the model at Mach numbers from 15.2 to 17.7 at angles of attack from -15 deg to -180 deg and at several roll angles. The tests were conducted over a Reynolds number range from 1300 to 58,000, based on model length.

Description and calibration of the Langley 6- by 19-inch transonic tunnel

NASA Technical Reports Server (NTRS)

Ladson, C. L.

1973-01-01

A description and calibration is presented of the Langley 6- by 19-inch transonic tunnel which is a two-dimensional facility with top and bottom slotted walls used for testing two-dimensional airfoil sections. Basic tunnel-empty Mach number distributions and schlieren flow photographs as well as integrated normal-force coefficients, pitching-moment coefficients, surface-pressure distributions, and schlieren flow photographs of an NACA 0012 airfoil calibration model are presented. The Mach number capability of the facility is from 0.5 to about 1.1 with a corresponding Reynolds number range of 1.5 million to 3 million based on a 4.0-in. model chord. Comparisons of experimental results from the tests with previous data are also presented.

Comparison of wind tunnel and flight test afterbody and nozzle pressures for a twin-jet fighter aircraft at transonic speeds

NASA Technical Reports Server (NTRS)

Nugent, Jack; Pendergraft, Odis C., Jr.

1987-01-01

Afterbody and nozzle pressures measured on a 1/12-scale model and in flight on a twin-jet fighter aircraft were compared as Mach number varied from 0.6 to 1.2, Reynolds number from 17.5 million to 302.5 million, and angle of attack from 1 to 7 deg. At Mach 0.6 and 0.8, nozzle pressure coefficient distributions and nozzle axial force coefficients agreed and showed good recompression. At Mach 0.9 and 1.2, flow complexity caused a loss in recompression for both flight and wind tunnel nozzle data. The flight data exhibited less negative values of pressure coefficient and lower axial force coefficients than did the wind tunnel data. Reynolds number effects were noted only at these Mach numbers. Jet temperature and mass flux ratio did not affect the comparisons of nozzle axial flow coefficient. At subsonic speeds, the levels of pressure coefficient distributions on the upper fuselage and lower nacelle surfaces for flight were less negative than those for the model. The model boundary layer thickness at the aft rake station exceeded that for the forward rake station and increased with increasing angle of attack. The flight boundary layer thickness at the aft rake station was less than that for the forward rake station and decreased with increasing angle of attack.

Control effectiveness and tip-fin dihedral effects for the HL-20 lifting-body configuration at Mach numbers from 1.6 to 4.5

NASA Technical Reports Server (NTRS)

Cruz, Christopher I.; Ware, George M.

1995-01-01

Wind tunnel tests were made with a scale model of the HL-20 in the Langley Unitary Plan Wind Tunnel. Pitch control was investigated by deflecting the elevon surfaces on the outboard fins and body flaps on the fuselage. Yaw control tests were made with the all movable center fin deflected 5 deg. Almost full negative body flap deflection (-30 deg) was required to trim the HL-20 (moment reference center at 0.54-percent body length from nose) to positive values of life in the Mach number range from 1.6 to 2.5. Elevons were twice as effective as body flaps as a longitudinal trim device. The elevons were effective as a roll control, but because of tip-fin dihedral angle, produced about as much adverse yawing moment as rolling moment. The body flaps were less effective in producing rolling moment, but produced little adverse yawing moment. The yaw effectiveness of the all movable center fin was essentially constant over the angle-of-attack range at each Mach number. The value of yawing moment, however, was small. Center-fin deflection produced almost no rolling moments. The model was directionally unstable over most of the Mach number range with tip-fin dihedral angles less than the baseline value of 50 deg.

Piloted simulation study of a balloon-assisted deployment of an aircraft at high altitude

NASA Technical Reports Server (NTRS)

Murray, James; Moes, Timothy; Norlin, Ken; Bauer, Jeffrey; Geenen, Robert; Moulton, Bryan; Hoang, Stephen

1992-01-01

A piloted simulation was used to study the feasibility of a balloon assisted deployment of a research aircraft at high altitude. In the simulation study, an unmanned, modified sailplane was carried to 110,000 ft with a high altitude balloon and released in a nose down attitude. A remote pilot controlled the aircraft through a pullout and then executed a zoom climb to a trimmed, 1 g flight condition. A small parachute was used to limit the Mach number during the pullout to avoid adverse transonic effects. The use of small rocket motor was studied for increasing the maximum attainable altitude. Aerodynamic modifications to the basic sailplane included applying supercritical airfoil gloves over the existing wing and tail surfaces. The aerodynamic model of the simulated aircraft was based on low Reynolds number wind tunnel tests and computational techniques, and included large Mach number and Reynolds number effects at high altitude. Parametric variations were performed to study the effects of launch altitude, gross weight, Mach number limit, and parachute size on the maximum attainable stabilized altitude. A test altitude of approx. 95,000 ft was attained, and altitudes in excess of 100,000 ft was attained.

Thin film deposition using rarefied gas jet

NASA Astrophysics Data System (ADS)

Pradhan, Sahadev, , Dr.

2017-01-01

The rarefied gas jet of aluminium is studied at Mach number Ma =(U_j /√{ kbTj / m }) in the range .01

Isolated Performance at Mach Numbers From 0.60 to 2.86 of Several Expendable Nozzle Concepts for Supersonic Applications

NASA Technical Reports Server (NTRS)

Re, Richard J.; Berrier, Bobby L.; Abeyounis, William K.

2001-01-01

Investigations have been conducted in the Langley 16-Foot Transonic Tunnel (at Mach numbers from 0.60 to 1.25) and in the Langley Unitary Plan Wind Tunnel (at Mach numbers from 2.16 to 2.86) at an angle of attack of 0 deg to determine the isolated performance of several expendable nozzle concepts for supersonic nonaugmented turbojet applications. The effects of centerbody base shape, shroud length, shroud ventilation, cruciform shroud expansion ratio, and cruciform shroud flap vectoring were investigated. The nozzle pressure ratio range, which was a function of Mach number, was between 1.9 and 11.8 in the 16-Foot Transonic Tunnel and between 7.9 and 54.9 in the Unitary Plan Wind Tunnel. Discharge coefficient, thrust-minus-drag, and the forces and moments generated by vectoring the divergent shroud flaps (for Mach numbers of 0.60 to 1.25 only) of a cruciform nozzle configuration were measured. The shortest nozzle had the best thrust-minus-drag performance at Mach numbers up to 0.95 but was approached in performance by other configurations at Mach numbers of 1.15 and 1.25. At Mach numbers above 1.25, the cruciform nozzle configuration having the same expansion ratio (2.64) as the fixed geometry nozzles had the best thrust-minus-drag performance. Ventilation of the fixed geometry divergent shrouds to the nozzle external boattail flow generally improved thrust-minus-drag performance at Mach numbers from 0.60 to 1.25, but decreased performance above a Mach number of 1.25.

Surface pressure data on a series of analytic forebodies at Mach numbers from 1.70 to 4.50 and combined angles of attack and sideslip. [Langley Unitary Plan wind tunnel

NASA Technical Reports Server (NTRS)

Townsend, J. C.; Howell, D. T.; Collins, I. K.; Hayes, C.

1979-01-01

Tabulated surface pressure data for a series of four forebodies which have analytically defined cross sections and which are based on a parabolic arc profile having a 20 deg half angle at the nose are presented without analysis. The first forebody has a circular cross section, and the second has a cross section which is an ellipse with an axis ratio of 2/1. The third has a cross section defined by a lobed analytic curve. The fourth forebody has cross sections which develop smoothly from circular at the pointed nose through the lobed analytic curve and back to circular at the aft end. The data generally cover angles of attack from -5 deg to 20 deg at angles of sideslip from 0 deg to 5 deg for Mach numbers of 1.70, 2.50, 3.95, and 4.50 at a constant Reynolds number.

Noncoplanar component of the magnetic field at low Mach number shocks

NASA Technical Reports Server (NTRS)

Friedman, M. A.; Russell, C. T.; Gosling, J. T.; Thomsen, M. F.

1990-01-01

The component of the magnetic field that deviates from the plane defined by the shock normal and the upstream magnetic field is examined for low Mach number bow shocks. The integrated value of this noncoplanar component is compared to the predictions of Jones and Ellison (1987). A test of this relationship was first reported by Gosling et al. (1988) who found good agreement only at the two low Mach number shocks that were included in their study. Analysis of a more extensive collection of low Mach number shocks confirms the Jones and Ellison relationship at very low Mach numbers as well as its deterioration for higher Mach numbers. However, there also is an indication that the relationship may break down for shocks that are nearly perpendicular.

HIGH STAR FORMATION RATES IN TURBULENT ATOMIC-DOMINATED GAS IN THE INTERACTING GALAXIES IC 2163 AND NGC 2207

DOE Office of Scientific and Technical Information (OSTI.GOV)

Elmegreen, Bruce G.; Kaufman, Michele; Bournaud, Frédéric

CO observations of the interacting galaxies IC 2163 and NGC 2207 are combined with HI, H α , and 24 μ m observations to study the star formation rate (SFR) surface density as a function of the gas surface density. More than half of the high-SFR regions are HI dominated. When compared to other galaxies, these HI-dominated regions have excess SFRs relative to their molecular gas surface densities but normal SFRs relative to their total gas surface densities. The HI-dominated regions are mostly located in the outer part of NGC 2207 where the HI velocity dispersion is high, 40–50 kmmore » s{sup −1}. We suggest that the star-forming clouds in these regions have envelopes at lower densities than normal, making them predominantly atomic, and cores at higher densities than normal because of the high turbulent Mach numbers. This is consistent with theoretical predictions of a flattening in the density probability distribution function for compressive, high Mach number turbulence.« less

Experimental Results on the Feasibility of an Aerospike for Hypersonic Missiles

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Mitchell, Anthony M.; Boudreaux, Ellis J.

1995-01-01

A series of wind tunnel tests have been performed on an aerospike-protected missile dome at a Mach number of 6 to obtain quantitative surface pressure and temperature-rise data, as well as qualitative flow visualization data. These data were used to determine aerospike concept feasibility and will also provide a database to be used for calibration of computational fluid dynamics codes. Data were obtained on the hemispherical missile dome with and without an aerospike that protrudes ahead of the dome along the axisymmetric center line. Data were obtained on two models (one pressure, one temperature) in the NASA Langley 20-Inch Mach 6 Tunnel at a freestream Reynolds number of 8.0 x 10(exp 6) per feet and angles of attack from 0 to 40 degrees. Surface pressure and temperature-rise results indicate that the aerospike is effective for very low angles of attack (less than 5 degrees) at Mach 6. Above 5 degrees, impingement of the aerospike bow shock and the flow separation shock from the recirculation region created by the aerospike causes pressure and temperature increases on the windward side of the dome which exceed values observed in the same region with the aerospike removed. Flow characterization obtained via oil-flow and schlieren photographs provides some insight into the quantitative surface data results, including vortical flow and shock-wave impingement.

Aerodynamic characteristics of a Sparrow 3 missile model in the flow field of a generalized parent body at Mach 2.86

NASA Technical Reports Server (NTRS)

Stallings, R. L., Jr.

1984-01-01

Longitudinal aerodynamic characteristics of a Sparrow 3 wing control missile model were measured through a range of separation distances relative to a flat plate surface that represented the parent-body configuration. Measurements were obtained with and without two dimensional circular arc protuberances attached to the flat plate surface. The tests were conducted at a Mach number of 2.86 and a Reynolds number per meter of 6.56 million. The behavior of these longitudinal characteristics with varying separation distance in the flow field created by the flat plate and protuberance was generally as would be expected on the basis of flow field boundaries determined from the second order approximation of Friedrich. In general, varying roll angle from 0 deg to 45 deg caused no significant effect on the store separation characteristics.

Investigation of two-stage air-cooled turbine suitable for flight at Mach number of 2.5 II : blade design

NASA Technical Reports Server (NTRS)

Miser, James W; Stewart, Warner L

1957-01-01

A blade design study is presented for a two-stage air-cooled turbine suitable for flight at a Mach number of 2.5 for which velocity diagrams have been previously obtained. The detailed procedure used in the design of the blades is given. In addition, the design blade shapes, surface velocity distributions, inner and outer wall contours, and other design data are presented. Of all the blade rows, the first-stage rotor has the highest solidity, with a value of 2.289 at the mean section. The second-stage stator also had a high mean-section solidity of 1.927, mainly because of its high inlet whirl. The second-stage rotor has the highest value of the suction-surface diffusion parameter, with a value of 0.151. All other blade rows have values for this parameter under 0.100.

Preliminary study of effects of winglets on wing flutter

NASA Technical Reports Server (NTRS)

Doggett, R. V., Jr.; Farmer, M. G.

1976-01-01

Some experimental flutter results are presented over a Mach number range from about 0.70 to 0.95 for a simple, swept, tapered, flat-plate wing model having a planform representative of subsonic transport airplanes and for the same wing model equipped with two different upper surface winglets. Both winglets had the same planform and area (about 2 percent of the basic-wing area); however, one weighed about 0.3 percent of the basic-wing weight, and the other weighed about 1.8 percent of the wing weight. The addition of the lighter winglet reduced the wing-flutter dynamic pressure by about 3 percent; the heavier winglet reduced the wing-flutter dynamic pressure by about 12 percent. The experimental flutter results are compared at a Mach number of 0.80 with analytical flutter results obtained by using doublet-lattice and lifting-surface (kernel-function) unsteady aerodynamic theories.

Unsteady aerodynamics of an oscillating cascade in a compressible flow field

NASA Technical Reports Server (NTRS)

Buffum, Daniel H.; Boldman, Donald R.; Fleeter, Sanford

1987-01-01

Fundamental experiments were performed in the NASA Lewis Transonic Oscillating Cascade Facility to investigate and quantify the unsteady aerodynamics of a cascade of biconvex airfoils executing torsion-mode oscillations at realistic reduced frequencies. Flush-mounted, high-response miniature pressure transducers were used to measure the unsteady airfoil surface pressures. The pressures were measured for three interblade phase angles at two inlet Mach numbers, 0.65 and 0.80, and two incidence angles, 0 and 7 deg. The time-variant pressures were analyzed by means of discrete Fourier transform techniques, and these unique data were then compared with predictions from a linearized unsteady cascade model. The experimental results indicate that the interblade phase angle had a major effect on the chordwise distributions of the airfoil surface unsteady pressure, and that reduced frequency, incidence angle, and Mach number had a somewhat less significant effect.

Oil-flow separation patterns on an ogive forebody

NASA Technical Reports Server (NTRS)

Keener, E. R.

1981-01-01

Oil flow patterns on a symmetric tangent ogive forebody having a fineness ratio of 3.5 are presented for angles of attack up to 88 deg at a transitional Reynolds number of 8 million (based on base diameter) and a Mach number of 0.25. Results show typical surface flow separation patterns, the magnitude of surface flow angles, and the extent of laminar and turbulent flow for symmetric, asymmetric, and wakelike flow regimes.

Effects of Cavities and Protuberances on Transition over Hypersonic Vehicles

NASA Technical Reports Server (NTRS)

Chang, Chau-Lyan; Choudhari, Meelan M.; Li, Fei; Venkatachari, Balaji

2011-01-01

Surface protuberances and cavities on a hypersonic vehicle are known to cause several aerodynamic or aerothermodynamic issues. Most important of all, premature transition due to these surface irregularities can lead to a significant rise in surface heating. To help understand laminar-turbulent transition induced by protuberances or cavities on a Crew Exploration Vehicle (CEV) surface, high-fidelity numerical simulations are carried out for both types of trips on a CEV wind tunnel model. Due to the large bluntness, these surface irregularities reside in an accelerating subsonic boundary layer. For the Mach 6 wind tunnel conditions with a roughness Reynolds number Re(sub kk) of 800, it was found that a protuberance with a height to boundary layer thickness ratio of 0.73 leads to strong wake instability and spontaneous vortex shedding, while a cavity with identical geometry only causes a rather weak flow unsteadiness. The same cavity with a larger Reynolds number also leads to similar spontaneous vortex shedding and wake instability. The wake development and the formation of hairpin vortices for both protuberance and cavity were found to be qualitatively similar to that observed for an isolated hemisphere submerged in a subsonic, low speed flat-plate boundary layer. However, the shed vortices and their accompanying instability waves were found to be slightly stabilized downstream by the accelerating boundary layer along the CEV surface. Despite this stabilizing influence, it was found that the wake instability spreads substantially in both wall-normal and azimuthal directions as the flow is evolving towards a transitional state. Similarities and differences between the wake instability behind a protuberance and a cavity are investigated. Computations for the Mach 6 boundary layer over a slender cylindrical roughness element with a height to the boundary layer thickness of about 1.1 also shows spontaneous vortex shedding and strong wake instability. Comparisons of detailed flow structures associated with protuberances at subsonic and supersonic edge Mach numbers indicate distinctively different instability mechanisms.

Local flow measurements at the inlet spike tip of a Mach 3 supersonic cruise airplane

NASA Technical Reports Server (NTRS)

Johnson, H. J.; Montoya, E. J.

1973-01-01

The flow field at the left inlet spike tip of a YF-12A airplane was examined using at 26 deg included angle conical flow sensor to obtain measurements at free-stream Mach numbers from 1.6 to 3.0. Local flow angularity, Mach number, impact pressure, and mass flow were determined and compared with free-stream values. Local flow changes occurred at the same time as free-stream changes. The local flow usually approached the spike centerline from the upper outboard side because of spike cant and toe-in. Free-stream Mach number influenced the local flow angularity; as Mach number increased above 2.2, local angle of attack increased and local sideslip angle decreased. Local Mach number was generally 3 percent less than free-stream Mach number. Impact-pressure ratio and mass flow ratio increased as free-stream Mach number increased above 2.2, indicating a beneficial forebody compression effect. No degradation of the spike tip instrumentation was observed after more than 40 flights in the high-speed thermal environment encountered by the airplane. The sensor is rugged, simple, and sensitive to small flow changes. It can provide accurate imputs necessary to control an inlet.

Courant Number and Mach Number Insensitive CE/SE Euler Solvers

NASA Technical Reports Server (NTRS)

Chang, Sin-Chung

2005-01-01

It has been known that the space-time CE/SE method can be used to obtain ID, 2D, and 3D steady and unsteady flow solutions with Mach numbers ranging from 0.0028 to 10. However, it is also known that a CE/SE solution may become overly dissipative when the Mach number is very small. As an initial attempt to remedy this weakness, new 1D Courant number and Mach number insensitive CE/SE Euler solvers are developed using several key concepts underlying the recent successful development of Courant number insensitive CE/SE schemes. Numerical results indicate that the new solvers are capable of resolving crisply a contact discontinuity embedded in a flow with the maximum Mach number = 0.01.

CFD-Based Design Optimization Tool Developed for Subsonic Inlet

NASA Technical Reports Server (NTRS)

1995-01-01

The traditional approach to the design of engine inlets for commercial transport aircraft is a tedious process that ends with a less-than-optimum design. With the advent of high-speed computers and the availability of more accurate and reliable computational fluid dynamics (CFD) solvers, numerical optimization processes can effectively be used to design an aerodynamic inlet lip that enhances engine performance. The designers' experience at Boeing Corporation showed that for a peak Mach number on the inlet surface beyond some upper limit, the performance of the engine degrades excessively. Thus, our objective was to optimize efficiency (minimize the peak Mach number) at maximum cruise without compromising performance at other operating conditions. Using a CFD code NPARC, the NASA Lewis Research Center, in collaboration with Boeing, developed an integrated procedure at Lewis to find the optimum shape of a subsonic inlet lip and a numerical optimization code, ADS. We used a GRAPE-based three-dimensional grid generator to help automate the optimization procedure. The inlet lip shape at the crown and the keel was described as a superellipse, and the superellipse exponents and radii ratios were considered as design variables. Three operating conditions: cruise, takeoff, and rolling takeoff, were considered in this study. Three-dimensional Euler computations were carried out to obtain the flow field. At the initial design, the peak Mach numbers for maximum cruise, takeoff, and rolling takeoff conditions were 0.88, 1.772, and 1.61, respectively. The acceptable upper limits on the takeoff and rolling takeoff Mach numbers were 1.55 and 1.45. Since the initial design provided by Boeing was found to be optimum with respect to the maximum cruise condition, the sum of the peak Mach numbers at takeoff and rolling takeoff were minimized in the current study while the maximum cruise Mach number was constrained to be close to that at the existing design. With this objective, the optimum design satisfied the upper limits at takeoff and rolling takeoff while retaining the desirable cruise performance. Further studies are being conducted to include static and cross-wind operating conditions in the design optimization procedure. This work was carried out in collaboration with Dr. E.S. Reddy of NYMA, Inc.

Theoretical Calculations of Supersonic Wave Drag at Zero Lift for a Particular Store Arrangement

NASA Technical Reports Server (NTRS)

Margolis, Kenneth; Malvestuto, Frank S , Jr; Maxie, Peter J , Jr

1958-01-01

An analysis, based on the linearized thin-airfoil theory for supersonic speeds, of the wave drag at zero lift has been carried out for a simple two-body arrangement consisting of two wedgelike surfaces, each with a rhombic lateral cross section and emanating from a common apex. Such an arrangement could be used as two stores, either embedded within or mounted below a wing, or as auxiliary bodies wherein the upper halves could be used as stores and the lower halves for bomb or missile purposes. The complete range of supersonic Mach numbers has been considered and it was found that by orienting the axes of the bodies relative to each other a given volume may be redistributed in a manner which enables the wave drag to be reduced within the lower supersonic speed range (where the leading edge is substantially subsonic). At the higher Mach numbers, the wave drag is always increased. If, in addition to a constant volume, a given maximum thickness-chord ratio is imposed, then canting the two surfaces results in higher wave drag at all Mach numbers. For purposes of comparison, analogous drag calculations for the case of two parallel winglike bodies with the same cross-sectional shapes as the canted configuration have been included. Consideration is also given to the favorable (dragwise) interference pressures acting on the blunt bases of both arrangements.

Variation with Mach Number of Static and Total Pressures Through Various Screens

NASA Technical Reports Server (NTRS)

Adler, Alfred A

1946-01-01

Tests were conducted in the Langley 24-inch highspeed tunnel to ascertain the static-pressure and total-pressure losses through screens ranging in mesh from 3 to 12 wires per inch and in wire diameter from 0.023 to 0.041 inch. Data were obtained from a Mach number of approximately 0.20 up to the maximum (choking) Mach number obtainable for each screen. The results of this investigation indicate that the pressure losses increase with increasing Mach number until the choking Mach number, which can be computed, is reached. Since choking imposes a restriction on the mass rate of flow and maximum losses are incurred at this condition, great care must be taken in selecting the screen mesh and wire dimmeter for an installation so that the choking Mach number is

Geometric Effects on the Amplification of First Mode Instability Waves

NASA Technical Reports Server (NTRS)

Kirk, Lindsay C.; Candler, Graham V.

2013-01-01

The effects of geometric changes on the amplification of first mode instability waves in an external supersonic boundary layer were investigated using numerical techniques. Boundary layer stability was analyzed at Mach 6 conditions similar to freestream conditions obtained in quiet ground test facilities so that results obtained in this study may be applied to future test article design to measure first mode instability waves. The DAKOTA optimization software package was used to optimize an axisymmetric geometry to maximize the amplification of the waves at first mode frequencies as computed by the 2D STABL hypersonic boundary layer stability analysis tool. First, geometric parameters such as nose radius, cone half angle, vehicle length, and surface curvature were examined separately to determine the individual effects on the first mode amplification. Finally, all geometric parameters were allowed to vary to produce a shape optimized to maximize the amplification of first mode instability waves while minimizing the amplification of second mode instability waves. Since first mode waves are known to be most unstable in the form of oblique wave, the geometries were optimized using a broad range of wave frequencies as well as a wide range of oblique wave angles to determine the geometry that most amplifies the first mode waves. Since first mode waves are seen most often in flows with low Mach numbers at the edge of the boundary layer, the edge Mach number for each geometry was recorded to determine any relationship between edge Mach number and the stability of first mode waves. Results indicate that an axisymmetric cone with a sharp nose and a slight flare at the aft end under the Mach 6 freestream conditions used here will lower the Mach number at the edge of the boundary layer to less than 4, and the corresponding stability analysis showed maximum first mode N factors of 3.

Impingement of Boundary-Reflected Disturbances Originating at the Nose of a Body of Revolution in the Langley Research Center 16-Foot Transonic Tunnel

NASA Technical Reports Server (NTRS)

Re, Richard, J.; Capone, Francis J.

1998-01-01

An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to determine boundary-reflected disturbance lengths at low supersonic Mach numbers in the octagonally shaped test section. A body of revolution that had a nose designed to produce a bow shock and flow field similar to that about the nose of a supersonic transport configuration was used. The impingement of reflected disturbances on the model was determined from static pressures measured on the surface of the model. Test variables included Mach number (0.90 to 1.25), model angle of attack (nominally -10, 0, and 10), and model roll angle.

Upper wing surface boundary layer measurements and static aerodynamic data obtained on a 0.015-scale model (42-0) or the SSV orbiter configuration 140A/B in the LTV HSWT at a Mach number of 4.6 (LA58)

NASA Technical Reports Server (NTRS)

Ball, J. W.; Lindahl, R. H.

1976-01-01

The purpose of the test was to investigate the nature of the Orbiter boundary layer characteristics at angles of attack from -4 to 32 degrees at a Mach number of 4.6. The effect of large grit, employed as transition strips, on both the nature of the boundary layer and the force and moment characteristics were investigated along with the effects of large negative elevon deflection on lee side separation. In addition, laminar and turbulent boundary layer separation phenomena which could cause asymmetric flow separation were investigated.

An oscillatory kernel function method for lifting surfaces in mixed transonic flow

NASA Technical Reports Server (NTRS)

Cunningham, A. M., Jr.

1974-01-01

A study was conducted on the use of combined subsonic and supersonic linear theory to obtain economical and yet realistic solutions to unsteady transonic flow problems. With some modification, existing linear theory methods were combined into a single computer program. The method was applied to problems for which measured steady Mach number distributions and unsteady pressure distributions were available. By comparing theory and experiment, the transonic method showed a significant improvement over uniform flow methods. The results also indicated that more exact local Mach number effects and normal shock boundary conditions on the perturbation potential were needed. The validity of these improvements was demonstrated by application to steady flow.

Heat transfer investigation of Langley Research Center transition models at a Mach number of 8, volume 2

NASA Technical Reports Server (NTRS)

Matthews, R. K.; Martindale, W. R.; Warmbrod, J. D.; Johnson, C. B.

1972-01-01

The results are presented of a wind tunnel test program to determine aerodynamic heat transfer distributions on delta body and straight body transition models of the space shuttle. Heat transfer rates were determined by the phase-change paint technique on Stycast and RTV models using Tempilag as the surface temperature indicator. The nominal test conditions were: Mach 8, length Reynolds numbers of 5 million and 7.4 million, and angles of attack of 20, 40, and 60 deg. Model details, test conditions, and reduced heat transfer data are included. Data reduction of the phase-change paint photographs was performed by utilizing a new technique.

Development and flight test results of an autothrottle control system at Mach 3 cruise

NASA Technical Reports Server (NTRS)

Gilyard, G. B.; Burken, J. J.

1980-01-01

Flight test results obtained with the original Mach hold autopilot designed the YF-12C airplane which uses elevator control and a newly developed Mach hold system having an autothrottle integrated with an altitude hold autopilot system are presented. The autothrottle tests demonstrate good speed control at high Mach numbers and high altitudes while simultaneously maintaining control over altitude and good ride qualities. The autothrottle system was designed to control either Mach number or knots equivalent airspeed (KEAS). Excellent control of Mach number or KEAS was obtained with the autothrottle system when combined with altitude hold. Ride qualities were significantly better than with the conventional Mach hold system.

Investigation of wing upper surface flow-field disturbance due to NASA DC-8-72 in-flight inboard thrust-reverser deployment

NASA Technical Reports Server (NTRS)

Hamid, Hedayat U.; Margason, Richard J.; Hardy, Gordon

1995-01-01

An investigation of the wing upper surface flow-field disturbance due to in-flight inboard thrust reverser deployment on the NASA DC-8-72, which was conducted cooperatively by NASA Ames, the Federal Aviation Administration (FAA), McDonnell Douglas, and the Aerospace Industry Association (AIA), is outlined and discussed in detail. The purpose of this flight test was to obtain tufted flow visualization data which demonstrates the effect of thrust reverser deployment on the wing upper surface flow field to determine if the disturbed flow regions could be modeled by computational methods. A total of six symmetric thrust reversals of the two inboard engines were performed to monitor tuft and flow cone patterns as well as the character of their movement at the nominal Mach numbers of 0.55, 0.70, and 0.85. The tufts and flow cones were photographed and video-taped to determine the type of flow field that occurs with and without the thrust reversers deployed. In addition, the normal NASA DC-8 onboard Data Acquisition Distribution System (DADS) was used to synchronize the cameras. Results of this flight test will be presented in two parts. First, three distinct flow patterns associated with the above Mach numbers were sketched from the motion videos and discussed in detail. Second, other relevant aircraft parameters, such as aircraft's angular orientation, altitude, Mach number, and vertical descent, are discussed. The flight test participants' comments were recorded on the videos and the interested reader is referred to the video supplement section of this report for that information.

Wind tunnel investigations of forebody strakes for yaw control on F/A-18 model at subsonic and transonic speeds

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Murri, Daniel G.

1993-01-01

Wind tunnel investigations have been conducted of forebody strakes for yaw control on 0.06-scale models of the F/A-18 aircraft at free-stream Mach numbers of 0.20 to 0.90. The testing was conducted in the 7- by 10-Foot Transonic Tunnel at the David Taylor Research Center and the Langley 7- by 10-Foot High-Speed Tunnel. The principal objectives of the testing were to determine the effects of the Mach number and the strake plan form on the strake yaw control effectiveness and the corresponding strake vortex induced flow field. The wind tunnel model configurations simulated an actuated conformal strake deployed for maximum yaw control at high angles of attack. The test data included six-component forces and moments on the complete model, surface static pressure distributions on the forebody and wing leading-edge extensions, and on-surface and off-surface flow visualizations. The results from these studies show that the strake produces large yaw control increments at high angles of attack that exceed the effect of conventional rudders at low angles of attack. The strake yaw control increments diminish with increasing Mach number but continue to exceed the effect of rudder deflection at angles of attack greater than 30 degrees. The character of the strake vortex induced flow field is similar at subsonic and transonic speeds. Cropping the strake planform to account for geometric and structural constraints on the F-18 aircraft has a small effect on the yaw control increments at subsonic speeds and no effect at transonic speeds.

Unsteady pressure measurements on a biconvex airfoil in a transonic oscillating cascade

NASA Technical Reports Server (NTRS)

Shaw, L. M.; Boldman, D. R.; Buggele, A. E.; Buffum, D. H.

1985-01-01

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfroils were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence and angle of 0 deg and A Mach number of 0.65 in order to obtain results in a shock-free compressible flowfield. Subsequent tests were performed at an incidence angle of 7 deg and Mach number of 0.8 in order to observe the surface pressures with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and -90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semi-chord reduced frequencies up to about 0.5 at a Mach number of 0.8). Results from the zero-incidence cascade are compared with a classical unsteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

Simulation Analysis of Fluid-Structure Interaction of High Velocity Environment Influence on Aircraft Wing Materials under Different Mach Numbers.

PubMed

Zhang, Lijun; Sun, Changyan

2018-04-18

Aircraft service process is in a state of the composite load of pressure and temperature for a long period of time, which inevitably affects the inherent characteristics of some components in aircraft accordingly. The flow field of aircraft wing materials under different Mach numbers is simulated by Fluent in order to extract pressure and temperature on the wing in this paper. To determine the effect of coupling stress on the wing’s material and structural properties, the fluid-structure interaction (FSI) method is used in ANSYS-Workbench to calculate the stress that is caused by pressure and temperature. Simulation analysis results show that with the increase of Mach number, the pressure and temperature on the wing’s surface both increase exponentially and thermal stress that is caused by temperature will be the main factor in the coupled stress. When compared with three kinds of materials, titanium alloy, aluminum alloy, and Haynes alloy, carbon fiber composite material has better performance in service at high speed, and natural frequency under coupling pre-stressing will get smaller.

Effects of Wing Sweep on In-flight Boundary-layer Transition for a Laminar Flow Wing at Mach Numbers from 0.60 to 0.79

NASA Technical Reports Server (NTRS)

Anderson, Bianca Trujillo; Meyer, Robert R., Jr.

1990-01-01

The variable sweep transition flight experiment (VSTFE) was conducted on an F-14A variable sweep wing fighter to examine the effect of wing sweep on natural boundary layer transition. Nearly full span upper surface gloves, extending to 60 percent chord, were attached to the F-14 aircraft's wings. The results are presented of the glove 2 flight tests. Glove 2 had an airfoil shape designed for natural laminar flow at a wing sweep of 20 deg. Sample pressure distributions and transition locations are presented with the complete results tabulated in a database. Data were obtained at wing sweeps of 15, 20, 25, 30, and 35 deg, at Mach numbers ranging from 0.60 to 0.79, and at altitudes ranging from 10,000 to 35,000 ft. Results show that a substantial amount of laminar flow was maintained at all the wing sweeps evaluated. The maximum transition Reynolds number obtained was 18.6 x 10(exp 6) at 15 deg of wing sweep, Mach 0.75, and at an altitude of 10,000 ft.

Simulation Analysis of Fluid-Structure Interaction of High Velocity Environment Influence on Aircraft Wing Materials under Different Mach Numbers

PubMed Central

Sun, Changyan

2018-01-01

Aircraft service process is in a state of the composite load of pressure and temperature for a long period of time, which inevitably affects the inherent characteristics of some components in aircraft accordingly. The flow field of aircraft wing materials under different Mach numbers is simulated by Fluent in order to extract pressure and temperature on the wing in this paper. To determine the effect of coupling stress on the wing’s material and structural properties, the fluid-structure interaction (FSI) method is used in ANSYS-Workbench to calculate the stress that is caused by pressure and temperature. Simulation analysis results show that with the increase of Mach number, the pressure and temperature on the wing’s surface both increase exponentially and thermal stress that is caused by temperature will be the main factor in the coupled stress. When compared with three kinds of materials, titanium alloy, aluminum alloy, and Haynes alloy, carbon fiber composite material has better performance in service at high speed, and natural frequency under coupling pre-stressing will get smaller. PMID:29670023

Numerical Solution of the Flow of a Perfect Gas Over A Circular Cylinder at Infinite Mach Number

NASA Technical Reports Server (NTRS)

Hamaker, Frank M.

1959-01-01

A solution for the two-dimensional flow of an inviscid perfect gas over a circular cylinder at infinite Mach number is obtained by numerical methods of analysis. Nonisentropic conditions of curved shock waves and vorticity are included in the solution. The analysis is divided into two distinct regions, the subsonic region which is analyzed by the relaxation method of Southwell and the supersonic region which was treated by the method of characteristics. Both these methods of analysis are inapplicable on the sonic line which is therefore considered separately. The shapes of the sonic line and the shock wave are obtained by iteration techniques. The striking result of the solution is the strong curvature of the sonic line and of the other lines of constant Mach number. Because of this the influence of the supersonic flow on the sonic line is negligible. On comparison with Newtonian flow methods, it is found that the approximate methods show a larger variation of surface pressure than is given by the present solution.

Simulating compressible-incompressible two-phase flows

NASA Astrophysics Data System (ADS)

Denner, Fabian; van Wachem, Berend

2017-11-01

Simulating compressible gas-liquid flows, e.g. air-water flows, presents considerable numerical issues and requires substantial computational resources, particularly because of the stiff equation of state for the liquid and the different Mach number regimes. Treating the liquid phase (low Mach number) as incompressible, yet concurrently considering the gas phase (high Mach number) as compressible, can improve the computational performance of such simulations significantly without sacrificing important physical mechanisms. A pressure-based algorithm for the simulation of two-phase flows is presented, in which a compressible and an incompressible fluid are separated by a sharp interface. The algorithm is based on a coupled finite-volume framework, discretised in conservative form, with a compressive VOF method to represent the interface. The bulk phases are coupled via a novel acoustically-conservative interface discretisation method that retains the acoustic properties of the compressible phase and does not require a Riemann solver. Representative test cases are presented to scrutinize the proposed algorithm, including the reflection of acoustic waves at the compressible-incompressible interface, shock-drop interaction and gas-liquid flows with surface tension. Financial support from the EPSRC (Grant EP/M021556/1) is gratefully acknowledged.

Comparison of jet Mach number decay data with a correlation and jet spreading contours for a large variety of nozzles

NASA Technical Reports Server (NTRS)

Groesbeck, D. E.; Huff, R. G.; Vonglahn, U. H.

1977-01-01

Small-scale circular, noncircular, single- and multi-element nozzles with flow areas as large as 122 sq cm were tested with cold airflow at exit Mach numbers from 0.28 to 1.15. The effects of multi-element nozzle shape and element spacing on jet Mach number decay were studied in an effort to reduce the noise caused by jet impingement on externally blown flap (EBF) STOL aircraft. The jet Mach number decay data are well represented by empirical relations. Jet spreading and Mach number decay contours are presented for all configurations tested.

Evaluation of a Quartz Bourdon Pressure Gage of Wind Tunnel Mach Number Control System Application

NASA Technical Reports Server (NTRS)

Chapin, W. G.

1986-01-01

A theoretical and experimental study was undertaken to determine the feasibility of using the National Transonic Facility's high accuracy Mach number measurement system as part of a closed loop Mach number control system. The theoretical and experimental procedures described are applicable to the engineering design of pressure control systems. The results show that the dynamic response characteristics of the NTF Mach number gage (a Ruska DDR-6000 quartz absolute pressure gage) coupled to a typical length of pressure tubing were only marginally acceptable within a limited range of the facility's total pressure envelope and could not be used in the Mach number control system.

Chaotic behaviour of high Mach number flows

NASA Technical Reports Server (NTRS)

Varvoglis, H.; Ghosh, S.

1985-01-01

The stability of the super-Alfvenic flow of a two-fluid plasma model with respect to the Mach number and the angle between the flow direction and the magnetic field is investigated. It is found that, in general, a large scale chaotic region develops around the initial equilibrium of the laminar flow when the Mach number exceeds a certain threshold value. After reaching a maximum the size of this region begins shrinking and goes to zero as the Mach number tends to infinity. As a result high Mach number flows in time independent astrophysical plasmas may lead to the formation of 'quasi-shocks' in the presence of little or no dissipation.

Heat transfer rate distributions on McDonnell-Douglas booster determined by phase change technique for nominal Mach number of 8

NASA Technical Reports Server (NTRS)

Matthews, R. K.; Martindale, W. R.; Warmbrod, J. D.

1972-01-01

The results of a wind tunnel test program to determine aerodynamic heat transfer distributions on the McDonnell Douglas Booster configuration are presented. Heat-transfer rates were determined by the phase-change paint technique on 0.009-scale Stycast models using Tempilaq as the surface temperature indicator. The nominal test conditions were; Mach 8, length Reynolds numbers 5 million and 7.3 million, and angles of attack of 40, 50, and 60 deg. At the higher Reynolds number, data were obtained with and without boundary layer trips. Model details, test conditions, and reduced heat-transfer data are presented. Data reduction of the phase-change paint photographs was performed by utilizing a new technique which is described.

Boundary layer, skin friction, and boattail pressure measurements from the YF-12 airplane at Mach numbers up to 3

NASA Technical Reports Server (NTRS)

Fisher, D. F.

1978-01-01

In-flight measurements of boundary layer and skin friction data were made on YF-12 airplanes for Mach numbers between 2.0 and 3.0. Boattail pressures were also obtained for Mach numbers between 0.7 and 3.0 with Reynolds numbers up to four hundred million. Boundary layer data measured along the lower fuselage centerline indicate local displacement and momentum thicknesses can be much larger than predicted. Skin friction coefficients measured at two of five lower fuselage stations were significantly less than predicted by flat plate theory. The presence of large differences between measured boattail pressure drag and values calculated by a potential flow solution indicates the presence of vortex effects on the upper boattail surface. At both subsonic and supersonic speeds, pressure drag on the longer of two boattail configurations was equal to or less than the pressure drag on the shorter configuration. At subsonic and transonic speeds, the difference in the drag coefficient was on the order of 0.0008 to 0.0010. In the supersonic cruise range, the difference in the drag coefficient was on the order of 0.002. Boattail drag coefficients are based on wing reference area.

Numerical Analysis of the Trailblazer Inlet Flowfield for Hypersonic Mach Numbers

NASA Technical Reports Server (NTRS)

Steffen, C. J., Jr.; DeBonis, J. R.

1999-01-01

A study of the Trailblazer vehicle inlet was conducted using the Global Air Sampling Program (GASP) code for flight Mach numbers ranging from 4-12. Both perfect gas and finite rate chemical analysis were performed with the intention of making detailed comparisons between the two results. Inlet performance was assessed using total pressure recovery and kinetic energy efficiency. These assessments were based upon a one-dimensional stream-thrust-average of the axisymmetric flowfield. Flow visualization utilized to examine the detailed shock structures internal to this mixed-compression inlet. Kinetic energy efficiency appeared to be the least sensitive to differences between the perfect gas and finite rate chemistry results. Total pressure recovery appeared to be the most sensitive discriminator between the perfect gas and finite rate chemistry results for flight Mach numbers above Mach 6. Adiabatic wall temperature was consistently overpredicted by the perfect gas model for flight Mach numbers above Mach 4. The predicted shock structures were noticeably different for Mach numbers from 6-12. At Mach 4, the perfect gas and finite rate chemistry models collapse to the same result.

14 CFR Appendix B to Part 91 - Authorizations To Exceed Mach 1 (§ 91.817)

Code of Federal Regulations, 2010 CFR

2010-01-01

..., including an environmental analysis of that area meeting the requirements of paragraph (b) of this section... reach the surface outside of the designated test area. (d) An application is denied if the Administrator... number of 1 will not cause a measurable sonic boom overpressure to reach the surface. (b) For a flight...

14 CFR Appendix B to Part 91 - Authorizations To Exceed Mach 1 (§ 91.817)

Code of Federal Regulations, 2011 CFR

2011-01-01

..., including an environmental analysis of that area meeting the requirements of paragraph (b) of this section... reach the surface outside of the designated test area. (d) An application is denied if the Administrator... number of 1 will not cause a measurable sonic boom overpressure to reach the surface. (b) For a flight...

Effect of gas injection on drag and surface heat transfer rates for a 30° semi-apex angle blunt body flying at Mach 5.75

NASA Astrophysics Data System (ADS)

Sahoo, N.; Kulkarni, V.; Jagadeesh, G.; Reddy, K. P. J.

Effect of coolant gas injection in the stagnation region on the surface heat transfer rates and aerodynamic drag for a large angle blunt body flying at hypersonic Mach number is reported for two stagnation enthalpies. A 60° apex-angle blunt cone model is employed for this purpose with air injection at the nose through a hole of 2mm diameter. The convective surface heating rates and aerodynamic drag are measured simultaneously using surface mounted platinum thin film sensors and internally mounted accelerometer balance system, respectively. About 35-40% reduction in surface heating rates is observed in the vicinity of stagnation region whereas 15-25% reduction in surface heating rates is felt beyond the stagnation region at stagnation enthalpy of 1.6MJ/kg. The aerodynamic drag expressed in terms of drag coefficient is found to increase by 20% due to the air injection.

Investigation of configuration effects on entry heating distributions at Mach no. equal 8.0 (OH41). [for wind tunnel model of space shuttle orbiter

NASA Technical Reports Server (NTRS)

Gorowitz, H.; White, R.; Derrico, A.

1973-01-01

Aerodynamic heating data were obtained on 0.006 scale models of four Rockwell International SSV double delta wing Orbiters in the Mach 8 variable density tunnel. A model of two previously tested Rockwell International Orbiters which are identified in the Configuration Description of this report were also tested. Orbiter surfaces were thermally mapped from the laminar through turbulent flight regimes during re-entry. Various modifications were made to model lower surfaces to determine the cause of transition in the vicinity of 3.0 million Reynolds number per foot. Re-entry data were acquired for angles of attack from 25 through 35 degrees at nominal Reynolds numbers per foot of 1.0, 2.0, 2.3, 2.5, 3.0, 3.5, 4.5 and 6.0 million utilizing the phase change paint technique. Launch data were acquired on the model upper surfaces for angles of attack of 0 and -5 degrees at nominal Reynolds numbers per foot of 3.0 and 6.0 million. A total of 70 orbiter heating runs and 6 material sample sphere runs were completed.

The Development of Cambered Airfoil Sections Having Favorable Lift Characteristics at Supercritical Mach Numbers

NASA Technical Reports Server (NTRS)

Graham, Donald J

1948-01-01

Several groups of new airfoil sections, designated as the NACA 8-series, are derived analytically to have lift characteristics at supercritical Mach numbers which are favorable in the sense that the abrupt loss of lift, characteristic of the usual airfoil section at Mach numbers above the critical, is avoided. Aerodynamic characteristics determined, from two-dimensional windtunnel tests at Mach numbers up to approximately 0.9 are presented for each of the derived airfoils. Comparisons are made between the characteristics of these airfoils and the corresponding characteristics of representative NPiCA 6-series airfoils. The experimental results confirm the design expectations in demonstrating for the NACA S-series airfoils either no variation, or an Increase from the low-speed design value, In the lift coefficient at a constant angle of attack with increasing Mach number above the critical. It was not found possible to improve the variation with Mach number of the slope of the lift curve for these airfoils above that for the NACA 6-series airfoils. The drag characteristics of the new airfoils are somewhat inferior to those of the NACA 6- series with respect to divergence with Mach number, but the pitching-moment characteristics are more favorable for the thinner new sections In demonstrating somewhat smaller variations of moment coefficient with both angle of attack and Mach number. The effect on the aero&ynamic characteristics at high Mach numbers of removing the cusp from the trailing-edge regions of two 10-percent-chord-thick NACA 6-series airfoils is determined to be negligible.

Aerodynamic Parameter Estimation for the X-43A (Hyper-X) from Flight Data

NASA Technical Reports Server (NTRS)

Morelli, Eugene A.; Derry, Stephen D.; Smith, Mark S.

2005-01-01

Aerodynamic parameters were estimated based on flight data from the third flight of the X-43A hypersonic research vehicle, also called Hyper-X. Maneuvers were flown using multiple orthogonal phase-optimized sweep inputs applied as simultaneous control surface perturbations at Mach 8, 7, 6, 5, 4, and 3 during the vehicle descent. Aerodynamic parameters, consisting of non-dimensional longitudinal and lateral stability and control derivatives, were estimated from flight data at each Mach number. Multi-step inputs at nearly the same flight conditions were also flown to assess the prediction capability of the identified models. Prediction errors were found to be comparable in magnitude to the modeling errors, which indicates accurate modeling. Aerodynamic parameter estimates were plotted as a function of Mach number, and compared with estimates from the pre-flight aerodynamic database, which was based on wind-tunnel tests and computational fluid dynamics. Agreement between flight estimates and values computed from the aerodynamic database was excellent overall.

Investigation of very low blockage ratio boattail models in the Langley 16-foot transonic tunnel

NASA Technical Reports Server (NTRS)

Reubush, D. E.

1976-01-01

An investigation at an angle of attack of 0 deg was conducted in a 16 foot transonic tunnel at Mach numbers from 0.4 to 1.05 to determine the limits in Mach number at which valid boattail pressure drag data may be obtained with very low blockage ratio bodies. Extreme care was exercised when examining any data taken at subsonic Mach numbers very near 1.0 and lower than the supersonic Mach number at which shock reflections miss the model. Boattail pressure coefficient distributions did not indicate any error, but when integrated boattail pressure drag data was plotted as a function of Mach number, data which were in error were identified.

Results of Two Free-fall Experiments on Flutter of Thin Unswept Wings in the Transonic Speed Range

NASA Technical Reports Server (NTRS)

Lauten, William T , Jr; Nelson, Herbert C

1957-01-01

Results of four thin, unswept, flutter airfoils attached to two freely falling bodies are reported. Two airfoils fluttered at a Mach number of 0.85, a third airfoil fluttered at a Mach number of 1.03, and a fourth fluttered at a Mach number of 1.07. Results of calculations of flutter speed using incompressible and compressible air-force coefficients, including a Mach number of 1.0, are presented.

Modelling two-way interactions between atmospheric pollution and weather using high-resolution GEM-MACH

NASA Astrophysics Data System (ADS)

Makar, Paul; Gong, Wanmin; Pabla, Balbir; Cheung, Philip; Milbrandt, Jason; Gravel, Sylvie; Moran, Michael; Gilbert, Samuel; Zhang, Junhua; Zheng, Qiong

2013-04-01

The Global Environmental Multiscale (GEM) model is the source of the Canadian government's operational numerical weather forecast guidance, and GEM-MACH is the Canadian operational air-quality forecast model. GEM-MACH comprises GEM and the 'Modelling Air-quality and Chemistry' module, a gas-phase, aqueous-phase and aerosol chemistry and microphysics subroutine package called from within GEM's physics module. The present operational GEM-MACH model is "on-line" (both chemistry and meteorology are part of the same modelling structure) but is not fully coupled (weather variables are provided as inputs to the chemistry, but the chemical variables are not used to modify the weather). In this work, we describe modifications made to GEM-MACH as part of the 2nd phase of the Air Quality Model Evaluation International Initiative, in order to bring the model to a fully coupled status and present the results of initial tests comparing uncoupled and coupled versions of the model to observations for a high-resolution forecasting system. Changes to GEM's cloud microphysics and radiative transfer packages were carried out to allow two-way coupling. The cloud microphysics package used here is the Milbrandt-Yau 2-moment (MY2) bulk microphysics scheme, which solves prognostic equations for the total droplet number concentration and the mass mixing ratios of six hydrometeor categories. Here, we have replaced the original cloud condensation nucleation parameterization of MY2 (empirically relating supersaturation and CCN number) with the aerosol activation scheme of Abdul-Razzak and Ghan (2002). The latter scheme makes use of the particle size and speciation distribution of GEM-MACH's chemistry code as well as meteorological inputs to predict the number of aerosol particles activated to form cloud droplets, which is then used in the MY2 microphysics. The radiative transfer routines of GEM assume a default constant concentration aerosol profile between the surface and 1500m, and a single set of optical properties for extinction, single scattering albedo, and asymmetry factor. Ozone in GEM is taken from a default 2D (latitude-height) monthly climatology. We have replaced the ozone below the model top with the ozone calculated from GEM-MACH's chemistry, and the default optical parameters associated with particulate matter have been replaced by those calculated with a Mie scattering algorithm. These changes were found to have a significant local impact on both weather and air-quality predictions for short-term test runs of 24 hours duration. In that particular case, the maximum number concentration of cloud droplets decreased by an order of magnitude, while the number of raindrops increased by an order of magnitude and changed in spatial distribution, but surface rainfall was found to decrease. The differences in meteorology had a profound effect on local pollutant plume concentrations at specific locations and times. We compare results over a longer time period, using two parallel forecast systems, one with feedbacks between meteorology and chemistry, one without. Both nest GEM-MACH from a North American domain (10 km horizontal grid spacing) to a 1535 x 1360 km, 2.5 km domain. These systems will be evaluated against monitoring networks within the high resolution domain.

Supersonic aerodynamic characteristics of a proposed Assured Crew Return Capability (ACRC) lifting-body configuration

NASA Technical Reports Server (NTRS)

Ware, George M.

1989-01-01

An investigation was conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers from 1.6 to 4.5. The model had a low-aspect-ratio body with a flat undersurface. A center fin and two outboard fins were mounted on the aft portion of the upper body. The outboard fins were rolled outboard 40 deg from the vertical. Elevon surfaces made up the trailing edges of the outboard fins, and body flaps were located on the upper and lower aft fuselage. The center fin pivoted about its midchord for yaw control. The model was longitudinally stable about the design center-of-gravity position at 54 percent of the body length. The configuration with undeflected longitudinal controls trimmed near 0 deg angle of attack at Mach numbers from 1.6 to 3.0 where lift and lift-drag ratio were negative. Longitudinal trim was near the maximum lift-drag ratio (1.4) at Mach 4.5. The model was directionally stable over Mach number range except at angles of attack around 4 deg at M = 2.5. Pitch control deflection of more than -10 deg with either elevons or body flaps is needed to trim the model to angles of attack at which lift becomes positive. With increased control deflection, the lifting-body configuration should perform the assured crew return mission through the supersonic speed range.

Aerodynamic Performance and Static Stability and Control of Flat-Top Hypersonic Gliders at Mach Numbers from 0.6 to 18

NASA Technical Reports Server (NTRS)

Syvertson, Clarence A; Gloria, Hermilo R; Sarabia, Michael F

1958-01-01

A study is made of aerodynamic performance and static stability and control at hypersonic speeds. In a first part of the study, the effect of interference lift is investigated by tests of asymmetric models having conical fuselages and arrow plan-form wings. The fuselage of the asymmetric model is located entirely beneath the wing and has a semicircular cross section. The fuselage of the symmetric model was centrally located and has a circular cross section. Results are obtained for Mach numbers from 3 to 12 in part by application of the hypersonic similarity rule. These results show a maximum effect of interference on lift-drag ratio occurring at Mach number of 5, the Mach number at which the asymmetric model was designed to exploit favorable lift interference. At this Mach number, the asymmetric model is indicated to have a lift-drag ratio 11 percent higher than the symmetric model and 15 percent higher than the asymmetric model when inverted. These differences decrease to a few percent at a Mach number of 12. In the course of this part of the study, the accuracy to the hypersonic similarity rule applied to wing-body combinations is demonstrated with experimental results. These results indicate that the rule may prove useful for determining the aerodynamic characteristics of slender configurations at Mach numbers higher than those for which test equipment is really available. In a second part of the study, the aerodynamic performance and static stability and control characteristics of a hypersonic glider are investigated in somewhat greater detail. Results for Mach numbers from 3 to 18 for performance and 0.6 to 12 for stability and control are obtained by standard text techniques, by application of the hypersonic stability rule, and/or by use of helium as a test medium. Lift-drag ratios of about 5 for Mach numbers up to 18 are shown to be obtainable. The glider studied is shown to have acceptable longitudinal and directional stability characteristics through the range of Mach numbers studied. Some roll instability (negative effective dihedral) is found at Mach numbers near 12.

F-18 high alpha research vehicle surface pressures: Initial in-flight results and correlation with flow visualization and wind-tunnel data

NASA Technical Reports Server (NTRS)

Fisher, David F.; Banks, Daniel W.; Richwine, David M.

1990-01-01

Pressure distributions measured on the forebody and the leading-edge extensions (LEX's) of the NASA F-18 high alpha research vehicle (HARV) were reported at 10 and 50 degree angles of attack and at Mach 0.20 to 0.60. The results were correlated with HARV flow visualization and 6-percent scale F-18 wind-tunnel-model test results. The general trend in the data from the forebody was for the maximum suction pressure peaks to first appear at an angle of attack (alpha) of approximately 19 degrees and increase in magnitude with angle of attack. The LEX pressure distribution general trend was the inward progression and increase in magnitude of the maximum suction peaks up to vortex core breakdown and then the decrease and general flattening of the pressure distribution beyond that. No significant effect of Mach number was noted for the forebody results. However, a substantial compressibility effect on the LEX's resulted in a significant reduction in vortex-induced suction pressure as Mach number increased. The forebody primary and the LEX secondary vortex separation lines, from surface flow visualization, correlated well with the end of pressure recovery, leeward and windward, respectively, of maximum suction pressure peaks. The flight to wind-tunnel correlations were generally good with some exceptions.

Heat Transfer and Flow on the First Stage Blade Tip of a Power Generation Gas Turbine. Part 1; Experimental Results

NASA Technical Reports Server (NTRS)

Bunker, Ronald S.; Bailey, Jeremy C.; Ameri, Ali A.

1999-01-01

A combined computational and experimental study has been performed to investigate the detailed distribution of convective heat transfer coefficients on the first stage blade tip surface for a geometry typical of large power generation turbines(>100MW). This paper is concerned with the design and execution of the experimental portion of the study. A stationary blade cascade experiment has been run consisting of three airfoils, the center airfoil having a variable tip gap clearance. The airfoil models the aerodynamic tip section of a high pressure turbine blade with inlet Mach number of 0.30, exit Mach number of 0.75, pressure ratio of 1.45, exit Reynolds number based on axial chord of 2.57 x 10(exp 6), and total turning of about 110 degrees. A hue detection based liquid crystal method is used to obtain the detailed heat transfer coefficient distribution on the blade tip surface for flat, smooth tip surfaces with both sharp and rounded edges. The cascade inlet turbulence intensity level took on values of either 5% or 9%. The cascade also models the casing recess in the shroud surface ahead of the blade. Experimental results are shown for the pressure distribution measurements on the airfoil near the tip gap, on the blade tip surface, and on the opposite shroud surface. Tip surface heat transfer coefficient distributions are shown for sharp-edge and rounded-edge tip geometries at each of the inlet turbulence intensity levels.

Boundary-layer and wake measurements on a swept, circulation-control wing

NASA Technical Reports Server (NTRS)

Spaid, Frank W.; Keener, Earl R.

1987-01-01

Wind-tunnel measurements of boundary-layer and wake velocity profiles and surface static pressure distributions are presented for a swept, circulation-control wing. The model is an aspect-ratio-four semispan wing mounted on the tunnel side wall at a sweep angle of 45 deg. A full-span, tangential, rearward blowing, circulation-control slot is located ahead of the trailing edge on the upper surface. Flow surveys were obtained at mid-semispan at freestream Mach numbers of 0.425 and 0.70. Boundary-layer profiles measured on the forward portions of the wing are approximately streamwise and two dimensional. The flow in the vicinity of the jet exit and in the near wake is highly three dimensional. The jet flow near the slot on the Coanda surface is directed normal to the slot. Near-wake surveys show large outboard flows at the center of the wake. At Mach 0.425 and a 5-deg angle of attack, a range of jet-blowing rates was found for which an abrupt transition from incipient separation to attached flow occurs in the boundary layer upstream of the slot. The variation in the lower-surface separation location with blowing rate was determined from boundary-layer measurements at Mach 0.425.

Study on the shock interference in a wedged convergent-divergent channel

NASA Astrophysics Data System (ADS)

Yu, F. M.; Wang, C. Z.

The investigation of shock reflection-to-diffraction phenomena upon a wedged convergent-divergent channel produced by a planar incident shock wave have been done in the shock tube facility of Institute of Aeronautics and Astronautics, National Cheng-Kung University. The experiment proceeds upon seven wedged convergent-divergent channels with the forward and rear wedge angles arrangement of them are (50°, 50°), (35°, 35°), (50°, 35°), (35°, 50°), (50°, 0°), (35°, 0°), and (90°, 0°), respectively. They were tested at Mach numbers of 1.1, 1.2, 1.3, 1.4, 1.5 and 1.6, respectively. On the first wedged channel, following the regular reflection on a 50°- wedged surface by the incident shock wave, shock diffraction with Mach stem has been observed as it moves to the downstream wedge surface. On the apex of the wedge, the secondary reflected shock behaviors as a sector of the blast shock moving toward the centerline of the channel. From the color schlieren pictures it has been observed that there exists a pattern of blast-wave-type high gas density gradient region near the wedge apex. Following the Mach reflection from the 35° -wedged surface on which only the Mach stem diffracted across the apex and following with a small region of disturbed acoustic wave front. The shock interference, which proceeds by the Mach reflection-to-diffraction generates a very complicate vortical flow structure. The measurement of the peak pressure along centerline of the channel downstream of the wedge apex indicates that it is larger near the apex and it decreases downstream. It is larger for larger convergent wedge angle and It is smaller for larger divergent wedge angle.

Projectile Combustion Effects on Ram Accelerator Performance

NASA Astrophysics Data System (ADS)

Chitale, Saarth Anjali

University of Washington Abstract Projectile Combustion Effects on Ram Accelerator Performance Saarth Anjali Chitale Chair of the Supervisory Committee: Prof. Carl Knowlen William E. Boeing Department of Aeronautics and Astronautics The ram accelerator facility at the University of Washington is used to propel projectiles at supersonic velocities. This concept is similar to an air-breathing ramjet engine in that sub-caliber projectiles, shaped like the ramjet engine center-body, are shot through smooth-bore steel-walled tubes having an internal diameter of 38 mm. The ram accelerator propulsive cycles operate between Mach 2 to 10 and have the potential to accelerate projectile to velocities greater than 8 km/s. The theoretical thrust versus Mach number characteristics can be obtained using knowledge of gas dynamics and thermodynamics that goes into the design of the ram accelerator. The corresponding velocity versus distance profiles obtained from the test runs at the University of Washington, however, are often not consistent with the theoretical predictions after the projectiles reach in-tube Mach numbers greater than 4. The experimental velocities are typically greater than the expected theoretical predictions; which has led to the proposition that the combustion process may be moving up onto the projectile. An alternative explanation for higher than predicted thrust, which is explored here, is that the performance differences can be attributed to the ablation of the projectile body which results in molten metal being added to the flow of the gaseous combustible mixture around the projectile. This molten metal is assumed to mix uniformly and react with the gaseous propellant; thereby enhancing the propellant energy release and altering the predicted thrust-Mach characteristics. This theory predicts at what Mach number the projectile will first experience enhanced thrust and the corresponding velocity-distance profile. Preliminary results are in good agreement with projectiles operating in methane/oxygen/nitrogen propellants. Effects of projectile surface to volume ratio are also explored by applying the model to experimental results from smaller (Tohoku University, 25-mm-bore) and larger (Institute of Saint-Louis 90-mm-bore) bore ram accelerators. Due to lower surface-to-volume ratio, large diameter projectiles are predicted to need to reach higher Mach numbers than smaller diameter projectiles before thrust enhancement due to metal ablation and burning would be experienced. This proposition was supported by published experimental data. The theoretical modeling of projectile ablation, metal combustion, and subsequent ram accelerator thrust characteristics are presented along comparisons to experiments from three different sized ram accelerator facilities.

Numerical simulation of the tip vortex off a low-aspect-ratio wing at transonic speed

NASA Technical Reports Server (NTRS)

Mansour, N. N.

1984-01-01

The viscous transonic flow around a low aspect ratio wing was computed by an implicit, three dimensional, thin-layer Navier-Stokes solver. The grid around the geometry of interest is obtained numerically as a solution to a Dirichlet problem for the cube. A low aspect ratio wing with large sweep, twist, taper, and camber is the chosen geometry. The topology chosen to wrap the mesh around the wing with good tip resolution is a C-O type mesh. The flow around the wing was computed for a free stream Mach number of 0.82 at an angle of attack of 5 deg. At this Mach number, an oblique shock forms on the upper surface of the wing, and a tip vortex and three dimensional flow separation off the wind surface are observed. Particle path lines indicate that the three dimensional flow separation on the wing surface is part of the roots of the tip vortex formation. The lifting of the tip vortex before the wing trailing edge is observed by following the trajectory of particles release around the wing tip.

Aerodynamic pressures and heating rates on surfaces between split elevons at Mach 6.6

NASA Technical Reports Server (NTRS)

Hunt, L. Roane

1988-01-01

An aerothermal study was performed in the Langley 8-Foot High Temperature Tunnel at Mach number 6.6 to define the pressures and heating rates on the surfaces between split elevons similar to those used on the Space Shuttle. Tests were performed with both laminar and turbulent boundary layers on the wing surface upstream of the elevons. The flow in the chordwise gap between the elevons was characterized by flow separation at the gap entrance and flow reattachment at a depth into the gap inversely proportional to the gap width. The gap pressure and heating rate increased significantly with decrease of elevon gap width, and the maximum gap heating rate was proportional to the maximum gap pressure. Correlation of the present results indicate that the gap heating was directly proportional to the elevon windward surface pressure and was not dependent upon whether the boundary layer on the windward elevon surface was laminar or turbulent.

Transonic Stability and Control Investigation of a 1/80-Scale Model of the Consolidated Vultee Skate 9 Seaplane, TED No. NACA DE 345: Transonic-Bump Method

NASA Technical Reports Server (NTRS)

Riebe, John M.; MacLeod, Richard G.

1950-01-01

An investigation of the longitudinal stability and of the all-movable horizontal tail and aileron control of a 1/80-scale reflection-plane model of the Consolidated Vultee Skate 9 seaplane has been made through a Mach number range of 0.6 to 1.16 on the transonic bump of the Langley high-speed 7- by 10-foot tunnel. At moderate lift coefficients (0.4 to 0.8) and below a Mach number of 1.0 the model was statically unstable longitudinally. The static longitudinal stability of the model at low lift coefficients increased with Mach number corresponding to a shift in aerodynamic center from 37 percent mean aerodynamic chord at a Mach number of 0.60 to 64 percent at a Mach number of 1.10. Estimates indicate that the tail deflection angle required for steady flight and for accelerated maneuvers of the Skate 9 airplane would probably not vary greatly with Mach number at sea level, but for accelerated maneuvers at altitude the tail deflection angle would probably vary erratically with Mach number. The variation of rolling-moment coefficient with aileron deflection angle was approximately linear, agreed well with theory, and held for the range of aileron deflections tested (-17.1 deg to 16.6 deg). At low lift coefficients the drag rise occurred at Mach numbers of 0.95 and 0.90 for the wing alone and the complete model, respectively. The effects of the canopy on the model were small. For the ranges investigated, angle-of-attack and Mach number changes caused no large pressure drops in the jet-engine duct.

A tabulation of pipe length to diameter ratios as a function of Mach number and pressure ratios for compressible flow

NASA Technical Reports Server (NTRS)

Dixon, G. V.; Barringer, S. R.; Gray, C. E.; Leatherman, A. D.

1975-01-01

Computer programs and resulting tabulations are presented of pipeline length-to-diameter ratios as a function of Mach number and pressure ratios for compressible flow. The tabulations are applicable to air, nitrogen, oxygen, and hydrogen for compressible isothermal flow with friction and compressible adiabatic flow with friction. Also included are equations for the determination of weight flow. The tabulations presented cover a wider range of Mach numbers for choked, adiabatic flow than available from commonly used engineering literature. Additional information presented, but which is not available from this literature, is unchoked, adiabatic flow over a wide range of Mach numbers, and choked and unchoked, isothermal flow for a wide range of Mach numbers.

Experimental investigation of tangential blowing for control of the strong shock boundary layer interaction on inlet ramps

NASA Technical Reports Server (NTRS)

Schwendemann, M. F.

1981-01-01

A 0.165-scale isolated inlet model was tested in the NASA Lewis Research Center 8-ft by 6-ft Supersonic Wind Tunnel. Ramp boundary layer control was provided by tangential blowing from a row of holes in an aft-facing step set into the ramp surface. Testing was performed at Mach numbers from 1.36 to 1.96 using both cold and heated air in the blowing system. Stable inlet flow was achieved at all Mach numbers. Blowing hole geometry was found to be significant at 1.96M. Blowing air temperature was found to have only a small effect on system performance. High blowing levels were required at the most severe test conditions.

Implementation of a Single-Stage-To-Orbit (SSTO) model for stability and control analysis

NASA Astrophysics Data System (ADS)

Ingalls, Stephen A.

1995-07-01

Three NASA centers: Marshall Space Flight Center (MSFC), Langley Research Center (LaRC), and Johnson Space Center (JSC) are currently involved in studying a family of single-stage- and two-stage-to-orbit (SSTO/TSTO) vehicles to serve as the next generation space transportation system (STS). A rocketed winged-body is the current focus. The configuration (WB001) is a vertically-launched, horizontally-landing system with circular cross-section. Preliminary aerodynamic data was generated by LaRC and is a combination of wind-tunnel data, empirical methods, and Aerodynamic Preliminary Analysis System-(APAS) generated values. JSC's efforts involve descent trajectory design, stability analysis, and flight control system synthesis. Analysis of WB001's static stability indicates instability in 'tuck' (C(sub mu) less than 0: Mach = 0.30, alpha greater than 3.25 deg; Mach = 0.60, alpha greater than 8.04), an unstable dihedral effects (C(sub l(beta)) greater than 0: Mach = 30,alpha less than 12 deg.; Mach = 0.60, alpha less than 10.00 deg.), and, most significantly, an unstable weathercock stability derivative, C(sub n(beta)), at all angles of attack and subsonic Mach numbers. Longitudinal trim solutions for Mach = 0.30 and 0.60 indicate flight path angle possibilities ranging from around 12 (M = 0.30) to slightly over 20 degrees at Mach = 0.60. Trim angles of attack increase from 6.24 at Mach 0.60 and 10,000 feet to 17.7 deg. at Mach 0.30, sea-level. Lateral trim was attempted for a design cross-wind of 25.0 knots. The current vehicle aerodynamic and geometric characteristics will only yield a lateral trim solution at impractical tip-fin deflections (approximately equal to 43 deg.) and bank angles (21 deg.). A study of the lateral control surfaces, tip-fin controllers for WB001, indicate increased surface area would help address these instabilities, particularly the deficiency in C(sub n(beta)), but obviously at the expense of increased vehicle weight. Growth factors of approximately 7 were determined using a design C(sub n(beta)) of 0.100/radian (approximate subsonic values for the orbiter).

Implementation of a Single-Stage-To-Orbit (SSTO) model for stability and control analysis

NASA Technical Reports Server (NTRS)

Ingalls, Stephen A.

1995-01-01

Three NASA centers: Marshall Space Flight Center (MSFC), Langley Research Center (LaRC), and Johnson Space Center (JSC) are currently involved in studying a family of single-stage- and two-stage-to-orbit (SSTO/TSTO) vehicles to serve as the next generation space transportation system (STS). A rocketed winged-body is the current focus. The configuration (WB001) is a vertically-launched, horizontally-landing system with circular cross-section. Preliminary aerodynamic data was generated by LaRC and is a combination of wind-tunnel data, empirical methods, and Aerodynamic Preliminary Analysis System-(APAS) generated values. JSC's efforts involve descent trajectory design, stability analysis, and flight control system synthesis. Analysis of WB001's static stability indicates instability in 'tuck' (C(sub mu) less than 0: Mach = 0.30, alpha greater than 3.25 deg; Mach = 0.60, alpha greater than 8.04), an unstable dihedral effects (C(sub l(beta)) greater than 0: Mach = 30,alpha less than 12 deg.; Mach = 0.60, alpha less than 10.00 deg.), and, most significantly, an unstable weathercock stability derivative, C(sub n(beta)), at all angles of attack and subsonic Mach numbers. Longitudinal trim solutions for Mach = 0.30 and 0.60 indicate flight path angle possibilities ranging from around 12 (M = 0.30) to slightly over 20 degrees at Mach = 0.60. Trim angles of attack increase from 6.24 at Mach 0.60 and 10,000 feet to 17.7 deg. at Mach 0.30, sea-level. Lateral trim was attempted for a design cross-wind of 25.0 knots. The current vehicle aerodynamic and geometric characteristics will only yield a lateral trim solution at impractical tip-fin deflections (approximately equal to 43 deg.) and bank angles (21 deg.). A study of the lateral control surfaces, tip-fin controllers for WB001, indicate increased surface area would help address these instabilities, particularly the deficiency in C(sub n(beta)), but obviously at the expense of increased vehicle weight. Growth factors of approximately 7 were determined using a design C(sub n(beta)) of 0.100/radian (approximate subsonic values for the orbiter).

Effect of Surface Waviness on Transition in Three-Dimensional Boundary-Layer Flow

NASA Technical Reports Server (NTRS)

Masad, Jamal A.

1996-01-01

The effect of a surface wave on transition in three-dimensional boundary-layer flow over an infinite swept wing was studied. The mean flow computed using interacting boundary-layer theory, and transition was predicted using linear stability theory coupled with the empirical eN method. It was found that decreasing the wave height, sweep angle, or freestream unit Reynolds number, and increasing the freestream Mach number or suction level all stabilized the flow and moved transition onset to downstream locations.

Columbia: The first five flights entry heating data series. Volume 2: The OMS Pod

NASA Technical Reports Server (NTRS)

Williams, S. D.

1983-01-01

Entry heating flight data and wind tunnel data on the OMS Pod are presented for the first five flights of the Space Shuttle Orbiter. The heating rate data are presented in terms of normalized film heat transfer coefficients as a function of angle-of-attack, Mach number, and normal shock Reynolds number. The surface heating rates and temperatures were obtained via the JSC NONLIN/INVERSE computer program. Time history plots of the surface heating rates and temperatures are also presented.

Modal propagation angles in a cylindrical duct with flow and their relation to sound radiation

NASA Technical Reports Server (NTRS)

Rice, E. J.; Heidmann, M. F.; Sofrin, T. G.

1979-01-01

The main emphasis is upon the propagation angle with respect to the duct axis and its relation to the far-field acoustic radiation pattern. When the steady flow Mach number is accounted for in the duct, the propagation angle in the duct is shown to be coincident with the angle of the principal lobe of far-field radiation obtained using the Wiener-Hopf technique. Different Mach numbers are allowed within the duct and in the external field. For static tests with a steady flow in an inlet but with no external Mach number the far-field radiation pattern is shifted considerably toward the inlet axis when compared to zero Mach number radiation theory. As the external Mach number is increased the noise radiation pattern is shifted away from the inlet axis. The theory is developed using approximations for sound propagation in circular ducts. An exact analysis using Hankel function solutions for the zero Mach number case is given to provide a check of the simpler approximate theory.

Pitot-probe displacement in a supersonic turbulent boundary layer

NASA Technical Reports Server (NTRS)

Allen, J. M.

1972-01-01

Eight circular pitot probes ranging in size from 2 to 70 percent of the boundary-layer thickness were tested to provide experimental probe displacement results in a two-dimensional turbulent boundary layer at a nominal free-stream Mach number of 2 and unit Reynolds number of 8 million per meter. The displacement obtained in the study was larger than that reported by previous investigators in either an incompressible turbulent boundary layer or a supersonic laminar boundary layer. The large probes indicated distorted Mach number profiles, probably due to separation. When the probes were small enough to cause no appreciable distortion, the displacement was constant over most of the boundary layer. The displacement in the near-wall region decreased to negative displacement in some cases. This near-wall region was found to extend to about one probe diameter from the test surface.

Predicted and experimental steady and unsteady transonic flows about a biconvex airfoil

NASA Technical Reports Server (NTRS)

Levy, L. L., Jr.

1981-01-01

Results of computer code time dependent solutions of the two dimensional compressible Navier-Stokes equations and the results of independent experiments are compared to verify the Mach number range for instabilities in the transonic flow field about a 14 percent thick biconvex airfoil at an angle of attack of 0 deg and a Reynolds number of 7 million. The experiments were conducted in a transonic, slotted wall wind tunnel. The computer code included an algebraic eddy viscosity turbulence model developed for steady flows, and all computations were made using free flight boundary conditions. All of the features documented experimentally for both steady and unsteady flows were predicted qualitatively; even with the above simplifications, the predictions were, on the whole, in good quantitative agreement with experiment. In particular, predicted time histories of shock wave position, surface pressures, lift, and pitching moment were found to be in very good agreement with experiment for an unsteady flow. Depending upon the free stream Mach number for steady flows, the surface pressure downstream of the shock wave or the shock wave location was not well predicted.

Supersonic Aerodynamic Characteristics of Proposed Mars '07 Smart Lander Configurations

NASA Technical Reports Server (NTRS)

Murphy, Kelly J.; Horvath, Thomas J.; Erickson, Gary E.; Green, Joseph M.

2002-01-01

Supersonic aerodynamic data were obtained for proposed Mars '07 Smart Lander configurations in NASA Langley Research Center's Unitary Plan Wind Tunnel. The primary objective of this test program was to assess the supersonic aerodynamic characteristics of the baseline Smart Lander configuration with and without fixed shelf/tab control surfaces. Data were obtained over a Mach number range of 2.3 to 4.5, at a free stream Reynolds Number of 1 x 10(exp 6) based on body diameter. All configurations were run at angles of attack from -5 to 20 degrees and angles of sideslip of -5 to 5 degrees. These results were complemented with computational fluid dynamic (CFD) predictions to enhance the understanding of experimentally observed aerodynamic trends. Inviscid and viscous full model CFD solutions compared well with experimental results for the baseline and 3 shelf/tab configurations. Over the range tested, Mach number effects were shown to be small on vehicle aerodynamic characteristics. Based on the results from 3 different shelf/tab configurations, a fixed control surface appears to be a feasible concept for meeting aerodynamic performance metrics necessary to satisfy mission requirements.

Correlation of transonic-cone Preston-tube data and skin friction. [characterizing the flow quality of a transonic wind tunnel

NASA Technical Reports Server (NTRS)

Reed, T. D.

1981-01-01

The distribution of Preston tube pressures within turbulent boundary layers along the surface of a sharp-nosed, ten degree cone was correlated with theoretical values of turbulent skin friction for freestream Mach numbers less than one. The mini-basic computer code, the Wu and Lock computer code, and the STAN-5 computer code were used to analyze the data and to solve the boundary layer conservation equations. The skin friction which results from using Preston tube pressures in the correlation equation, has a rms error of 1.125 percent. It was found that the effective center of the probe is not a constant but increases as the surface distance increases. For a specified unit Reynolds number, the effective center of the probe decreases as the Mach number increases. The variation of the fluid (air) properties across the face of the probe may be neglected for subsonic flows. The possible transverse errors caused by the use of the concept of a virtual origin for the turbulent boundary layer were investigated and found to be negligible.

An Investigation of the Longitudinal Characteristics of the MX-656 Configuration Using Rocket-Propelled Models Preliminary Results at Mach Numbers from 0.65 to 1.25

NASA Technical Reports Server (NTRS)

Mitchell, Jesse L.; Peck, Robert F.

1950-01-01

A rocket-propelled model of the Mx-656 configuration has been flown through the Mach number range from 0.65 to 1.25. An analysis of the response of the model to rapid deflections of the horizontal tail gave information on the lift, drag, longitudinal stability and control, and longitudinal-trim change. The lift-coefficient range covered by the test was from -0.2 to 0,3 throughout most of the Mach number range, The model was statically and dynamically stable throughout the lift-coefficient and Mach number range of the test. At subsonic speeds the aerodynamic center moved f o m r d with increasing lift coefficient. The most forward position of the aerodynamic center was about 12,5 percent of the mean aerodynamic chord at a small positive lift coefficient and at a Mach number of about 0.84. A t supersonic speeds the aerodynamic center was well aft, varying from 33 to 39 percent of the mean aerodynamic chord at Mach numbers of 1.0 and 1.25, respectively. Transonic-trim change, as measured by the change in trim lift coefficient with Mach number at a constant t a i l setting, was of small magnitude (about 0.1 lift coefficient for zero tail setting). The zero-lift/drag coefficient increased about 0.042 in the region between a Mach number of 0.9 and 1.1

Studies of Shock Wave Interactions with Homogeneous and Isotropic Turbulence

NASA Technical Reports Server (NTRS)

Briassulis, G.; Agui, J.; Watkins, C. B.; Andreopoulos, Y.

1998-01-01

A nearly homogeneous nearly isotropic compressible turbulent flow interacting with a normal shock wave has been studied experimentally in a large shock tube facility. Spatial resolution of the order of 8 Kolmogorov viscous length scales was achieved in the measurements of turbulence. A variety of turbulence generating grids provide a wide range of turbulence scales. Integral length scales were found to substantially decrease through the interaction with the shock wave in all investigated cases with flow Mach numbers ranging from 0.3 to 0.7 and shock Mach numbers from 1.2 to 1.6. The outcome of the interaction depends strongly on the state of compressibility of the incoming turbulence. The length scales in the lateral direction are amplified at small Mach numbers and attenuated at large Mach numbers. Even at large Mach numbers amplification of lateral length scales has been observed in the case of fine grids. In addition to the interaction with the shock the present work has documented substantial compressibility effects in the incoming homogeneous and isotropic turbulent flow. The decay of Mach number fluctuations was found to follow a power law similar to that describing the decay of incompressible isotropic turbulence. It was found that the decay coefficient and the decay exponent decrease with increasing Mach number while the virtual origin increases with increasing Mach number. A mechanism possibly responsible for these effects appears to be the inherently low growth rate of compressible shear layers emanating from the cylindrical rods of the grid.

Design and Performance Calculations of a Propeller for Very High Altitude Flight. Degree awarded by Case Western Univ.

NASA Technical Reports Server (NTRS)

Koch, L. Danielle

1998-01-01

Reported here is a design study of a propeller for a vehicle capable of subsonic flight in Earth's stratosphere. All propellers presented were required to absorb 63.4 kW (85 hp) at 25.9 km (85,000 ft) while aircraft cruise velocity was maintained at Mach 0.40. To produce the final design, classic momentum and blade-element theories were combined with two and three-dimensional results from the Advanced Ducted Propfan Analysis Code (ADPAC), a numerical Navier-Stokes analysis code. The Eppler 387 airfoil was used for each of the constant section propeller designs compared. Experimental data from the Langley Low-Turbulence Pressure Tunnel was used in the strip theory design and analysis programs written. The experimental data was also used to validate ADPAC at a Reynolds numbers of 60,000 and a Mach number of 0.20. Experimental and calculated surface pressure coefficients are compared for a range of angles of attack. Since low Reynolds number transonic experimental data was unavailable, ADPAC was used to generate two-dimensional section performance predictions for Reynolds numbers of 60,000 and 100,000 and Mach numbers ranging from 0.45 to 0.75. Surface pressure coefficients are presented for selected angles of attack. in addition to the variation of lift and drag coefficients at each flow condition. A three-dimensional model of the final design was made which ADPAC used to calculated propeller performance. ADPAC performance predictions were compared with strip-theory calculations at design point. Propeller efficiency predicted by ADPAC was within 1.5% of that calculated by strip theory methods, although ADPAC predictions of thrust, power, and torque coefficients were approximately 5% lower than the strip theory results. Simplifying assumptions made in the strip theory account for the differences seen.

Effects of rocket jet on stability and control at high Mach numbers

NASA Technical Reports Server (NTRS)

Fetterman, David E , Jr

1958-01-01

Paper presents the results of an investigation to determine the jet-interference effects which may occur at high jet static-pressure ratios and high Mach numbers. Tests were made in the Langley 11-inch hypersonic tunnel at a Mach number of 6.86.

Wave phenomena in a high Reynolds number compressible boundary layer

NASA Technical Reports Server (NTRS)

Bayliss, A.; Maestrello, L.; Parikh, P.; Turkel, E.

1985-01-01

Growth of unstable disturbances in a high Reynolds number compressible boundary layer is numerically simulated. Localized periodic surface heating and cooling as a means of active control of these disturbances is studied. It is shown that compressibility in itself stabilizes the flow but at a lower Mach number, significant nonlinear distortions are produced. Phase cancellation is shown to be an effective mechanism for active boundary layer control.

Effects of body shape on the aerodynamics of a body of revolution at Mach numbers from 1.6 to 4.6

NASA Technical Reports Server (NTRS)

Spearman, M. L.

1985-01-01

The aerodnamic characteristics for several bodies of revolution have been determined from wind tunnel tests at Mach numbers from 1.6 to 4.63. Six bodies, each having a length-to-diameter ratio of 6.67, were investigated. Geometric modifications included forebody shape, afterbody shape, and midsection slope. Significant aerodynamic changes were observed to be functions of geometric change and Mach number. Because of the aerodynamic dependence on geometry as well as Mach number, it is obvious that a number of trades must be considered in selecting a projectile shape.

Heat transfer measurements on an incidence-tolerant low pressure turbine blade in a high speed linear cascade at low to moderate Reynolds numbers

NASA Astrophysics Data System (ADS)

Moualeu, Leolein Patrick Gouemeni

Runway-independent aircraft are expected to be the future for short-haul flights by improving air transportation and reducing area congestion encountered in airports. The Vehicle Systems Program of NASA identified a Large Civil Tilt-Rotor, equipped with variable-speed power-turbine engines, as the best concept. At cruise altitude, the engine rotor-speed will be reduced by as much as the 50% of take-off speed. The large incidence variation in the low pressure turbine associated with the change in speed can be detrimental to the engine performance. Low pressure turbine blades in cruise altitude are more predisposed to develop regions of boundary layer separation. Typical phenomenon such as impinging wakes on downstream blades and mainstream turbulences enhance the complexity of the flow in low pressure turbines. It is therefore important to be able to understand the flow behavior to accurately predict the losses. Research facilities are seldom able to experimentally reproduce low Reynolds numbers at relevant engine Mach number. Having large incidence swing as an additional parameter in the investigation of the boundary layer development, on a low pressure turbine blade, makes this topic unique and as a consequence requires a unique facility to conduct the experimental research. The compressible flow wind tunnel facility at the University of North Dakota had been updated to perform steady state experiments on a modular-cascade, designed to replicate a large variation of the incidence angles. The high speed and low Reynolds number facility maintained a sealed and closed loop configuration for each incidence angle. The updated facility is capable to produce experimental Reynolds numbers as low as 45,000 and as high as 570,000 at an exit Mach number of 0.72. Pressure and surface temperature measurements were performed at these low pressure turbine conditions. The present thesis investigates the boundary layer development on the surface of an Incidence-tolerant blade. The heat transfer approach is the method used to obtain knowledge of the state of the boundary layer on the surface of the blade. Pressure and temperature distributions are acquired for Reynolds numbers of 50,000, 66,000, 228,000, and 568,000 at an exit Mach number of 0.72, and Reynolds numbers of 228,000, and 568,000 at an exit Mach number of 0.35. These experimental flow conditions are conducted at different flow inlet angles of 40°, 34.2°, 28°, 18°, 8°, -2.6°, -12°, and -17°, and at two free-stream turbulence levels. Results of the analyses performed show that as the incidence angle decreases, a region of laminar separation bubble forms on the pressure surface and grows toward the trailing-edge. It is also noted that the position of the leading-edge moves as the incidence angle varies. A transitional flow is observed on both the pressure and suction surfaces, mainly at the two highest incidence angles, for the high turbulence case. This investigation also reveals that the Stanton number increases as the mainstream turbulence increases, and that the Stanton number at the leading-edge increases as the Reynolds number decreases, as it is documented in the literature.

In-flight acoustic results from an advanced-design propeller at Mach numbers to 0.8

NASA Technical Reports Server (NTRS)

Mackall, K. G.; Lasagna, P. L.; Walsh, K.; Dittmar, J. H.

1982-01-01

Acoustic data for the advanced-design SR-3 propeller at Mach numbers to 0.8 and helical tip Mach numbers to 1.14 are presented. Several advanced-design propellers, previously tested in wind tunnels at the Lewis Research Center, are being tested in flight at the Dryden Flight Research Facility. The flight-test propellers are mounted on a pylon on the top of the fuselage of a JetStar airplane. Instrumentation provides near-field acoustic data for the SR-3. Acoustic data for the SR-3 propeller at Mach numbers up to 0.8, for propeller helical tip Mach numbers up to 1.14, and comparison of wind tunnel to flight data are included. Flowfield profiles measured in the area adjacent to the propeller are also included.

Aerodynamic characteristics at Mach numbers from 0.6 to 2.16 of a supersonic cruise fighter configuration with a design Mach number of 1.8

NASA Technical Reports Server (NTRS)

Shrout, B. L.

1977-01-01

An investigation was made in the Langley 8-foot transonic tunnel and the Langley Unitary Plan wind tunnel, over a Mach number range of 0.6 to 2.16, to determine the static longitudinal and lateral aerodynamic characteristics of a model of a supersonic-cruise fighter. The configuration, which is designed for efficient cruise at Mach number 1.8, is a twin-engine tailless arrow-wing concept with a single rectangular inlet beneath the fuselage and outboard vertical tails and ventral fins. It had untrimmed values of lift-drage ratio ranging from 10 at subsonic speeds to 6.4 at the design Mach number. The configuration was statically stable both longitudinally and laterally.

The NASA Langley 8-foot Transonic Pressure Tunnel calibration

NASA Technical Reports Server (NTRS)

Brooks, Cuyler W., Jr.; Harris, Charles D.; Reagon, Patricia G.

1994-01-01

The NASA Langley 8-Foot Transonic Pressure Tunnel is a continuous-flow, variable-pressure wind tunnel with control capability to independently vary Mach number, stagnation pressure, stagnation temperature, and humidity. The top and bottom walls of the test section are axially slotted to permit continuous variation of the test section Mach number from 0.2 to 1.2, the slot-width contour provides a gradient-free test section 50 in. long for Mach numbers equal to or greater than 1.0 and 100 in. long for Mach numbers less than 1.0. The stagnation pressure may be varied from 0.25 to 2.0 atm. The tunnel test section has been recalibrated to determine the relationship between the free-stream Mach number and the test chamber reference Mach number. The hardware was the same as that of an earlier calibration in 1972 but the pressure measurement instrumentation available for the recalibration was about an order of magnitude more precise. The principal result of the recalibration was a slightly different schedule of reentry flap settings for Mach numbers from 0.80 to 1.05 than that determined during the 1972 calibration. Detailed tunnel contraction geometry, test section geometry, and limited test section wall boundary layer data are presented.

Longitudinal Stability and Drag Characteristics at Mach Numbers from 0.70 to 1.37 of Rocket-propelled Models Having a Modified Triangular Wing

NASA Technical Reports Server (NTRS)

Chapman, Rowe, Jr; Morrow, John D

1952-01-01

A modified triangular wing of aspect ratio 2.53 having an airfoil section 3.7 percent thick at the root and 5.98 percent thick at the tip was designed in an attempt to improve the lift and drag characteristics of triangular wings. Free-flight drag and stability tests were made using rocket-propelled models equipped with the modified wing. The Mach number range of the test was from 0.70 to 1.37. Test results indicated the following: The lift-curve slope of wing plus fuselage approaches the theoretical value of wing alone at supersonic Mach numbers. The drag coefficient, based on total wing area, for wing plus interference was approximately 0.0035 at subsonic Mach numbers and 0.0080 at supersonic Mach numbers. The maximum shift in aerodynamic center for the complete configuration was 14 percent in the rearward direction from the forward position of 51.5 percent of mean aerodynamic chord at subsonic Mach numbers. The variation of lift and moment with angle of attack was linear at supersonic Mach numbers for the range of coefficients covered in the test. The high value of lift-curve slope was considered to be a significant result attributable to the wing modifications.

Results of tests OA63 and IA29 on an 0.015 scale model of the space shuttle configuration 140 A/B in the NASA/ARC 6- by 6-foot transonic wind tunnel, volume 1

NASA Technical Reports Server (NTRS)

Spangler, R. H.; Thornton, D. E.

1974-01-01

Tests were conducted in the NASA/ARC 6- by 6-foot transonic wind tunnel from September 12 to September 28, 1973 on an 0.015-scale model of the space shuttle configuration 140 A/B. Surface pressure data were obtained for the orbiter for both launch and entry configuration at Mach numbers from 0.6 to 2.0. The surface pressures were obtained in the vicinity of the cargo bay door hinge and parting lines, the side of the fuselage at the crew compartment and below the OMS pods at the aft compartment. Data were obtained at angles of attack and sideslip consistent with the expected divergencies along the nominal trajectory. These tests were first in a series of tests supporting the orbiter venting analysis. The series will include tests in three facilities covering a total Mach number range from 0.6 to 10.4.

Hypersonic vehicle simulation model: Winged-cone configuration

NASA Technical Reports Server (NTRS)

Shaughnessy, John D.; Pinckney, S. Zane; Mcminn, John D.; Cruz, Christopher I.; Kelley, Marie-Louise

1990-01-01

Aerodynamic, propulsion, and mass models for a generic, horizontal-takeoff, single-stage-to-orbit (SSTO) configuration are presented which are suitable for use in point mass as well as batch and real-time six degree-of-freedom simulations. The simulations can be used to investigate ascent performance issues and to allow research, refinement, and evaluation of integrated guidance/flight/propulsion/thermal control systems, design concepts, and methodologies for SSTO missions. Aerodynamic force and moment coefficients are given as functions of angle of attack, Mach number, and control surface deflections. The model data were estimated by using a subsonic/supersonic panel code and a hypersonic local surface inclination code. Thrust coefficient and engine specific impulse were estimated using a two-dimensional forebody, inlet, nozzle code and a one-dimensional combustor code and are given as functions of Mach number, dynamic pressure, and fuel equivalence ratio. Rigid-body mass moments of inertia and center of gravity location are functions of vehicle weight which is in turn a function of fuel flow.

Aerodynamic Characteristics of Two Waverider-Derived Hypersonic Cruise Configurations

NASA Technical Reports Server (NTRS)

Cockrell, Charles E., Jr.; Huebner, Lawrence D.; Finley, Dennis B.

1996-01-01

An evaluation was made on the effects of integrating the required aircraft components with hypersonic high-lift configurations known as waveriders to create hypersonic cruise vehicles. Previous studies suggest that waveriders offer advantages in aerodynamic performance and propulsion/airframe integration (PAI) characteristics over conventional non-waverider hypersonic shapes. A wind-tunnel model was developed that integrates vehicle components, including canopies, engine components, and control surfaces, with two pure waverider shapes, both conical-flow-derived waveriders for a design Mach number of 4.0. Experimental data and limited computational fluid dynamics (CFD) solutions were obtained over a Mach number range of 1.6 to 4.63. The experimental data show the component build-up effects and the aerodynamic characteristics of the fully integrated configurations, including control surface effectiveness. The aerodynamic performance of the fully integrated configurations is not comparable to that of the pure waverider shapes, but is comparable to previously tested hypersonic models. Both configurations exhibit good lateral-directional stability characteristics.

Flush Airdata Sensing (FADS) System Calibration Procedures and Results for Blunt Forebodies

NASA Technical Reports Server (NTRS)

Cobleigh, Brent R.; Whitmore, Stephen A.; Haering, Edward A., Jr.; Borrer, Jerry; Roback, V. Eric

1999-01-01

Blunt-forebody pressure data are used to study the behavior of the NASA Dryden Flight Research Center flush airdata sensing (FADS) pressure model and solution algorithm. The model relates surface pressure measurements to the airdata state. Spliced from the potential flow solution for uniform flow over a sphere and the modified Newtonian impact theory, the model was shown to apply to a wide range of blunt-forebody shapes and Mach numbers. Calibrations of a sphere, spherical cones, a Rankine half body, and the F-14, F/A-18, X-33, X-34, and X-38 configurations are shown. The three calibration parameters are well-behaved from Mach 0.25 to Mach 5.0, an angle-of-attack range extending to greater than 30 deg, and an angle-of-sideslip range extending to greater than 15 deg. Contrary to the sharp calibration changes found on traditional pitot-static systems at transonic speeds, the FADS calibrations are smooth, monotonic functions of Mach number and effective angles of attack and sideslip. Because the FADS calibration is sensitive to pressure port location, detailed measurements of the actual pressure port locations on the flight vehicle are required and the wind-tunnel calibration model should have pressure ports in similar locations. The procedure for calibrating a FADS system is outlined.

Flow-field measurements in the windward surface shock layer of space shuttle orbiter configurations at Mach number 8. [wind tunnel tests of scale models

NASA Technical Reports Server (NTRS)

Martindale, W. R.; Carter, L. D.

1975-01-01

Pitot pressure and total-temperature measurements were made in the windward surface shock layer of two 0.0175-scale space shuttle orbiter models at simulated re-entry conditions. Corresponding surface static pressure measurements were also made. Flow properties at the edge of the model boundary layer were derived from these measurements and compared with values calculated using conventional methods.

Flight Test Measurements From The Tu-144LL Structure/Cabin Noise Follow-On Experiment

NASA Technical Reports Server (NTRS)

Rizzi, Stephen A.; Rackl, Robert G.; Andrianov, Eduard V.

2000-01-01

This follow-on flight experiment on the TU-144LL Supersonic Flying Laboratory, conducted during the period September 1998 to April 1999, was a continuation of previous Structure/Cabin Noise Experiment 2.1. Data was obtained over a wide range of altitudes and Mach numbers. Measured were: turbulent boundary layer pressure fluctuations on the fuselage over its length; structural response on skin panels using accelerometers; and flow direction over three windows using 'flow cones'. The effect of steps in the flow was also measured using two window blank pairs; each pair bridged by a plate which created small sharp forward and aft facing steps. The effect of transducer flushness with the exterior surface was also measured during flight. Height test points were chosen to cover much of the TU-144's flight envelope, as well as to obtain as large a unit Reynolds number range as possible at various Mach numbers: takeoff, subsonic, transonic, and supersonic cruise conditions up to Mach 2. Data on engine runups and background noise were acquired on the ground. The data in the form of time histories of the acoustic signals, together with auxiliary data and basic MATLAB processing modules, are available on CD-R disks.

Measurement of the steady surface pressure distribution on a single rotation large scale advanced prop-fan blade at Mach numbers from 0.03 to 0.78

NASA Technical Reports Server (NTRS)

Bushnell, Peter

1988-01-01

The aerodynamic pressure distribution was determined on a rotating Prop-Fan blade at the S1-MA wind tunnel facility operated by the Office National D'Etudes et de Recherches Aerospatiale (ONERA) in Modane, France. The pressure distributions were measured at thirteen radial stations on a single rotation Large Scale Advanced Prop-Fan (LAP/SR7) blade, for a sequence of operating conditions including inflow Mach numbers ranging from 0.03 to 0.78. Pressure distributions for more than one power coefficient and/or advanced ratio setting were measured for most of the inflow Mach numbers investigated. Due to facility power limitations the Prop-Fan test installation was a two bladed version of the eight design configuration. The power coefficient range investigated was therefore selected to cover typical power loading per blade conditions which occur within the Prop-Fan operating envelope. The experimental results provide an extensive source of information on the aerodynamic behavior of the swept Prop-Fan blade, including details which were elusive to current computational models and do not appear in the two-dimensional airfoil data.

Results of a study of Mach number and Reynolds number effects on the crossflow drag characteristics of ogive cylinders and ogive-cylinder-frustum-cylinders at angles of attack to 30 degrees

NASA Technical Reports Server (NTRS)

Foley, J. E.

1971-01-01

An analysis was made to determine the effects of Mach number and Reynolds number on the local and total crossflow drag characteristics of ogive-cylinders and ogive-cylinder-frustum-cylinders at angles of the MSFC 14 in TWT and the LTV 4 ft HSWT, and pressure data obtained in the TWT, at Mach numbers 0.14, 0.8, 1.2, and 2.0, and a wide range of Reynolds numbers. Results indicate that the streamwise Reynolds number, VD/nusin alpha, is an important correlation parameter in the subcritical Reynolds number range at imcompressible speeds and that the crossflow Mach number correlates compressibility effects.

On the transonic aerodynamics of a compressor blade row

NASA Technical Reports Server (NTRS)

Erickson, J. C., Jr.; Lordi, J. A.; Rae, W. J.

1971-01-01

Linearized analyses have been carried out for the induced velocity and pressure fields within a compressor blade row operating in an infinite annulus at transonic Mach numbers of the flow relative to the blades. In addition, the relationship between the induced velocity and the shape of the mean blade surface has been determined. A computational scheme has been developed for evaluating the blade mean surface ordinates and surface pressure distributions. The separation of the effects of a specified blade thickness distribution from the effects of a specified distribution of the blade lift has been established. In this way, blade mean surface shapes that are necessary for the blades to be locally nonlifting have been computed and are presented for two examples of blades with biconvex parabolic arc sections of radially tapering thickness. Blade shapes that are required to achieve a zero thickness, uniform chordwise loading, constant work spanwise loading are also presented for two examples. In addition, corresponding surface pressure distributions are given. The flow relative to the blade tips has a high subsonic Mach number in the examples that have been computed. The results suggest that at near-sonic relative tip speeds the effective blade shape is dominated by the thickness distribution, with the lift distribution playing only a minor role.

Experimental Aerodynamic Characteristics of the Pegasus Air-Launched Booster and Comparisons with Predicted and Flight Results

NASA Technical Reports Server (NTRS)

Rhode, M. N.; Engelund, Walter C.; Mendenhall, Michael R.

1995-01-01

Experimental longitudinal and lateral-directional aerodynamic characteristics were obtained for the Pegasus and Pegasus XL configurations over a Mach number range from 1.6 to 6 and angles of attack from -4 to +24 degrees. Angle of sideslip was varied from -6 to +6 degrees, and control surfaces were deflected to obtain elevon, aileron, and rudder effectiveness. Experimental data for the Pegasus configuration are compared with engineering code predictions performed by Nielsen Engineering & Research, Inc. (NEAR) in the aerodynamic design of the Pegasus vehicle, and with results from the Aerodynamic Preliminary Analysis System (APAS) code. Comparisons of experimental results are also made with longitudinal flight data from Flight #2 of the Pegasus vehicle. Results show that the longitudinal aerodynamic characteristics of the Pegasus and Pegasus XL configurations are similar, having the same lift-curve slope and drag levels across the Mach number range. Both configurations are longitudinally stable, with stability decreasing towards neutral levels as Mach number increases. Directional stability is negative at moderate to high angles of attack due to separated flow over the vertical tail. Dihedral effect is positive for both configurations, but is reduced 30-50 percent for the Pegasus XL configuration because of the horizontal tail anhedral. Predicted longitudinal characteristics and both longitudinal and lateral-directional control effectiveness are generally in good agreement with experiment. Due to the complex leeside flowfield, lateral-directional characteristics are not as well predicted by the engineering codes. Experiment and flight data are in good agreement across the Mach number range.

Transonic and Supersonic Wind-Tunnel Tests of Wing-Body Combinations Designed for High Efficiency at a Mach Number of 1.41

NASA Technical Reports Server (NTRS)

Grant, Frederick C.; Sevier, John R., Jr.

1960-01-01

Wind-tunnel force tests of a number of wing-body combinations designed for high lift-drag ratio at a Mach number of 1.41 are reported. Five wings and six bodies were used in making up the various wing-body combinations investigated. All the wings had the same highly swept dis- continuously tapered plan form with NACA 65A-series airfoil sections 4 percent thick at the root tapering linearly to 3 percent thick at the tip. The bodies were based on the area distribution of a Sears-Haack body of revolution for minimum drag with a given length and volume. These wings and bodies were used to determine the effects of wing twist., wing twist and camber, wing leading-edge droop, a change from circular to elliptical body cross-sectional shape, and body indentation by the area-rule and streamline methods. The supersonic test Mach numbers were 1.41 and 2.01. The transonic test Mach number range was from 0.6 to 1.2. For the transition-fixed condition and at a Reynolds number of 2.7 x 10(exp 6) based on the mean aerodynamic chord, the maximum value of lift- drag ratio at a Mach number of 1.41 was 9.6 for a combination with a twisted wing and an indented body of elliptical cross section. The tests indicated that the transonic rise in minimum drag was low and did not change appreciably up to the highest test Mach number of 2.01. The lower values of lift-drag ratio obtained at a Mach number of 2.01 can be attributed to the increase of drag due to lift with Mach number.

Measurements in Flight of the Pressure Distribution on the Right Wing of a Pursuit-Type Airplane at Several Values of Mach Number

NASA Technical Reports Server (NTRS)

Clousing, Lawrence A; Turner, William N; Rolls, L Stewart

1946-01-01

Pressure-distribution measurements were made on the right wing of a pursuit-type airplane at values of Mach number up to 0.80. The results showed that a considerable portion of the lift was carried by components of the airplane other than the wings, and that the proportion of lift carried by the wings may vary considerably with Mach number, thus changing the bending moment at the wing root whether or not there is a shift in the lateral position of the center of pressure. It was also shown that the center of pressure does not necessarily move outward at high Mach numbers, even though the wing-thickness ratio decreases toward the wing tip. The wing pitching-moment coefficient increased sharply in a negative direction at a Mach lift-curve slope increased with Mach number up to values of above the critical value. Pressures inside the wing were small and negative.

Analysis of sonic boom measurements near shock wave extremities for flight near Mach 1.0 and for airplane accelerations

NASA Technical Reports Server (NTRS)

Haglund, G. T.; Kane, E. J.

1974-01-01

The analysis of the 14 low-altitude transonic flights showed that the prevailing meteorological consideration of the acoustic disturbances below the cutoff altitude during threshold Mach number flight has shown that a theoretical safe altitude appears to be valid over a wide range of meteorological conditions and provides a reasonable estimate of the airplane ground speed reduction to avoid sonic boom noise during threshold Mach number flight. Recent theoretical results for the acoustic pressure waves below the threshold Mach number caustic showed excellent agreement with observations near the caustic, but the predicted overpressure levels were significantly lower than those observed far from the caustic. The analysis of caustics produced by inadvertent low-magnitude accelerations during flight at Mach numbers slightly greater than the threshold Mach number showed that folds and associated caustics were produced by slight changes in the airplane ground speed. These caustic intensities ranged from 1 to 3 time the nominal steady, level flight intensity.

Longitudinal Stability Characteristics of the Consolidated Vultee XFY-1 Airplane with Windmilling Propellers as Obtained from Flight of 0.133-Scale Rocket-Propelled Model at Mach Numbers from 0.70 to 1.13

NASA Technical Reports Server (NTRS)

Hastings, Earl C.; Mitcham, Grady L.

1954-01-01

A flight test has been conducted to determine the longitudinal stability and control characteristics of a 0.133-scale model of the Consolidated Vultee XFY-1 airplane with windmilling propellers for the Mach number range between 0.70 and 1.13. The variation of lift-curve slope C(sub L(sub alpha) with Mach number was gradual with a maximum value of 0.074 occurring at a Mach number of 0.97. Propellers had little effect upon the values of lift-curve slope or the linearity of lift coefficient with angle of attack. At lift coefficients between approximately 0.25 and 0.45 with an elevon angle of approximately -l0 deg, there was a region of neutral longitudinal stability at Mach numbers below 0.93 introduced by the addition of windmilling propellers. Below a lift coefficient of 0.10 and above a lift coefficient of 0.45, the model was longitudinally stable throughout the Mach number range of the test. There was a forward shift in the aerodynamic center of about 3-percent mean aerodynamic chord introduced by the addition of propellers. The aerodynamic center as determined at low lift moved gradually from a value of 28.5-percent mean aerodynamic chord at a Mach number of 0.75 to a value of 47-percent mean aerodynamic chord at a Mach number of 1.10. There was an abrupt decrease in pitch damping between Mach numbers of 0.88 and 0.99 followed by a rapid increase in damping to a Mach number of 1.06. The propellers had little effect upon the pitch damping characteristics . The transonic trim change was a large pitching-down tendency with and without windmilling propellers. The elevons were effective pitch controls throughout the speed range; however, their effectiveness was reduced about 50 percent at supersonic speeds. The propellers had no appreciable effect upon the control effectiveness.

Ground Measurements of Airplane Shock-Wave Noise at Mach Numbers to 2.0 and at Altitudes to 60,000 Feet

NASA Technical Reports Server (NTRS)

Lina, Lindsay J.; Maglieri, Domenic J.

1960-01-01

The intensity of shock-wave noise at the ground resulting from flights at Mach numbers to 2.0 and altitudes to 60,000 feet was measured. Meagurements near the ground track for flights of a supersonic fighter and one flight of a supersonic bomber are presented. Level cruising flight at an altitude of 60,000 feet and a Mach number of 2.0 produced sonic booms which were considered to be tolerable, and it is reasonable t o expect that cruising flight at higher altitudes will produce booms of tolerable intensity for airplanes of the size and weight of the test airplanes. The measured variation of sonic-boom intensity with altitude was in good agreement with the variation calculated by an equation given in NASA Technical Note D-48. The effect of Mach number on the ground overpressure is small between Mach numbers of 1.4 and 2.0, a result in agreement with the theory. No amplification of the shock-wave overpressures due to refraction effects was apparent near the cutoff Mach number. A method for estimating the effect of fligh-path angle on cutoff Mach number is shown. Experimental results indicate agreement with the method, since a climb maneuver produced booms of a much decreased intensity as compared with the intensity of those measured in level flight at about the same altitude and Mach number. Comparison of sound pressure levels for the fighter and bomber airp lanes indicated little effect of either airplane size or weight at an altitude of 40,000 feet.

Shuttle model tailcone pressure distribution at low subsonic speeds of a 0.03614-scale model in the NASA/LaRC low-turbulence pressure tunnel (LA81), volume 1

NASA Technical Reports Server (NTRS)

Ball, J. W.; Lindahl, R. H.

1976-01-01

An investigation was conducted in the NASA/LaRC Low-Turbulence Pressure Tunnel on a 0.03614-scale orbiter model of a 089B configuration with a 139B configuration nose forward of F.S. 500. The tailcone was the TC sub 4 design and was instrumented with eighty-nine pressure orifices. Control surfaces were deflected and three wind tunnel mounting techniques were investigated over an angle-of-attack range from -2 deg to a maximum of 18 deg. In order to determine the sensitivity of the tailcone to changes in Reynolds number, most of the test was made at a Mach number of 0.20 over a Reynolds number range of 2.0 to 10 million per foot. A few runs were made at a Mach number of 0.30 at Reynolds numbers of 4.0, 6.0, and 8 million per foot.

Aerodynamic characteristics of two 10-percent-thick NASA supercritical airfoils with different upper surface curvature distributions. [Langley 8 foot transonic tunnel tests

NASA Technical Reports Server (NTRS)

Harris, C. D.

1974-01-01

In order to assess the degree to which the characteristic region of low curvature of the supercritical airfoil can be practically extended on the upper surface, the aerodynamic characteristics of two supercritical airfoils with different upper surface curvature distributions were measured at Mach numbers from 0.60 to 0.81. Integrated section force and moment data, surface pressure distributions, and typical wake survey profiles are presented.

A Low Mach Number Model for Moist Atmospheric Flows

DOE PAGES

Duarte, Max; Almgren, Ann S.; Bell, John B.

2015-04-01

A low Mach number model for moist atmospheric flows is introduced that accurately incorporates reversible moist processes in flows whose features of interest occur on advective rather than acoustic time scales. Total water is used as a prognostic variable, so that water vapor and liquid water are diagnostically recovered as needed from an exact Clausius–Clapeyron formula for moist thermodynamics. Low Mach number models can be computationally more efficient than a fully compressible model, but the low Mach number formulation introduces additional mathematical and computational complexity because of the divergence constraint imposed on the velocity field. Here in this paper, latentmore » heat release is accounted for in the source term of the constraint by estimating the rate of phase change based on the time variation of saturated water vapor subject to the thermodynamic equilibrium constraint. Finally, the authors numerically assess the validity of the low Mach number approximation for moist atmospheric flows by contrasting the low Mach number solution to reference solutions computed with a fully compressible formulation for a variety of test problems.« less

Flight measured and calculated exhaust jet conditions for an F100 engine in an F-15 airplane

NASA Technical Reports Server (NTRS)

Hernandez, Francisco J.; Burcham, Frank W., Jr.

1988-01-01

The exhaust jet conditions, in terms of temperature and Mach number, were determined for a nozzle-aft end acoustic study flown on an F-15 aircraft. Jet properties for the F100 EMD engines were calculated using the engine manufacturer's specification deck. The effects of atmospheric temperature on jet Mach number, M10, were calculated. Values of turbine discharge pressure, PT6M, jet Mach number, and jet temperature were calculated as a function of aircraft Mach number, altitude, and power lever angle for the test day conditions. At a typical test point with a Mach number of 0.9, intermediate power setting, and an altitude of 20,000 ft, M10 was equal to 1.63. Flight measured and calculated values of PT6M were compared for intermediate power at altitudes of 15500, 20500, and 31000 ft. It was found that at 31000 ft, there was excellent agreement between both, but for lower altitudes the specification deck overpredicted the flight data. The calculated jet Mach numbers were believed to be accurate to within 2 percent.

A Low Mach Number Model for Moist Atmospheric Flows

DOE Office of Scientific and Technical Information (OSTI.GOV)

Duarte, Max; Almgren, Ann S.; Bell, John B.

A low Mach number model for moist atmospheric flows is introduced that accurately incorporates reversible moist processes in flows whose features of interest occur on advective rather than acoustic time scales. Total water is used as a prognostic variable, so that water vapor and liquid water are diagnostically recovered as needed from an exact Clausius–Clapeyron formula for moist thermodynamics. Low Mach number models can be computationally more efficient than a fully compressible model, but the low Mach number formulation introduces additional mathematical and computational complexity because of the divergence constraint imposed on the velocity field. Here in this paper, latentmore » heat release is accounted for in the source term of the constraint by estimating the rate of phase change based on the time variation of saturated water vapor subject to the thermodynamic equilibrium constraint. Finally, the authors numerically assess the validity of the low Mach number approximation for moist atmospheric flows by contrasting the low Mach number solution to reference solutions computed with a fully compressible formulation for a variety of test problems.« less

Preliminary Investigation of a New Type of Supersonic Inlet

NASA Technical Reports Server (NTRS)

Ferri, Antonio; Nucci, Louis M

1946-01-01

A supersonic inlet with supersonic deceleration of the flow entirely outside of the inlet is considered. A particular arrangement with fixed geometry having a central body with a circular annular intake is analyzed, and it is shown theoretically that this arrangement gives high pressure recovery for a large range of Mach number and mass flow and therefore is practical for use on supersonic airplanes and missiles. For some Mach numbers the drag coefficient for this type of inlet is larger than the drag coefficient for the type of inlet with supersonic compression entirely inside, but the pressure recovery is larger for all flight conditions. The differences in drag can be eliminated for the design Mach number. Experimental results confirm the results of the theoretical analysis and show that pressure recoveries of 95 percent for Mach numbers of 1.33 and 1.52, 92 percent for a Mach number of 1.72, and 86 percent for a Mach number oof 2.10 are possible with the configurations considered. If the mass flow decreases, the total drag coefficient increases gradually and the pressure recovery does not change appreciably.

Estimation of Static Longitudinal Stability of Aircraft Configurations at High Mach Numbers and at Angles of Attack Between 0 deg and +/-180 deg

NASA Technical Reports Server (NTRS)

Dugan, Duane W.

1959-01-01

The possibility of obtaining useful estimates of the static longitudinal stability of aircraft flying at high supersonic Mach numbers at angles of attack between 0 and +/-180 deg is explored. Existing theories, empirical formulas, and graphical procedures are employed to estimate the normal-force and pitching-moment characteristics of an example airplane configuration consisting of an ogive-cylinder body, trapezoidal wing, and cruciform trapezoidal tail. Existing wind-tunnel data for this configuration at a Mach number of 6.86 provide an evaluation of the estimates up to an angle of attack of 35 deg. Evaluation at higher angles of attack is afforded by data obtained from wind-tunnel tests made with the same configuration at angles of attack between 30 and 150 deg at five Mach numbers between 2.5 and 3.55. Over the ranges of Mach numbers and angles of attack investigated, predictions of normal force and center-of-pressure locations for the configuration considered agree well with those obtained experimentally, particularly at the higher Mach numbers.

Orbiter BLT Flight Experiment Wind Tunnel Simulations: Nearfield Flowfield Imaging and Surface Thermography

NASA Technical Reports Server (NTRS)

Danehy, Paul M.; Ivey, Christoper B.; Barthel, Brett F.; Inman, Jennifer A.; Jones, Stephen B.; Watkins, Anthony N.; Goodman, Kyle Z.; McCrea, Andrew C.; Leighty, Bradley D.; Lipford, William K.;

Pressure recovery performance of conical diffusers at high subsonic Mach numbers

NASA Technical Reports Server (NTRS)

Dolan, F. X.; Runstadler, P. W., Jr.

1973-01-01

The pressure recovery performance of conical diffusers has been measured for a wide range of geometries and inlet flow conditions. The approximate level and location (in terms of diffuser geometry of optimum performance were determined. Throat Mach numbers from low subsonic (m sub t equals 0.2) through choking (m sub t equals 1.0) were investigated in combination with throat blockage from 0.03 to 0.12. For fixed Mach number, performance was measured over a fourfold range of inlet Reynolds number. Maps of pressure recovery are presented as a function of diffuser geometry for fixed sets of inlet conditions. The influence of inlet blockage, throat Mach number, and inlet Reynolds number is discussed.

Wind-Tunnel and Flight Test Results for the Measurements of Flow Variables at Supersonic Speeds Using Improved Wedge and Conical Probes

NASA Technical Reports Server (NTRS)

Bobbitt, Percy J.; Maglieri, Domenic J.; Banks, Daniel W.; Frederick, Michael A.; Fuchs, Aaron W.

2012-01-01

The results of supersonic wind-tunnel tests on three probes at nominal Mach numbers of 1.6, 1.8 and 2.0 and flight tests on two of these probes up to a Mach number of 1.9 are described. One probe is an 8 deg. half-angle wedge with two total-pressure measurements and one static. The second, a conical probe, is a cylinder that has a 15 deg., semi-angle cone tip with one total-pressure orifice at the apex and four static-pressure orifices on the surface of the cone, 90 deg. apart, and about two-thirds of the distance from the cone apex to the base of the cone. The third is a 2 deg. semi-angle cone that has two static ports located 180 deg. apart about 1.5 inches behind the apex of the cone. The latter probe was included since it has been the "probe of choice" for wind-tunnel flow-field pressure measurements (or one similar to it) for the past half-century. The wedge and 15 deg. conical probes used in these tests were designed for flight diagnostic measurements for flight Mach numbers down to 1.35 and 1.15 respectively, and have improved capabilities over earlier probes of similar shape. The 15. conical probe also has a temperature sensor that is located inside the cylindrical part of the probe that is exposed to free-stream flow through an annulus at the apex of the cone. It enables the determination of free-stream temperature, density, speed of sound, and velocity, in addition to free-stream pressure, Mach number, angle of attack and angle of sideslip. With the time-varying velocity, acceleration can be calculated. Wind-tunnel tests of the two probes were made in NASA Langley Research Center fs Unitary Plan Wind Tunnel (UPWT) at Mach numbers of 1.6, 1.8, and 2.0. Flight tests were carried out at the NASA Dryden Flight Research Center (DFRC) on its F-15B aircraft up to Mach numbers of 1.9. The probes were attached to a fixture, referred to as the Centerline Instrumented Pylon (CLIP), under the fuselage of the aircraft. Problems controlling the velocity of the flow through the conical probe required for accurate temperature measurements are noted, as well as some calibration problems of the miniature pressure sensors that required a re-calculation of the flow variables. Data are presented for angle of attack, pressure and Mach number obtained in the wind tunnel and in flight. In the wind tunnel some transient data were obtained by translating the probes through the shock flow field created by a bump on the wind-tunnel wall.

A Wind-Tunnel Investigation of a Transonic-Transport Configuration Utilizing Drag-Reducing Devices at Mach Numbers from 0.20 to 1.03

NASA Technical Reports Server (NTRS)

Loving, Donald L.

1961-01-01

The static longitudinal stability and control and lateral characteristics of a transonic-transport model, incorporating recent drag-reducing devices, has been investigated in the Langley 8-foot transonic pressure tunnel. The wing was cambered, had a thickened root and a taper ratio of 0.3. Wing sweepback angles of 45 degrees and 40 degrees were investigated with corresponding aspect ratios of 7 and 8, respectively. Modifications to the model for reducing the drag were: a forward fuselage addition and special bodies (four big enough to house jet engines) added to the upper surface of the wing. Other components and changes investigated included an empennage, a wing-tip body, wing fences, wing trailing-edge flaps, horizontal-tail settings, and wing dihedral angle. The investigation covered the Mach number range from 0.20 to 1.03 for the angle-of-attack range from -5 degrees to 15.4 degrees, and a sideslip angle of -5 degrees, in the Reynolds number range from 0.52 times 10(exp 6) to 1.94 times 10(exp 6) based on the wing mean aerodynamic chord. The various fuselage and wing additions delayed the drag-rise Mach number and greatly reduced the drag beyond the drag rise. The wing bodies markedly alleviated unstable pitch tendencies throughout the test Mach number range. At low landing speeds, the wing bodies exhibited little interference with the ability of trailing-edge flaps to increase the lift near maximum lift coefficient; and the use of fences greatly reduced the severe longitudinal instability trend at landing attitudes. The model with a 6 degree dihedral angle exhibited positive lateral and directional stability characteristics in the presence of the fuselage and wing additions. An increase in drag-rise Mach number associated with the fuselage and wing additions on the 40 degree sweptback wing combination was similar to that for the comparable 45 degree combination. These additions did, however, reduce the drag of the 40 degree sweptback configurations more than the 45 degree configurations in the transonic speed range.

Heat transfer rate distribution on North American Rockwell delta wing orbiter determined by phase change paint technique at a Mach number of 8, volume 1

NASA Technical Reports Server (NTRS)

Matthews, R. K.; Martindale, W. R.; Warmbrod, J. D.

1972-01-01

The results of a wind tunnel test program to determine aerodynamic heat transfer distributions on an orbiter configuration are presented. Heat-transfer rates were determined by the phase change paint technique on 0.013-scale Stycast models using Tempilaq as the surface temperature indicator. The nominal test conditions were; Mach 8, length Reynolds numbers of 6.0 x 1 million and 8.9 x 1 million, and angles of attack from 10 to 50 deg in 10-deg increments. At the higher Reynolds number, data were obtained with and without boundary layer trips. Model details, test conditions, and reduced heat-transfer data are presented. Data reduction of the phase-change paint photographs was performed by utilizing a new technique which is described in the data presentation section.

Aerothermal tests of quilted dome models on a flat plate at a Mach number of 6.5

NASA Technical Reports Server (NTRS)

Glass, Christopher E.; Hunt, L. Roane

1988-01-01

Aerothermal tests were conducted in the NASA Langley 8 Foot High Temperature Tunnel (8'HTT) at a Mach number of 6.5 on simulated arrays of thermally bowed metallic thermal protection system (TPS) tiles at an angle of attack of 5 deg. Detailed surface pressures and heating rates were obtained for arrays aligned with the flow and skewed 45 deg diagonally to the flow with nominal bowed heights of 0.1, 0.2, and 0.4 inch submerged in both laminar and turbulent boundary layers. Aerothermal tests were made at a nominal total temperature of 3300 R, a total pressure of 400 psia, a total enthalpy of 950 Btu/lbm, a dynamic pressure of 2.7 psi, and a unit Reynolds number of 400,000 per foot. The experimental results form a data base that can be used to help protect aerothermal load increases from bowed arrays of TPS tiles.

Flight-measured base pressure coefficients for thick boundary-layer flow over an aft-facing step for Mach numbers from 0.4 to 2.5

NASA Technical Reports Server (NTRS)

Goecke, S. A.

1973-01-01

A 0.56-inch thick aft-facing step was located 52.1 feet from the leading edge of the left wing of an XB-70 airplane. A boundary-layer rake at a mirror location on the right wing was used to obtain local flow properties. Reynolds numbers were near 10 to the 8th power, resulting in a relatively thick boundary-layer. The momentum thickness ranged from slightly thinner to slightly thicker than the step height. Surface static pressures forward of the step were obtained for Mach numbers near 0.9, 1.5, 2.0, and 2.4. The data were compared with thin boundary-layer results from flight and wind-tunnel experiments and semiempirical relationships. Significant differences were found between the thick and the thin boundary-layer data.

Laser anemometer measurements in a transonic axial-flow fan rotor

NASA Technical Reports Server (NTRS)

Strazisar, Anthony J.; Wood, Jerry R.; Hathaway, Michael D.; Suder, Kenneth L.

1989-01-01

Laser anemometer surveys were made of the 3-D flow field in NASA rotor 67, a low aspect ratio transonic axial-flow fan rotor. The test rotor has a tip relative Mach number of 1.38. The flowfield was surveyed at design speed at near peak efficiency and near stall operating conditions. Data is presented in the form of relative Mach number and relative flow angle distributions on surfaces of revolution at nine spanwise locations evenly spaced from hub to tip. At each spanwise location, data was acquired upstream, within, and downstream of the rotor. Aerodynamic performance measurements and detailed rotor blade and annulus geometry are also presented so that the experimental results can be used as a test case for 3-D turbomachinery flow analysis codes.

Heat transfer tests of the McDonnell-Douglas delta wing orbiter mated with -17A booster at Mach number 8

NASA Technical Reports Server (NTRS)

Matthews, R. K.; Martindale, W. R.; Warmbrod, J. D.

1972-01-01

A wind tunnel test program to determine aerodynamic heat transfer distributions on the McDonnell-Douglas configurations is reported. The tests were conducted at the Arnold Engineering Development Center (AEDC) in Tunnel B of the von Karman Gas Dynamics Facility (VKF). Heat-transfer rates were determined by the phase-change paint technique on 0.011-scale Stycast models using Tempilaq as the surface temperature indicator. The nominal test conditions were; Mach 8, freestream unit Reynolds numbers of 0.8 x one million, 2.5 x one million, and 3.7 x one million, and angles of attack of -5 deg, 0 deg, +5deg. Model details, test conditions, phase-change paint photographs and reduced heat-transfer coefficients are presented.

Numerical models of jet disruption in cluster cooling flows

NASA Technical Reports Server (NTRS)

Loken, Chris; Burns, Jack O.; Roettiger, Kurt; Norman, Mike

1993-01-01

We present a coherent picture for the formation of the observed diverse radio morphological structures in dominant cluster galaxies based on the jet Mach number. Realistic, supersonic, steady-state cooling flow atmospheres are evolved numerically and then used as the ambient medium through which jets of various properties are propagated. Low Mach number jets effectively stagnate due to the ram pressure of the cooling flow atmosphere while medium Mach number jets become unstable and disrupt in the cooling flow to form amorphous structures. High Mach number jets manage to avoid disruption and are able to propagate through the cooling flow.

Study by the Prandtl-Glauert method of compressibility effects and critical Mach number for ellipsoids of various aspect ratios and thickness ratios

NASA Technical Reports Server (NTRS)

Hess, Robert V; Gardner, Clifford S

1947-01-01

By using the Prandtl-Glauert method that is valid for three-dimensional flow problems, the value of the maximum incremental velocity for compressible flow about thin ellipsoids at zero angle of attack is calculated as a function of the Mach number for various aspect ratios and thickness ratios. The critical Mach numbers of the various ellipsoids are also determined. The results indicate an increase in critical Mach number with decrease in aspect ratio which is large enough to explain experimental results on low-aspect-ratio wings at zero lift.

The development of cambered airfoil sections having favorable lift characteristics at supercritical Mach numbers

NASA Technical Reports Server (NTRS)

Graham, Donald J

1949-01-01

Several groups of new airfoil sections, designated as the NACA 8-series, are derived analytically to have lift characteristics at supercritical Mach numbers which are favorable in the sense that the abrupt loss of lift, characteristic of the usual airfoil section at Mach numbers above the critical, is avoided. Aerodynamic characteristics determined from two-dimensional wind-tunnel tests at Mach numbers up to approximately 0.9 are presented for each of the derived airfoils. Comparisons are made between the characteristics of these airfoils and the corresponding characteristics of representative NACA 6-series airfoils.

The shocking properties of supersonic flows: Dependence of the thermal overstability on M, α, and Tc / T0

NASA Astrophysics Data System (ADS)

Pittard, J. M.; Dobson, M. S.; Durisen, R. H.; Dyson, J. E.; Hartquist, T. W.; O'Brien, J. T.

2005-07-01

We present hydrodynamical calculations of radiative shocks with low Mach numbers and find that the well-known global overstability can occur if the temperature exponent (α) of the cooling is sufficiently negative. We find that the stability of radiative shocks increases with decreasing Mach number, with the result that M=2 shocks require α ⪉ -1.2 in order to be overstable. Such values occur within a limited temperature range of many cooling curves. We observe that Mach numbers of order 100 are needed before the strong shock limit of α_cr ≈ 0.4 is reached, and we discover that the frequency of oscillation of the fundamental mode also has a strong Mach number dependence. We find that feedback between the cooling region and the cold dense layer (CDL) further downstream is a function of Mach number, with stronger feedback and oscillation of the boundary between the CDL and the cooling region occuring at lower Mach numbers. This feedback can be quantified in terms of the reflection coefficient of sound waves, and in those cases where the cooling layer completely disappears at the end of each oscillation cycle, the initial velocity of the waves driven into the upstream pre-shock flow and into the downstream CDL, and the velocity of the the boundary between the CDL and the cooling layer, can be understood in terms of the solution to the Riemann problem. An interesting finding is that the stability properties of low Mach number shocks can be dramatically altered if the shocked gas is able to cool to temperatures less than the pre-shock value (i.e. when χ < 1, where χ is the ratio of the temperature of the cold dense layer to the pre-shock temperature). In such circumstances, low Mach number shocks have values of α_cr which are comparable to values obtained for higher Mach number shocks when χ = 1. For instance, α_cr=-0.1 when M=2 and χ=0.1, comparable to that when M=10 and χ=1. Thus, it is probable that low Mach number astrophysical shocks will be overstable in a variety of situations. We also explore the effect of different assumptions for the initial hydrodynamic set up and the type of boundary condition imposed downstream, and find that the properties of low Mach number shocks are relatively insensitive to these issues. The results of this work are relevant to astrophysical shocks with low Mach numbers, such as supernova remnants (SNRs) immersed in a hot interstellar medium (e.g., within a starburst region), and shocks in molecular clouds, where time-dependent chemistry can lead to overstability.

Aerodynamics of a Gulfstream G550 Nose Landing Gear Model

NASA Technical Reports Server (NTRS)

Neuhart, Dan H.; Khorrami, Mehdi R.; Choudhari, Meelan M.

2009-01-01

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.

Visualization of Flow Separation Around an Atmospheric Entry Capsule at Low-Subsonic Mach Number Using Background-Oriented Schlieren (BOS)

NASA Technical Reports Server (NTRS)

Mizukaki, Toshiharu; Borg, Stephen E.; Danehy, Paul M.; Murman, Scott M.

2014-01-01

This paper presents the results of visualization of separated flow around a generic entry capsule that resembles the Apollo Command Module (CM) and the Orion Multi-Purpose Crew Vehicle (MPCV). The model was tested at flow speeds up to Mach 0.4 at a single angle of attack of 28 degrees. For manned spacecraft using capsule-shaped vehicles, certain flight operations such as emergency abort maneuvers soon after launch and flight just prior to parachute deployment during the final stages of entry, the command module may fly at low Mach number. Under these flow conditions, the separated flow generated from the heat-shield surface on both windward and leeward sides of the capsule dominates the wake flow downstream of the capsule. In this paper, flow visualization of the separated flow was conducted using the background-oriented schlieren (BOS) method, which has the capability of visualizing significantly separated wake flows without the particle seeding required by other techniques. Experimental results herein show that BOS has detection capability of density changes on the order of 10(sup-5).

Symposium on Numerical and Physical Aspects of Aerodynamic Flows

DTIC Science & Technology

1992-01-15

ReT/R. -1 wide range of Mach numbers including pressure gradient, a = - (a*) (10) surface roughness. surface heating and cooling . and surface 9 I + ReT...and specific heat at constant pressure, respectively. Favre equation is then obtained by contracting the Reynolds-stress decomposition is applied to...their near- wall behavior. In other words, if highly cooled -wall flows are to be predicted correctly, heat fluxes should be modeled e = 2ak + 4bky

Aeroacoustic Study of a High-Fidelity Aircraft Model. Part 2; Unsteady Surface Pressures

NASA Technical Reports Server (NTRS)

Khorrami, Mehdi R.; Neuhart, Danny H.

2012-01-01

In this paper, we present unsteady surface pressure measurements for an 18%-scale, semi-span Gulfstream aircraft model. This high-fidelity model is being used to perform detailed studies of airframe noise associated with main landing gear, flap components, and gear-flap interaction noise, as well as to evaluate novel noise reduction concepts. The aerodynamic segment of the tests, conducted in the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel, was completed in November 2010. To discern the characteristics of the surface pressure fluctuations in the vicinity of the prominent noise sources, unsteady sensors were installed on the inboard and outboard flap edges, and on the main gear wheels, struts, and door. Various configurations were tested, including flap deflections of 0?, 20?, and 39?, with and without the main landing gear. The majority of unsteady surface pressure measurements were acquired for the nominal landing configuration where the main gear was deployed and the flap was deflected 39?. To assess the Mach number variation of the surface pressure amplitudes, measurements were obtained at Mach numbers of 0.16, 0.20, and 0.24. Comparison of the unsteady surface pressures with the main gear on and off shows significant interaction between the gear wake and the inboard flap edge, resulting in higher amplitude fluctuations when the gear is present.

Measurements of Aerodynamic Heat Transfer and Boundary-Layer Transition on a 15 deg. Cone in Free Flight at Supersonic Mach Numbers up to 5.2

NASA Technical Reports Server (NTRS)

Rumsey, Charles B.; Lee, Dorothy B.

1961-01-01

Measurements of aerodynamic heat transfer have been made at several stations on the 15 deg total-angle conical nose of a rocket-propelled model in free flight at Mach numbers up to 5.2. Data are presented for a range of local Mach number just outside the boundary layer from 1.40 to 4.65 and a range of local Reynolds number from 3.8 x 10(exp 6) to 46.5 x 10(exp 6), based on length from the nose tip to a measurement station. Laminar, transitional, and turbulent heat-transfer coefficients were measured. The laminar data were in agreement with laminar theory for cones, and the turbulent data agreed well with turbulent theory for cones using Reynolds number based on length from the nose tip. At a nearly constant ratio of wall to local static temperature of 1.2 the Reynolds number of transition increased from 14 x 10(exp 6) to 30 x 10(exp 6) as Mach number increased from 1.4 to 2.9 and then decreased to 17 x 10(exp 6) as Mach number increased to 3.7. At Mach numbers near 3.5, transition Reynolds numbers appeared to be independent of skin temperature at skin temperatures very cold with respect to adiabatic wall temperature. The transition Reynolds number was 17.7 x 10(exp 6) at a condition of Mach number and ratio of wall to local static temperature near that for which three-dimensional disturbance theory has been evaluated and has predicted laminar boundary-layer stability to very high Reynolds numbers (approximately 10(exp 12)).

14 CFR Appendix J to Part 36 - Alternative Noise Certification Procedure for Helicopters Under Subpart H Having a Maximum...

Code of Federal Regulations, 2010 CFR

2010-01-01

... produce the same advancing blade tip Mach number as associated with the reference conditions; (i) Advancing blade tip Mach number (MAT) is defined as the ratio of the arithmetic sum of blade tip rotational... the reference advancing blade tip Mach number. The adjusted reference airspeed shall be maintained...

14 CFR Appendix J to Part 36 - Alternative Noise Certification Procedure for Helicopters Under Subpart H Having a Maximum...

Code of Federal Regulations, 2011 CFR

2011-01-01

... produce the same advancing blade tip Mach number as associated with the reference conditions; (i) Advancing blade tip Mach number (MAT) is defined as the ratio of the arithmetic sum of blade tip rotational... the reference advancing blade tip Mach number. The adjusted reference airspeed shall be maintained...

14 CFR Appendix J to Part 36 - Alternative Noise Certification Procedure for Helicopters Under Subpart H Having a Maximum...

Code of Federal Regulations, 2012 CFR

2012-01-01

... produce the same advancing blade tip Mach number as associated with the reference conditions; (i) Advancing blade tip Mach number (MAT) is defined as the ratio of the arithmetic sum of blade tip rotational... the reference advancing blade tip Mach number. The adjusted reference airspeed shall be maintained...

14 CFR Appendix J to Part 36 - Alternative Noise Certification Procedure for Helicopters Under Subpart H Having a Maximum...

Code of Federal Regulations, 2013 CFR

2013-01-01

... produce the same advancing blade tip Mach number as associated with the reference conditions; (i) Advancing blade tip Mach number (MAT) is defined as the ratio of the arithmetic sum of blade tip rotational... the reference advancing blade tip Mach number. The adjusted reference airspeed shall be maintained...

14 CFR Appendix J to Part 36 - Alternative Noise Certification Procedure for Helicopters Under Subpart H Having a Maximum...

Code of Federal Regulations, 2014 CFR

2014-01-01

... produce the same advancing blade tip Mach number as associated with the reference conditions; (i) Advancing blade tip Mach number (MAT) is defined as the ratio of the arithmetic sum of blade tip rotational... the reference advancing blade tip Mach number. The adjusted reference airspeed shall be maintained...

Blunt Body Near-Wake Flow Field at Mach 10

NASA Technical Reports Server (NTRS)

Horvath, Thomas; Hannemann, Klaus

1997-01-01

Tests were conducted in a Mach 10 air flow to examine the reattachment process of a free shear layer associated with the near wake of a 70 deg half angle, spherically blunted cone having a cylindrical after body. The nominal free-stream Reynolds number based on model diameter ranged from 0.25 x l0(exp 6) to 1 x l0(exp 6) and the angle of incidence set at 0 and +/- 20 deg. The present study was designed to complement previously reported Mach 6 perfect air tests as well as results obtained in several hypervelocity facilities capable of producing real gas effects. Surface heating rates were inferred from temperature time histories from coaxial surface thermocouples on the model forebody and thin film resistance gages along the model base and cylindrical after body. Limited forebody, base, and support sting surface pressures were obtained with piezoresistive Experimental results are compared to laminar perfect gas predictions provided by a 3-0 Navier Stokes code (NSHYP). Shear layer impingement on the instrumented cylindrical after body resulted in a localized heating maximum that was 16 to 18percent of the forebody stagnation point and a factor of 2 higher than laminar predictions, suggesting a transitional or turbulent shear layer. transducers.

An Investigation of the Effects of Nose and Lip Shapes for an Underslung Scoop Inlet at Mach Numbers from 0 to 1.9

NASA Technical Reports Server (NTRS)

Pfyl, Frank A.

1955-01-01

An experimental investigation was conducted to determine the performance characteristics an underslung nose-scoop air-induction system for a supersonic airplane. Five different nose shapes, three lip shapes, and two internal diffusers were investigated. Tests were made at Mach numbers from 0 to 1.9, angles of attack from 0 deg to approximately l5 deg, and mass-flow ratios from 0 to maximum obtainable. It was found that the underslung nose-scoop inlet was able to operate at Mach numbers from 0.6 to 1.9 over a large positive angle-of-attack range without adverse effects on the pressure recovery. Although there was no one inlet configuration that was markedly superior over the entire range of operating variables, the arrangement having a nose designed to give increased supersonic compression at low angles of attack, and a sharp lip (configuration designated N3L3) showed the most favorable performance characteristics over the supersonic Mach number range. Inlets with sizable lip radii gave satisfactory performance up to a Mach number of 1.5; however, as a result of an increase in drag, the performance of such inlets was markedly inferior to the sharp-lip configuration above Mach numbers of 1.5. Throughout the range of test Mach numbers all inlet configurations evidenced stable air-flow characteristics over the mass-flow range for normal engine operation. Analysis of the inlet performance on the basis of a propulsive thrust parameter showed that a fixed inlet area could be used for Mach numbers up to 1.5 with only a small sacrifice in performance.

Turbulent Heat-Transfer Coefficients in the Vicinity of Surface Protuberances

NASA Technical Reports Server (NTRS)

Wisniewski, Richard J.

1958-01-01

Local turbulent heating rates were obtained in the vicinity of surface protuberances mounted on the cylinder section of a cone cylinder model at a Mach number of 3.12. Data were obtained at Reynolds number per foot of 4.5 and 6 million for an unswept cylinder, a 45 deg swept cylinder, a 45 deg elbow, and several 90 deg elbows. The unswept cylinder and the 90 deg elbows increased the local turbulent heating rates in the vicinity of the surface protuberances. The data of the 45 deg swept cylinder and the 45 deg elbow resulted in heating rates lower than those observed without surface protuberances. In general, sweeping a surface protuberance resulted in heating rates comparable or lower than those measured without surface protuberances.

Mach-like capillary-gravity wakes.

PubMed

Moisy, Frédéric; Rabaud, Marc

2014-08-01

We determine experimentally the angle α of maximum wave amplitude in the far-field wake behind a vertical surface-piercing cylinder translated at constant velocity U for Bond numbers Bo(D)=D/λ(c) ranging between 0.1 and 4.2, where D is the cylinder diameter and λ(c) the capillary length. In all cases the wake angle is found to follow a Mach-like law at large velocity, α∼U(-1), but with different prefactors depending on the value of Bo(D). For small Bo(D) (large capillary effects), the wake angle approximately follows the law α≃c(g,min)/U, where c(g,min) is the minimum group velocity of capillary-gravity waves. For larger Bo(D) (weak capillary effects), we recover a law α∼√[gD]/U similar to that found for ship wakes at large velocity [Rabaud and Moisy, Phys. Rev. Lett. 110, 214503 (2013)]. Using the general property of dispersive waves that the characteristic wavelength of the wave packet emitted by a disturbance is of order of the disturbance size, we propose a simple model that describes the transition between these two Mach-like regimes as the Bond number is varied. We show that the new capillary law α≃c(g,min)/U originates from the presence of a capillary cusp angle (distinct from the usual gravity cusp angle), along which the energy radiated by the disturbance accumulates for Bond numbers of order of unity. This model, complemented by numerical simulations of the surface elevation induced by a moving Gaussian pressure disturbance, is in qualitative agreement with experimental measurements.

Free-Flight Investigation of the Static and Dynamic Longitudinal Stability Characteristics of 1/3.7-Scale Rocket-Powered Models of the Bell MX-776A

NASA Technical Reports Server (NTRS)

Michal, David H.

1950-01-01

An investigation of the static and dynamic longitudinal stability characteristics of 1/3.7 scale rocket-powered model of the Bell MX-776A has been made for a Mach number range from 0.8 to 1.6. Two models were tested with all control surfaces at 0 degree deflection and centers of gravity located 1/4 and 1/2 body diameters, respectively, ahead of the equivalent design location. Both models were stable about the trim conditions but did not trim at 0 degree angle of attack because of slight constructional asymmetries. The results indicated that the variation of lift and pitching moment was not linear with angle of attack. Both lift-curve slope and pitching-moment-curve slope were of the smallest magnitude near 0 degree angle of attack. In general, an increase in angle of attack was accompanied by a rearward movement of the aerodynamic center as the rear wing moved out of the downwash from the forward surfaces. This characteristic was more pronounced in the transonic region. The dynamic stability in the form of total damping factor varied with normal-force coefficient but was greatest for both models at a Mach number of approximately 1.25. The damping factor was greater at the lower trim normal-force coefficients except at a Mach number of 1.0. At that speed the damping factor was of about the same magnitude for both models. The drag coefficient increased with trim normal-force coefficient and was largest in the transonic region.

Film cooling effectiveness on a large angle blunt cone flying at hypersonic speed

NASA Astrophysics Data System (ADS)

Sahoo, Niranjan; Kulkarni, Vinayak; Saravanan, S.; Jagadeesh, G.; Reddy, K. P. J.

2005-03-01

Effectiveness of film cooling technique to reduce convective heating rates for a large angle blunt cone flying at hypersonic Mach number and its effect on the aerodynamic characteristics is investigated experimentally by measuring surface heat-transfer rates and aerodynamic drag coefficient simultaneously. The test model is a 60° apex-angle blunt cone with an internally mounted accelerometer balance system for measuring aerodynamic drag and an array of surface mounted platinum thin film gauges for measuring heat-transfer rates. The coolant gas (air, carbon dioxide, and/or helium) is injected into the hypersonic flow at the nose of the test model. The experiments are performed at a flow free stream Mach number of 5.75 and 0° angle of attack for stagnation enthalpies of 1.16MJ/kg and 1.6MJ/kg with and without gas injection. About 30%-45% overall reduction in heat-transfer rates is observed with helium as coolant gas except at stagnation regions. With all other coolants, the reduction in surface heat-transfer rate is between 10%-25%. The aerodynamic drag coefficient is found to increase by 12% with helium injection whereas with other gases this increase is about 27%.

CFD code calibration and inlet-fairing effects on a 3D hypersonic powered-simulation model

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Tatum, Kenneth E.

1993-01-01

A three-dimensional (3D) computational study has been performed addressing issues related to the wind tunnel testing of a hypersonic powered-simulation model. The study consisted of three objectives. The first objective was to calibrate a state-of-the-art computational fluid dynamics (CFD) code in its ability to predict hypersonic powered-simulation flows by comparing CFD solutions with experimental surface pressure dam. Aftbody lower surface pressures were well predicted, but lower surface wing pressures were less accurately predicted. The second objective was to determine the 3D effects on the aftbody created by fairing over the inlet; this was accomplished by comparing the CFD solutions of two closed-inlet powered configurations with a flowing-inlet powered configuration. Although results at four freestream Mach numbers indicate that the exhaust plume tends to isolate the aftbody surface from most forebody flowfield differences, a smooth inlet fairing provides the least aftbody force and moment variation compared to a flowing inlet. The final objective was to predict and understand the 3D characteristics of exhaust plume development at selected points on a representative flight path. Results showed a dramatic effect of plume expansion onto the wings as the freestream Mach number and corresponding nozzle pressure ratio are increased.

CFD Code Calibration and Inlet-Fairing Effects On a 3D Hypersonic Powered-Simulation Model

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Tatum, Kenneth E.

1993-01-01

A three-dimensional (3D) computational study has been performed addressing issues related to the wind tunnel testing of a hypersonic powered-simulation model. The study consisted of three objectives. The first objective was to calibrate a state-of-the-art computational fluid dynamics (CFD) code in its ability to predict hypersonic powered-simulation flows by comparing CFD solutions with experimental surface pressure data. Aftbody lower surface pressures were well predicted, but lower surface wing pressures were less accurately predicted. The second objective was to determine the 3D effects on the aftbody created by fairing over the inlet; this was accomplished by comparing the CFD solutions of two closed-inlet powered configurations with a flowing- inlet powered configuration. Although results at four freestream Mach numbers indicate that the exhaust plume tends to isolate the aftbody surface from most forebody flow- field differences, a smooth inlet fairing provides the least aftbody force and moment variation compared to a flowing inlet. The final objective was to predict and understand the 3D characteristics of exhaust plume development at selected points on a representative flight path. Results showed a dramatic effect of plume expansion onto the wings as the freestream Mach number and corresponding nozzle pressure ratio are increased.

Exploratory Investigation of the Effects of Boundary-Layer Control on the Pressure-Recovery Characteristics of a Circular Internal-Contraction Inlet with Translating Centerbody at Mach Numbers of 2.00 and 2.35

NASA Technical Reports Server (NTRS)

Martin, Norman J.

1959-01-01

Exploratory tests of a circular internal-contraction inlet were made at Mach numbers of 2.00 and 2.35 to determine the effect of a cowl-type boundary-layer control located downstream of the inlet throat. The inlet was designed for a Mach number of 2.5. Tests were also made of the inlet modified to correspond to design Mach numbers of 2.35 and 2.25. Surveys near the minimum area section of the inlet without boundary-layer control indicated maximum averaged pressure recoveries between 0.90 and 0.92 at a free-stream Mach number, M(sub infinity), of 2.35 for the inlets. Farther downstream, after partial subsonic diffusion, a maximum pressure recovery of 0.842 was obtained with the inlet at M(sub infinity) = 2.35. The pressure recovery of the inlet was increased by 0.03 at a Mach number of 2.35 and decreased by 0.02 at a Mach number of 2.00 by the application of cowl-type boundary-layer control. Further investigation with the inlet without bleed demonstrated that an increase of angle of attack from 0 deg to 3 deg reduced the pressure recovery 0.04. The effect of Reynolds number was to increase pressure recovery 0.07 (from 0.785 to 0.855) with an increase in Reynolds number (based on inlet diameter) from 0.79 x 10(exp 6) to 3.19 x 10(exp 6).

Ram-recovery Characteristics of NACA Submerged Inlets at High Subsonic Speeds

NASA Technical Reports Server (NTRS)

Hall, Charles F; Frank, Joseph L

1948-01-01

Results are presented of an experimental investigation of the characteristics of NACA submerged inlets on a model of a fighter airplane for Mach numbers from 0.30 to 0.875. The effects on the ram-recovery ratio at the inlets of Mach number, angle of attack, boundary-layer thickness on the fuselage, inlet location, and boundary-layer deflectors are shown. The data indicate only a slight decrease in ram-recovery ratio for the inlets ahead of or just behind the wing leading edge as Mach number increased, but showed large decreases at high Mach numbers for the inlets aft of the point of maximum thickness of the wing.

Sidewall Mach Number Distributions for the NASA Langley Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Florance, James R.; Rivera, Jose A., Jr.

2001-01-01

The Transonic Dynamics Tunnel(TDT) was recalibrated due to the conversion of the heavy gas test medium from R-12 to R-134a. The objectives of the tests were to determine the relationship between the free-stream Mach number and the measured test section Mach number, and to quantify any necessary corrections. Other tests included the measurement of pressure distributions along the test-section walls, test-section centerline, at certain tunnel stations via a rake apparatus, and in the tunnel settling chamber. Wall boundary layer, turbulence, and flow angularity measurements were also performed. This paper discusses the determination of sidewall Mach number distributions.

Sensitivity of transport aircraft performance and economics to advanced technology and cruise Mach number

NASA Technical Reports Server (NTRS)

Ardema, M. D.

1974-01-01

Sensitivity data for advanced technology transports has been systematically collected. This data has been generated in two separate studies. In the first of these, three nominal, or base point, vehicles designed to cruise at Mach numbers .85, .93, and .98, respectively, were defined. The effects on performance and economics of perturbations to basic parameters in the areas of structures, aerodynamics, and propulsion were then determined. In all cases, aircraft were sized to meet the same payload and range as the nominals. This sensitivity data may be used to assess the relative effects of technology changes. The second study was an assessment of the effect of cruise Mach number. Three families of aircraft were investigated in the Mach number range 0.70 to 0.98: straight wing aircraft from 0.70 to 0.80; sweptwing, non-area ruled aircraft from 0.80 to 0.95; and area ruled aircraft from 0.90 to 0.98. At each Mach number, the values of wing loading, aspect ratio, and bypass ratio which resulted in minimum gross takeoff weight were used. As part of the Mach number study, an assessment of the effect of increased fuel costs was made.

Experimental evaluation of wall Mach number distributions of the octagonal test section proposed for NASA Lewis Research Center's altitude wind tunnel

NASA Technical Reports Server (NTRS)

Harrington, Douglas E.; Burley, Richard R.; Corban, Robert R.

1986-01-01

Wall Mach number distributions were determined over a range of test-section free-stream Mach numbers from 0.2 to 0.92. The test section was slotted and had a nominal porosity of 11 percent. Reentry flaps located at the test-section exit were varied from 0 (fully closed) to 9 (fully open) degrees. Flow was bled through the test-section slots by means of a plenum evacuation system (PES) and varied from 0 to 3 percent of tunnel flow. Variations in reentry flap angle or PES flow rate had little or no effect on the Mach number distributions in the first 70 percent of the test section. However, in the aft region of the test section, flap angle and PES flow rate had a major impact on the Mach number distributions. Optimum PES flow rates were nominally 2 to 2.5 percent wtih the flaps fully closed and less than 1 percent when the flaps were fully open. The standard deviation of the test-section wall Mach numbers at the optimum PES flow rates was 0.003 or less.

Experimental and Computational Evaluation of Flush-Mounted, S-Duct Inlets

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Allan, Brian G.

2004-01-01

A new high Reynolds number test capability for boundary layer ingesting inlets has been developed for the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. Using this new capability. an experimental investigation of four S-duct inlet configurations was conducted. A computational study of one of the inlets was also conducted using a Navier-Stokes solver. The objectives of this investigation were to: 1) develop a new high Reynolds number inlet test capability for flush-mounted inlets; 2) provide a database for CFD tool validation; 3) evaluate the performance of S-duct inlets with large amounts of boundary layer ingestion; and 4) provide a baseline inlet for future inlet flow-control studies. Tests were conducted at Mach numbers from 0.25 to 0.83. Reynolds numbers (based on duct exit diameter) from 5.1 million to a full-scale value of 13.9 million, and inlet mass-flow ratios from 0.39 to 1.58 depending on Mach number. Results of the experimental study indicate that inlet pressure recovery generally decreased and inlet distortion generally increased with increasing Mach number. Except at low Mach numbers, increasing inlet mass-flow increased pressure recovery and increased distortion. Increasing the amount of boundary layer ingestion or ingesting a boundary layer with a distorted profile decreased pressure recovery and increased distortion. Finally, increasing Reynolds number had almost no effect on inlet distortion but increased inlet recovery by about one-half percent at a Mach number near cruise. The computational results captured the inlet pressure recovery and distortion trends with Mach number and inlet mass-flow well: the reversal of the pressure recovery trend with increasing inlet mass-flow at low and high Mach numbers was predicted by CFD. However, CFD results were generally more pessimistic (larger losses) than measured experimentally.

Mach 6 flowfield survey at the engine inlet of a research airplane

NASA Technical Reports Server (NTRS)

Johnson, C. B.; Lawing, P. L.

1977-01-01

A flowfield survey was conducted to better define the nature of vehicle forebody flowfield at the inlet location of an airframe-integrated scramjet engine mounted on the lower surface of a high-speed research airplane to be air launched from a B-52 and rocket boosted to Mach 6. The tests were conducted on a 1/30-scale brass model in a Mach-6 20-in. wind tunnel at Reynolds number of 11,200,000 based on distance to engine inlet. Boundary layer profiles at five spanwise locations indicate that the boundary layer in the area of the forebody centerline is more than twice as thick as the boundary layer at three outboard stations. It is shown that the cold streak found in heating contours on the centerline of the forebody is caused by a thickening of the boundary layer on the centerline, and that this thickening decreases with angle of attack.

Flow quality of NAL two-dimensional transonic wind tunnel. Part 1: Mach number distributions, flow angularities and preliminary study of side wall boundary layer suction

NASA Technical Reports Server (NTRS)

Sakakibara, Seizo; Takashima, Kazuaki; Miwa, Hitoshi; Oguni, Yasuo; Sato, Mamoru; Kanda, Hiroshi

1988-01-01

Experimental data on the flow quality of the National Aerospace Laboratory two-dimensional transonic wind tunnel are presented. Mach number distributions on the test section axis show good uniformity which is characterized by the two sigma (standard deviation) values of 0.0003 to 0.001 for a range of Mach numbers from 0.4 to 1.0. Flow angularities, which were measured by using a wing model with a symmetrical cross section, remained within 0.04 deg for Mach numbers from 0.2 to 0.8. Side wall boundary layer suction was applied through a pair of porous plates. The variation of aerodynamic properties of the model due to the suction mass flow rate change is presented with a brief discussion. Two dimensionality of the flow over the wing span is expected to be improved by applying the appropriate suction rate, which depends on the Mach number, Reynolds number, and lift coefficient.

Experimental and theoretical study of shuttle lee-side heat transfer rates

NASA Technical Reports Server (NTRS)

Mruk, G. K.; Bertin, J.; Lamb, J. P.

1975-01-01

The experimental program which was conducted in the Calspan 96-inch hypersonic shock tunnel to investigate what effect the windward surface temperature had on the heat transfer to the leeward surface of the space shuttle orbiter is discussed. Heat-transfer distributions, surface-pressure distributions, and schlieren photographs were obtained for an 0.01-scale model of the 139 configuration space shuttle orbiter at angles-of-attack of 30 and 40 deg. Similar data were obtained for an 0.01 scale wingless model of the 139 configuration at angles-of-attack of 30 and 90 deg. Data were obtained for Mach numbers from Reynolds numbers, and surface temperatures and compared with theoretical results.

A Transonic Wind-Tunnel Investigation of the Longitudinal Aerodynamic Characteristics of a Model of the Lockheed XF-104 Airplane

NASA Technical Reports Server (NTRS)

Hieser, Gerald; Reid, Charles F.

1954-01-01

The transonic longitudinal aerodynamic characteristics of a 0.0858-scale model of the Lockheed XF-104 airplane have been obtained from tests at the Langley 16-foot transonic tunnel. The results of the investigation provide some general information applicable to the transonic properties of thin, low-aspect-ratio, unswept wing configurations utilizing a high horizontal tail . The model employs a horizontal tail mounted at the top of the vertical tail and a wing with an aspect ratio of 2.5, a taper ratio of 0.385, and 3.4-percent-thick airfoil sections. The lift, drag, and static longitudinal pitching moment were measured at Mach numbers from 0.80 t o 1.09 and angles of attack from -2.5 deg to 22.5 deg. Some of the dynamic longitudinal stability properties of the airplane have been predicted from the test results. In addition, some visual flow studies on the wing surfaces obtained at Mach numbers of 0.80 and 1.00 are included. Results of the investigation show that the transonic rise in drag coefficient at zero lift is about 0.030. At high angles of attack, the model becomes longitudinally unstable at Mach numbers from 0.80 t o 0.90, whereas a reduction in static stability is experienced when very high angles of attack are reached at Mach numbers above 0.90. Longitudinal dynamic stability calculations show that the longitudinal control is good at angles of attack below the unstable break in the static pitching-moment curves, but a typical corrective control applied after the occurrence of neutral stability has little effect in averting pitch-up.

Exhaust Simulation Testing of a Hypersonic Airbreathing Model at Transonic Speeds

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Witte, David W.; Andrews, Earl H., Jr.

2004-01-01

An experimental study was performed to examine jet-effects for an airframe-integrated, scramjet-rocket combined-cycle vehicle configuration at transonic test conditions. This investigation was performed by testing an existing exhaust simulation wind tunnel model, known as Model 5B, in the NASA Langley 16-Ft. Transonic Tunnel. Tests were conducted at freestream Mach numbers from 0.7 to 1.2, at angles of attack from 2 to +14 degrees, and at up to seven nozzle static pressure ratio values for a set of horizontal-tail and body-flap deflections. The model aftbody, horizontal tails, and body flaps were extensively pressure instrumented to provide an understanding of jet-effects and control-surface/plume interactions, as well as for the development of analytical methodologies and calibration of computational fluid dynamic codes to predict this type of flow phenomenon. At all transonic test conditions examined, the exhaust flow at the exit of the internal nozzle was over-expanded, generating an exhaust plume that turned toward the aftbody. Pressure contour plots for the aftbody of Model 5B are presented for freestream transonic Mach numbers of 0.70, 0.95, and 1.20. These pressure data, along with shadowgraph images, indicated the impingement of an internal plume shock and at least one reflected shock onto the aftbody for all transonic conditions tested. These results also provided evidence of the highly three-dimensional nature of the aftbody exhaust flowfield. Parametric testing showed that angle-of-attack, static nozzle pressure ratio, and freestream Mach number all affected the exhaust-plume size, exhaust-flowfield shock structure, and the aftbody-pressure distribution, with Mach number having the largest effect. Integration of the aftbody pressure data showed large variations in the pitching moment throughout the transonic regime.

Flight Investigation at Low Angles of Attack to Determine the Longitudinal Stability and Control Characteristics of a Cruciform Canard Missile Configuration with a Low-Aspect-Ratio Wing and Blunt Nose at Mach Numbers from 1.2 to 2.1

NASA Technical Reports Server (NTRS)

Brown, Clarence A , Jr

1957-01-01

A full- scale rocket-powered model of a cruciform canard missile configuration with a low- aspect - ratio wing and blunt nose has been flight tested by the Langley Pilotless Aircraft Research Division. Static and dynamic longitudinal stability and control derivatives of this interdigitated canard-wing missile configuration were determined by using the pulsed- control technique at low angles of attack and for a Mach number range of 1.2 to 2.1. The lift - curve slope showed only small nonlinearities with changes in control deflection or angle of attack but indicated a difference in lift- .curve slope of approximately 7 percent for the two control deflections of delta = 3.0 deg and delta= -0.3 deg . The large tail length of the missile tested was effective in producing damping in pitch throughout the Mach number range tested. The aerodynamic- center location was nearly constant with Mach number for the two control deflections but was shown to be less stable with the larger control deflection. The increment of lift produced by the controls was small and positive throughout the Mach number range tested, whereas the pitching moment produced by the controls exhibited a normal trend of reduced effectiveness with increasing Mach number.The effectiveness of the controls in producing angle of attack, lift, and pitching moment was good at all Mach numbers tested.

Flight Investigation at Low Angles of Attack to Determine the Longitudinal Stability and Control Characteristics of a Cruciform Canard Missile Configuration with a Low-Aspect-Ratio Wing and Blunt Nose at Mach Numbers from 1.2 to 2.1

NASA Technical Reports Server (NTRS)

Brown, C. A., Jr.

1957-01-01

A full-scale rocket-powered model of a cruciform canard missile configuration with a low-aspect-ratio wing and blunt nose has been flight tested by the Langley Pilotless Aircraft Research Division. Static and dynamic longitudinal stability and control derivatives of this interdigitated canard-wing missile configuration were determined by using the pulsed-control technique at low angles of attack and for a Mach number range of 1.2 to 2.1. The lift-curve slope showed only small nonlinearities with changes in control deflection or angle of attack but indicated a difference in lift-curve slope of approximately 7 percent for the two control deflections of delta = 3.0 deg and delta = -0.3 deg. The large tail length of the missile tested was effective in producing damping in pitch throughout the Mach number range tested. The aerodynamic-center location was nearly constant with Mach number for the two control deflections but was shown to be less stable with the larger control deflection. The increment of lift produced by the controls was small and positive throughout the Mach number range tested, whereas the pitching moment produced by the controls exhibited a normal trend of reduced effectiveness with increasing Mach number. The effectiveness of the controls in producing angle of attack, lift, and pitching moment was good at all Mach numbers tested.

Krypton tagging velocimetry in a turbulent Mach 2.7 boundary layer

NASA Astrophysics Data System (ADS)

Zahradka, D.; Parziale, N. J.; Smith, M. S.; Marineau, E. C.

2016-05-01

The krypton tagging velocimetry (KTV) technique is applied to the turbulent boundary layer on the wall of the "Mach 3 Calibration Tunnel" at Arnold Engineering Development Complex (AEDC) White Oak. Profiles of velocity were measured with KTV and Pitot-pressure probes in the Mach 2.7 turbulent boundary layer comprised of 99 % {N}2/1 % Kr at momentum-thickness Reynolds numbers of {Re}_{\\varTheta }= 800, 1400, and 2400. Agreement between the KTV- and Pitot-derived velocity profiles is excellent. The KTV and Pitot velocity data follow the law of the wall in the logarithmic region with application of the Van Driest I transformation. The velocity data are analyzed in the outer region of the boundary layer with the law of the wake and a velocity-defect law. KTV-derived streamwise velocity fluctuation measurements are reported and are consistent with data from the literature. To enable near-wall measurement with KTV (y/δ ≈ 0.1-0.2), an 800-nm longpass filter was used to block the 760.2-nm read-laser pulse. With the longpass filter, the 819.0-nm emission from the re-excited Kr can be imaged to track the displacement of the metastable tracer without imaging the reflection and scatter from the read-laser off of solid surfaces. To operate the Mach 3 AEDC Calibration Tunnel at several discrete unit Reynolds numbers, a modification was required and is described herein.

High-Speed Schlieren Movies of Decelerators at Supersonic Speeds

NASA Technical Reports Server (NTRS)

1960-01-01

Tests were conducted on several types of porous parachutes, a paraglider, and a simulated retrorocket. Mach numbers ranged from 1.8-3.0, porosity from 20-80 percent, and camera speeds from 1680-3000 feet per second (fps) in trials with porous parachutes. Trials of reefed parachutes were conducted at Mach number 2.0 and reefing of 12-33 percent at camera speeds of 600 fps. A flexible parachute with an inflatable ring in the periphery of the canopy was tested at Reynolds number 750,000 per foot, Mach number 2.85, porosity of 28 percent, and camera speed of 36oo fps. A vortex-ring parachute was tested at Mach number 2.2 and camera speed of 3000 fps. The paraglider, with a sweepback of 45 degrees at an angle of attack of 45 degrees was tested at Mach number 2.65, drag coefficient of 0.200, and lift coefficient of 0.278 at a camera speed of 600 fps. A cold air jet exhausting upstream from the center of a bluff body was used to simulate a retrorocket. The free-stream Mach number was 2.0, free-stream dynamic pressure was 620 lb/sq ft, jet-exit static pressure ratio was 10.9, and camera speed was 600 fps.

Wind-tunnel measurements of the chordwise pressure distribution and profile drag of a research airplane model incorporating a 17-percent-thick supercritical wing

NASA Technical Reports Server (NTRS)

Ferris, J. C.

1973-01-01

The Langley 8-foot transonic pressure tunnel to determine the wing chordwise pressure distribution for a 0.09-scale model of a research airplane incorporating a 17-percent-thick supercritical wing. Airfoil profile drag was determined from wake pressure measurements at the 42-percent-semispan wing station. The investigation was conducted at Mach numbers from 0.30 to 0.80 over an angle-of-attack range sufficient to include buffet onset. The Reynolds number based on the mean geometric chord varied from 2 x 10 to the 6th power at Mach number 0.30 to 3.33 x 10 to the 6th power at Mach number 0.65 and was maintained at a constant value of 3.86 x 10 to the 6th power at Mach numbers from 0.70 to 0.80. Pressure coefficients for four wing semispan stations and wing-section normal-force and pitching-moment coefficients for two semispan stations are presented in tabular form over the Mach number range from 0.30 to 0.80. Plotted chordwise pressure distributions and wake profiles are given for a selected range of section normal-force coefficients over the same Mach number range.

INTERSTELLAR SONIC AND ALFVENIC MACH NUMBERS AND THE TSALLIS DISTRIBUTION

DOE Office of Scientific and Technical Information (OSTI.GOV)

Tofflemire, Benjamin M.; Burkhart, Blakesley; Lazarian, A.

2011-07-20

In an effort to characterize the Mach numbers of interstellar medium (ISM) magnetohydrodynamic (MHD) turbulence, we study the probability distribution functions (PDFs) of spatial increments of density, velocity, and magnetic field for 14 ideal isothermal MHD simulations at a resolution of 512{sup 3}. In particular, we fit the PDFs using the Tsallis function and study the dependency of the fit parameters on the compressibility and magnetization of the gas. We find that the Tsallis function fits PDFs of MHD turbulence well, with fit parameters showing sensitivities to the sonic and Alfven Mach numbers. For three-dimensional density, column density, and Position-Position-Velocitymore » data, we find that the amplitude and width of the PDFs show a dependency on the sonic Mach number. We also find that the width of the PDF is sensitive to the global Alfvenic Mach number especially in cases where the sonic number is high. These dependencies are also found for mock observational cases, where cloud-like boundary conditions, smoothing, and noise are introduced. The ability of Tsallis statistics to characterize the sonic and Alfvenic Mach numbers of simulated ISM turbulence points to it being a useful tool in the analysis of the observed ISM, especially when used simultaneously with other statistical techniques.« less

Natural laminar flow flight experiments on a swept wing business jet-boundary layer stability analyses

NASA Technical Reports Server (NTRS)

Rozendaal, R. A.

1986-01-01

The linear boundary layer stability analyses and their correlation with data of 18 cases from a natural laminar flow (NLF) flight test program using a Cessna Citation 3 business jet are described. The transition point varied from 5% to 35% chord for these conditions, and both upper and lower wing surfaces were included. Altitude varied from 10,000 to 43,000 ft and Mach number from 0.3 to 0.8. Four cases were at nonzero sideslip. Although there was much scatter in the results, the analyses of boundary layer stability at the 18 conditions led to the conclusion that crossflow instability was the primary cause of transition. However, the sideslip cases did show some interaction of crossflow and Tollmien-Schlichting disturbances. The lower surface showed much lower Tollmien-Schlichting amplification at transition than the upper surface, but similar crossflow amplifications. No relationship between Mach number and disturbance amplification at transition could be found. The quality of these results is open to question from questionable wing surface quality, inadequate density of transition sensors on the wing upper surface, and an unresolved pressure shift in the wing pressure data. The results of this study show the need for careful preparation for transition experiments. Preparation should include flow analyses of the test surface, boundary layer disturbance amplification analyses, and assurance of adequate surface quality in the test area. The placement of necessary instruments and usefulness of the resulting data could largely be determined during the pretest phase.

Pressure-distribution measurements on a transonic low-aspect ratio wing

NASA Technical Reports Server (NTRS)

Keener, E. R.

1985-01-01

Experimental surface pressure distributions and oil flow photographs are presented for a 0.90 m semispan model of NASA/Lockheed Wing C, a generic transonic, supercritical, low aspect ratio, highly 3-dimensional configuration. This wing was tested at the design angle of attack of 5 deg over a Mach number range from 0.25 to 0.96, and a Reynolds number range from 3.4 x 1,000,000 to 10 x 1,000,000. Pressures were measured with both the tunnel floor and ceiling suction slots open for most of the tests but taped closed for some tests to simulate solid walls. A comparison is made with the measured pressures from a small model in high Reynolds number facility and with predicted pressures using two three dimesional, transonic full potential flow wing codes: design code FLO22 (nonconservative) and TWING code (conservative). At the given design condition, a small region of flow separation occurred. At a Mach number of 0.82 the flow was unseparated and the surface flow angles were less than 10 deg, indicating that the boundary layer flow was not 3-D. Evidence indicate that wings that are optimized for mild shock waves and mild pressure recovery gradients generally have small 3-D boundary layer flow at design conditions for unseparated flow.

Experimental evaluation of nacelle-airframe interference forces and pressures at Mach numbers of 0.9 to 1.4

NASA Technical Reports Server (NTRS)

Bencze, D. P.

1977-01-01

Detailed interference force-and-pressure data were obtained on a representative supersonic transport wing-body-nacelle combination at Mach numbers of 0.9 to 1.4. The basic model consisted of a delta wing-body aerodynamic model with a length of 158.0 cm (62.2 in.) and a wingspan of 103.6 cm (40.8 in.) and four independently supported nacelles positioned beneath the model. The experimental program was conducted in the Ames 11- by 11-Foot Wind Tunnel at a constant unit Reynolds number. The primary variables examined included Mach number, angle of attack, nacelle position, and nacelle mass-flow ratio. Under the most favorable conditions, the net interference drag was equal to 50 percent the drag of four isolated nacelles at M = 1.4, 75 percent at M = 1.15, and 144 percent at M = 0.90. The overall interference effects were found to be rather constant over the operating angle-of-attack range of the configuration. The effects of mass-flow ratio on the interference pressure distributions were limited to the lip region of the nacelle and the local wing surface in the immediate vicinity of the nacelle lip. The net change in the measured interference forces resulting from variations in the nacelle mass-flow ratio were found to be quite small.

Splitting methods for low Mach number Euler and Navier-Stokes equations

NASA Technical Reports Server (NTRS)

Abarbanel, Saul; Dutt, Pravir; Gottlieb, David

1987-01-01

Examined are some splitting techniques for low Mach number Euler flows. Shortcomings of some of the proposed methods are pointed out and an explanation for their inadequacy suggested. A symmetric splitting for both the Euler and Navier-Stokes equations is then presented which removes the stiffness of these equations when the Mach number is small. The splitting is shown to be stable.

Tabulations of static pressure coefficients on the surfaces of 3 pylon-mounted axisymmetric flow-through nacelles at Mach numbers from 0.40 to 0.98

NASA Technical Reports Server (NTRS)

Re, R. J.; Peddrew, K. H.

1982-01-01

Three flow through nacelles mounted on an 82 deg swept pylon (10 percent thickness-to-chord ratio) were tested in the Langley 16 foot Transonic Tunnel. The long uncambered pylon was supported from a small body of revolution so that pressure measurements on the nacelle and pylon represent a pylon nacelle flow field without a wing present. Two nacelles had NACA 1-85-100 inlets and different circular arc afterbodies. The third nacelle had an NACA 1-70-100 inlet with a circular arc afterbody having the same external shape as one of the other nacelles. Nacelle length to maximum diameter ratio was 3.5. Data were obtained at angles of attack from 2 deg to 8 deg at selected Mach numbers.

Ascent heat transfer rate distribution on the North American Rockwell delta wing orbiter and the General Dynamics/Convair booster at a Mach number of 8 (mated)

NASA Technical Reports Server (NTRS)

Matthews, R. K.; Martindale, W. R.; Warmbrod, J. D.

1972-01-01

A wind tunnel test program to determine aerodynamic interference heating on the North American Rockwell orbiter mated with the General Dynamics Convair booster is discussed. The tests were conducted at the Arnold Engineering Development Center (AEDC) in Tunnel B of the von Karman Gas Dynamics Facility (VKF). The test period was June 1971. Heat-transfer rates were determined by the phase-change paint technique on 0.013-scale Stycast models using Tempilaq as the surface temperature indicator. The nominal test conditions were: Mach 8, free-stream unit length Reynolds numbers of 1.25 x one million and 2.55 x one million angles of attack of -5, 0, +5 deg. Model details, test conditions, phase-change paint photographs and reduced heat-transfer coefficients are presented.

Flight-measured X-24A lifting body control surface hinge moments and correlation with wind tunnel predictions

NASA Technical Reports Server (NTRS)

Tang, M. H.; Pearson, G. P. E.

1973-01-01

Control-surface hinge-moment measurements obtained in the X-24A lifting body flight-test program are compared with results from wind-tunnel tests. The effects of variations in angle of attack, angle of sideslip, rudder bias, rudder deflection, upper-flap deflection, lower-flap deflection, Mach number, and rocket-engine operation on the control-surface hinge moments are presented. In-flight motion pictures of tufts attached to the inboard side of the right fin and the rudder and upper-flap surfaces are discussed.

Diffusive wave in the low Mach limit for non-viscous and heat-conductive gas

NASA Astrophysics Data System (ADS)

Liu, Yechi

2018-06-01

The low Mach number limit for one-dimensional non-isentropic compressible Navier-Stokes system without viscosity is investigated, where the density and temperature have different asymptotic states at far fields. It is proved that the solution of the system converges to a nonlinear diffusion wave globally in time as Mach number goes to zero. It is remarked that the velocity of diffusion wave is proportional with the variation of temperature. Furthermore, it is shown that the solution of compressible Navier-Stokes system also has the same phenomenon when Mach number is suitably small.

Instability of Poiseuille flow at extreme Mach numbers: linear analysis and simulations.

PubMed

Xie, Zhimin; Girimaji, Sharath S

2014-04-01

We develop the perturbation equations to describe instability evolution in Poiseuille flow at the limit of very high Mach numbers. At this limit the equation governing the flow is the pressure-released Navier-Stokes equation. The ensuing semianalytical solution is compared against simulations performed using the gas-kinetic method (GKM), resulting in excellent agreement. A similar comparison between analytical and computational results of small perturbation growth is performed at the incompressible (zero Mach number) limit, again leading to excellent agreement. The study accomplishes two important goals: it (i) contrasts the small perturbation evolution in Poiseuille flows at extreme Mach numbers and (ii) provides important verification of the GKM simulation scheme.

Cold spray nozzle mach number limitation

NASA Astrophysics Data System (ADS)

Jodoin, B.

2002-12-01

The classic one-dimensional isentropic flow approach is used along with a two-dimensional axisymmetric numerical model to show that the exit Mach number of a cold spray nozzle should be limited due to two factors. To show this, the two-dimensional model is validated with experimental data. Although both models show that the stagnation temperature is an important limiting factor, the one-dimensional approach fails to show how important the shock-particle interactions are at limiting the nozzle Mach number. It is concluded that for an air nozzle spraying solid powder particles, the nozzle Mach number should be set between 1.5 and 3 to limit the negative effects of the high stagnation temperature and of the shock-particle interactions.

Effect of Rocket-Motor Operation on the Drag of Three 1/5-Scale Hermes A-3A Models in Free Flight

NASA Technical Reports Server (NTRS)

Jackson, H. Herbert

1954-01-01

Three 1/5-scale models of the Hermes A-3A missile have been flown to determine the effect of rocket-motor operation on the drag corresponding to various altitude and Mach number combinations. The flights covered a Mach number range from 0.5 to 1.8, and ratios of jet-exit static pressure to free-stream static pressure from 0.8 to 1.8. The results indicate that the power-on drag of the missile should be the same as the power-off drag at Mach number 1.3 and slightly less than the power-off drag at Mach number 1.55.

Evaluation of Flush-Mounted, S-Duct Inlets with Large Amounts of Boundary Layer Ingestion

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Morehouse, Melissa B.

2003-01-01

A new high Reynolds number test capability for boundary layer ingesting inlets has been developed for the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. Using this new capability, an experimental investigation of four S-duct inlet configurations with large amounts of boundary layer ingestion (nominal boundary layer thickness of about 40% of inlet height) was conducted at realistic operating conditions (high subsonic Mach numbers and full-scale Reynolds numbers). The objectives of this investigation were to 1) provide a database for CFD tool validation on boundary layer ingesting inlets operating at realistic conditions and 2) provide a baseline inlet for future inlet flow-control studies. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on duct exit diameter) from 5.1 million to a full-scale value of 13.9 million, and inlet mass-flow ratios from 0.39 to 1.58 depending on Mach number. Results of this investigation indicate that inlet pressure recovery generally decreased and inlet distortion generally increased with increasing Mach number. Except at low Mach numbers, increasing inlet mass-flow increased pressure recovery and increased distortion. Increasing the amount of boundary layer ingestion (by decreasing inlet throat height) or ingesting a boundary layer with a distorted (adverse) profile decreased pressure recovery and increased distortion. Finally, increasing Reynolds number had almost no effect on inlet distortion but increased inlet recovery by about one-half percent at a Mach number near cruise.

Flight-determined characteristics of an air intake system on an F-111A airplane

NASA Technical Reports Server (NTRS)

Hughes, D. L.; Johnson, H. J.

1972-01-01

Flow phenomena of the F-111A air intake system were investigated over a large range of Mach number, altitude, and angle of attack. Boundary-layer variations are shown for the fuselage splitter plate and inlet entrance stations. Inlet performance is shown in terms of pressure recovery, airflow, mass-flow ratio, turbulence factor, distortion factor, and power spectral density. The fuselage boundary layer was found to be not completely removed from the upper portion of the splitter plate at all Mach numbers investigated. Inlet boundary-layer ingestion started at approximately Mach 1.6 near the translating spike and cone. Pressure-recovery distribution at the compressor face showed increasing distortion with increasing angle of attack and increasing Mach number. The time-averaged distortion-factor value approached 1300, which is near the distortion tolerance of the engine at Mach numbers above 2.1.

Analysis of experimental results of the inlet for the NASA hypersonic research engine aerothermodynamic integration model. [wind tunnel tests of ramjet engine hypersonic inlets

NASA Technical Reports Server (NTRS)

Andrews, E. H., Jr.; Mackley, E. A.

1976-01-01

An aerodynamic engine inlet analysis was performed on the experimental results obtained at nominal Mach numbers of 5, 6, and 7 from the NASA Hypersonic Research Engine (HRE) Aerothermodynamic Integration Model (AIM). Incorporation on the AIM of the mixed-compression inlet design represented the final phase of an inlet development program of the HRE Project. The purpose of this analysis was to compare the AIM inlet experimental results with theoretical results. Experimental performance was based on measured surface pressures used in a one-dimensional force-momentum theorem. Results of the analysis indicate that surface static-pressure measurements agree reasonably well with theoretical predictions except in the regions where the theory predicts large pressure discontinuities. Experimental and theoretical results both based on the one-dimensional force-momentum theorem yielded inlet performance parameters as functions of Mach number that exhibited reasonable agreement. Previous predictions of inlet unstart that resulted from pressure disturbances created by fuel injection and combustion appeared to be pessimistic.

Body-surface pressure data on two monoplane-wing missile configurations with elliptical cross sections at Mach 2.50

NASA Technical Reports Server (NTRS)

Allen, J. M.; Hernandez, G.; Lamb, M.

1983-01-01

Tabulated body surface pressure data for two monoplane-wing missile configurations are presented and analyzed. Body pressure data are presented for body-alone, body-tail, and body-wing-tail combinations. For the lost combination, data are presented for tail-fin deflection angles of 0 deg and 30 deg to simulate pitch, yaw, and roll control for both configurations. The data cover angles of attack from -5 deg to 25 deg and angles of roll from 0 deg to 90 deg at a Mach number of 2.50 and a Reynolds number of 6.56 x 1,000,000 per meter. Very consistent, systematic trends with angle of attack and angle of roll were observed in the data, and very good symmetry was found at a roll angle of 0 deg. Body pressures depended strongly on the local body cross-section shape, with very little dependence on the upstream shape. Undeflected fins had only a small influence on the pressures on the aft end of the body; however, tail-fin deflections caused large changes in the pressures.

The effect of swirl recovery vanes on the cruise noise of an advanced propeller

NASA Technical Reports Server (NTRS)

Dittmar, James H.; Hall, David G.

1990-01-01

The SR-7A propeller was acoustically tested with and without downstream swirl recovery vanes to determine if any extra noise was caused by the interaction of the propeller wakes and vortices with these vanes. No additional noise was observed at the cruise condition over the angular range tested. The presence of the swirl recovery vanes did unload the propeller and some small peak noise reductions were observed from lower propeller loading noise. The propeller was also tested alone to investigate the behavior of the peak propeller noise with helical tip Mach number. As observed before on other propellers, the peak noise first rose with helical tip Mach number and then leveled off or decreased at higher helical tip Mach numbers. Detailed pressure-time histories indicate that a portion of the primary pressure pulse is progressively cancelled by a secondary pulse as the helical tip Mach number is increased. This cancellation appears to be responsible for the peak noise behavior at high helical tip Mach numbers.

Some anomalies observed in wind-tunnel tests of a blunt body at transonic and supersonic speeds

NASA Technical Reports Server (NTRS)

Brooks, J. D.

1976-01-01

An investigation of anomalies observed in wind tunnel force tests of a blunt body configuration was conducted at Mach numbers from 0.20 to 1.35 in the Langley 8-foot transonic pressure tunnel and at Mach numbers of 1.50, 1,80, and 2.16 in the Langley Unitary Plan wind tunnel. At a Mach number of 1.35, large variations occurred in axial force coefficient at a given angle of attack. At transonic and low supersonic speeds, the total drag measured in the wind tunnel was much lower than that measured during earlier ballistic range tests. Accurate measurements of total drag for blunt bodies will require the use of models smaller than those tested thus far; however, it appears that accurate forebody drag results can be obtained by using relatively large models. Shock standoff distance is presented from experimental data over the Mach number range from 1.05 to 4.34. Theory accurately predicts the shock standoff distance at Mach numbers up to 1.75.

Measurements in Flight of the Longitudinal-Stability Characteristics of a Republic YF-84A Airplane (Army Serial No. 45-59488) at High Subsonic Mach Numbers

NASA Technical Reports Server (NTRS)

Turner, Howard L.; Cooper, George E.

1948-01-01

A brief investigation was made of the longitudinal-stability characteristics of a YF-84A airplane (Army Serial No. 45-79488). The airplane developed a pitching-up tendency at approximately 0.80 Mach number which necessitated large push forces and down-elevator deflections for further increases in speed. In steady turns at 35,000 feet with the center of gravity at 28.3 percent mean aerodynamic chord for normal accelerations up to the maximum test value, the control-force gradients were excessive at Mach numbers over 0.78. Airplane buffeting did not present a serious problem in accelerated or unaccelerated flight at 15,000 and 35,000 feet up to the maximum test Mach number of 0.84. It is believed that excessive control force would be the limiting factor in attaining speeds in excess of 0.84 Mach number, especially at altitudes below 35,000 feet.

On solving the compressible Navier-Stokes equations for unsteady flows at very low Mach numbers

NASA Technical Reports Server (NTRS)

Pletcher, R. H.; Chen, K.-H.

1993-01-01

The properties of a preconditioned, coupled, strongly implicit finite difference scheme for solving the compressible Navier-Stokes equations in primitive variables are investigated for two unsteady flows at low speeds, namely the impulsively started driven cavity and the startup of pipe flow. For the shear-driven cavity flow, the computational effort was observed to be nearly independent of Mach number, especially at the low end of the range considered. This Mach number independence was also observed for steady pipe flow calculations; however, rather different conclusions were drawn for the unsteady calculations. In the pressure-driven pipe startup problem, the compressibility of the fluid began to significantly influence the physics of the flow development at quite low Mach numbers. The present scheme was observed to produce the expected characteristics of completely incompressible flow when the Mach number was set at very low values. Good agreement with incompressible results available in the literature was observed.

Pressure and heat-transfer distributions in a simulated wing-elevon cove with variable leakage at a free-stream Mach number of 6.9

NASA Technical Reports Server (NTRS)

Deveikis, W. D.; Bartlett, W.

1978-01-01

An experimental aerodynamic heating investigation was conducted to determine effects of hot boundary-layer ingestion into the cove on the windward surface between a wing and elevon for cove seal leak areas nominally between 0 and 100 percent of cove entrance area. Pressure and heating-rate distributions were obtained on the wing and elevon surfaces and on the cove walls of a full-scale model that represented a section of the cove region on the space shuttle orbiter. Data were obtained for both attached and separated turbulent boundary layers upstream of the unswept cove entrance. Average free-stream Mach number was 6.9, average free-stream unit Reynolds numbers were 1.31 x 10 to the 6th power and 4.40 x 10 to the 6th power per meter (0.40 x 10 to the 6th power and 1.34 x 10 to the 6th power per foot), and average total temperature was 1888 K (3400 R). Cove pressures and heating rates varied as a function of seal leak area independent of leak aspect ratio. Although cove heating rates for attached flow did not appear intolerable, it was postulated that convective heating in the cove may increase with time. For separated flow, the cove environment was considered too severe for unprotected interior structures of control surfaces.

Effect of Full-Chord Porosity on Aerodynamic Characteristics of the NACA 0012 Airfoil

NASA Technical Reports Server (NTRS)

Mineck, Raymond E.; Hartwich, Peter M.

1996-01-01

A test was conducted on a model of the NACA 0012 airfoil section with a solid upper surface or a porous upper surface with a cavity beneath for passive venting. The purposes of the test were to investigate the aerodynamic characteristics of an airfoil with full-chord porosity and to assess the ability of porosity to provide a multipoint or self-adaptive design. The tests were conducted in the Langley 8-Foot Transonic Pressure Tunnel over a Mach number range from 0.50 to 0.82 at chord Reynolds numbers of 2 x 10(exp 6), 4 x 10(exp 6), and 6 x 10(exp 6). The angle of attack was varied from -1 deg to 6 deg. At the lower Mach numbers, porosity leads to a dependence of the drag on the normal force. At subcritical conditions, porosity tends to flatten the pressure distribution, which reduces the suction peak near the leading edge and increases the suction over the middle of the chord. At supercritical conditions, the compression region on the porous upper surface is spread over a longer portion of the chord. In all cases, the pressure coefficient in the cavity beneath the porous surface is fairly constant with a very small increase over the rear portion. For the porous upper surface, the trailing edge pressure coefficients exhibit a creep at the lower section normal force coefficients, which suggests that the boundary layer on the rear portion of the airfoil is significantly thickening with increasing normal force coefficient.

Scramjet Isolator Modeling and Control

DTIC Science & Technology

2011-12-01

12 γ Ratio of specific heats . . . . . . . . . . . . . . . . . . . . 12 p1 Static pressure entering shock . . . . . . . . . . . . . . . . 12 M1 Mach...138 MAve Average stream Mach number . . . . . . . . . . . . . . . . 138 γ Ratio of specific heats ... heats , p1 is the static pressure entering the shock, and M1 is the Mach number of the flow entering the shock. Subsequent researchers [9] took a

Investigation of two pitot-static tubes at supersonic speeds

NASA Technical Reports Server (NTRS)

Hasel, Lowell E; Coletti, Donald E

1948-01-01

The results of tests at a Mach number of 1.94 of an ogives-nose cylindrical pitot-static tube and similar tests at Mach numbers of 1.93 and 1.62 of a service pitot-static tube to determine body static pressures and indicated Mach numbers are presented and discussed. The radial pressure distribution on the cylindrical bodies is compared with that calculated by an approximate theory.

Lateral Stability and Control Measurements of a 0.0858-Scale Model of the Lockheed XF-104 Airplane at Transonic Speeds

NASA Technical Reports Server (NTRS)

Arabian, Donald D.; Schmeer, James W.

1955-01-01

An investigation of the lateral stability and control effectiveness of a 0.0858-scale model of the Lockheed XF-104 airplane has been conducted in the Langley 16-foot transonic tunnel. The model has a low aspect ratio, 3.4-percent-thick wing with negative dihedral. The horizontal tail is located on top of the vertical tail. The investigation was made through a Mach number range of 0.80 to 1.06 at sideslip angles of -5 deg. to 5 deg. and angles of attack from 0 deg. to 16 deg. The control effectiveness of the aileron, rudder, and yaw damper were determined through the Mach number and angle-of-attack range. The results of the investigation indicated that the directional stability derivative was stable and that positive effective dihedral existed throughout the lift-coefficient range and Mach number range tested. The total aileron effectiveness, which in general produced favorable yaw with rolling moment, remained fairly constant for lift coefficients up to about 0.8 for the Mach number range tested. Yawing-moment effectiveness of the rudder changed little through the Mach number range. However, the yaw damper effectiveness decreased about 30 percent at the intermediate test Mach numbers.

Determining the standoff distance of the bow shock: Mach number dependence and use of models

NASA Technical Reports Server (NTRS)

Farris, M. H.; Russell, C. T.

1994-01-01

We explore the factors that determine the bow shock standoff distance. These factors include the parameters of the solar wind, as well as the size and shape of the obstacle. In this report we develop a semiempirical Mach number relation for the bow shock standoff distance in order to take into account the shock's behavior at low Mach numbers. This is done by determining which properties of the shock are most important in controlling the standoff distance and using this knowledge to modify the current Mach number relation. While the present relation has proven useful at higher Mach numbers, it has lacked effectiveness at the low Mach number limit. We also analyze the bow shock dependence upon the size and shape of the obstacle, noting that it is most appropriate to compare the standoff distance of the bow shock to the radius of curvature of the obstacle, as opposed to the distance from the focus of the object to the nose. Last, we focus our attention on the use of bow shock models in determining the standoff distance. We note that the physical behavior of the shock must correctly be taken into account, specifically the behavior as a function of solar wind dynamic pressure; otherwise, erroneous results can be obtained for the bow shock standoff distance.

Surface-Pressure and Flow-Visualization Data at Mach Number of 1.60 for Three 65 deg Delta Wings Varying in Leading-Edge Radius and Camber

NASA Technical Reports Server (NTRS)

McMillin, S. Naomi; Bryd, James E.; Parmar, Devendra S.; Bezos-OConnor, Gaudy M.; Forrest, Dana K.; Bowen, Susan

1996-01-01

An experimental investigation of the effect of leading-edge radius, camber, Reynolds number, and boundary-layer state on the incipient separation of a delta wing at supersonic speeds was conducted at the Langley Unitary Plan Wind Tunnel at Mach number of 1.60 over a free-stream Reynolds number range of 1 x 106 to 5 x 106 ft-1. The three delta wing models examined had a 65 deg swept leading edge and varied in cross-sectional shape: a sharp wedge, a 20:1 ellipse, and a 20:1 ellipse with a -9.750 circular camber imposed across the span. The wings were tested with and without transition grit applied. Surface-pressure coefficient data and flow-visualization data indicated that by rounding the wing leading edge or cambering the wing in the spanwise direction, the onset of leading-edge separation on a delta wing can be raised to a higher angle of attack than that observed on a sharp-edged delta wing. The data also showed that the onset of leading-edge separation can be raised to a higher angle of attack by forcing boundary-layer transition to occur closer to the wing leading edge by the application of grit or the increase in free-stream Reynolds number.

Longitudinal Stability and Control Characteristics of a Semispan Model of the XF7U-1 Tailless Airplane at Transonic Speeds by the NACA Wing-Flow Method, TED No. NACA DE307

NASA Technical Reports Server (NTRS)

Sawyer, Richard H.; Trant, James P., Jr.

1947-01-01

An investigation was made by the NACA wing-flow method to determine the longitudinal stability and control characteristics at transonic speeds of a semispan model of the XF7U-1 tailless airplane. The 25-percent chord line of the wing of the model was swept back 35 deg. The airfoil sections of the wing perpendicular to the 25-percent chord line were 12 percent thick. Measurements were made of the normal force and pitching moment through an angle-of-attack range from about -3 deg to 14 deg for several ailavator deflections at Mach numbers from 0.65 to about 1.08. The results of the tests indicated no adverse effects of compressibility up to a Mach number of at least 0.85 at low normal-force coefficients and small ailavator deflections. Up to a Mach number of 0.85, the neutral point at low normal-force coefficients was at about 25 percent of the mean aerodynamic chord and moved rearward irregularly to 41 or 42 percent with a further increase in Mach number to about 1.05. For deflections up to -8.0 percent, the ailavator was effective in changing the pitching moment except at Mach numbers from 0.93 to about 1.0 where ineffectiveness or reversal was indicated for deflections and normal-force coefficients. With -13.2 deg deflection at normal-force coefficients above about 0.3, reversal of ailavator effectiveness occurred at Mach numbers as low as 0.81. A nose-down trim change, which began at a Mach number of about 0.85, together with the loss in effectiveness of the ailavator, indicated that with increase in the Mach number from about 0.95 to 1.05 an abrupt ailavator movement of 5 deg or 6 deg first up and then down would be required to maintain level flight.

Longitudinal Stability and Control Characteristics from a Flight Investigation of a Cruciform Canard Missile Configuration Having an Exposed Wing-canard Area Ratio of 16:1

NASA Technical Reports Server (NTRS)

Moul, Martin T; Wineman, Andrew R

1952-01-01

A flight investigation has been made to determine the longitudinal stability and control characteristics of a 60 0 delta-wing-canard missile configuration with an exposed wing-canard area ratio of 16:1. The results presented include the longitudinal stability derivatives, control effectiveness, and drag characteristics for a Mach number range of 0.75 to 1.80 and are compared with the results of a similar configuration having larger 6ontrols. Stability characteristics are also presented from the flights of an interdigitated canard configuration at a Mach number of 2.08 and a wing-body configuration at Mach numbers of 1.25 to 1.45. The stability derivatives varied gradually with Mach number with the exception of the damping-in-pitch derivative. Aerodynamic damping in pitch decreased to a minimum at a Mach number of 1.0 3, then increased to a peak value at a Mach number of 1.26 followed by a gradual decrease at higher Mach numbers. The aerodynamic-center location of the in-line canard configuration shifted rearward 13 percent of the mean aerodynamic chord at transonic speeds. The pitching-moment curve slope was 25 percent greater for the model having no canards than for the in-line configuration. No large effects of interdigitation were noted in the stability derivatives. Pitching effectiveness of the in-line configuration was maintained throughout the Mach number range. A comparison of the stability and control characteristics of two canard configurations having different area controls showed that decreasing the control area 44 percent decreased the pitching effectiveness proportionally, shifted the aerodynamic-center location rearward 9 to 14 percent of the mean aerodynamic chord, and reduced the total hinge moments required for 10 trimmed flight about 50 percent at transonic speeds.

Experimental aerodynamic characteristics for a cylindrical body of revolution with various noses at angles of attack from 0 deg to 58 deg and Mach numbers from 0.6 to 2.0

NASA Technical Reports Server (NTRS)

Jorgensen, L. H.; Nelson, E. R.

1974-01-01

An experimental investigation was conducted to determine the effect of forebody geometry, a grit ring around the nose, Reynolds number, Mach number, and angle of attack on the aerodynamic characteristics of a body of revolution. Aerodynamic force and moment characteristics were measured for a cylindrical body with tangent ogive noses of fineness ratio 2.5, 3.0, 3.5, and 5.0. The cylindrical body was tested with an ogive nose having a rounded tip and an ogive nose with two different nose strake arrangements. Aerodynamic configurations were tested at various Mach numbers, angles of attack, and Reynolds numbers. The data demonstrate that the aerodynamic characteristics for a body of revolution can be significantly affected by changes in nose fineness ratio, nose bluntness, Reynolds number, Mach number, and, of course, angle of attack. Nose strakes increased the normal forces but had little effect on the side forces that developed at subsonic Mach numbers for alpha greater than about 25. A grit ring around the nose had little or no effect on the aerodynamic characteristics.

Skin Friction Measurements at a Mach Number of Three and Momentum Thickness Reynolds Numbers Up to a Half Million.

DTIC Science & Technology

1980-09-01

k ADAOGZ 826 ~~~~~~~~~AIR FORCE FLIGHT DNMC A RGTPTESNABO / / I SKIN FRICTION MEASUREMENTS AT A MACH NUMBER OF THREE AND MOMENT--ETCIU) UNCASSFIE...AT A MACH NUMBER OF THREE AND MOMENTUM THICKNESS REYNOLDS NUMEERS UP TO A HALF MILLION Anthony W. Fiore Aeromechanics Division DTIC September 1980...NOR(&) S. CONTRACT OR GRANT NUMBER (s) 9. PER ORMING ORGANIZATION NAME AND ADDRESS PROR UNT N WU Flight Dynamics Laboratory (AFWAL/FIMG)RA; WOKUNTU

Experimental investigation of a 10-percent-thick helicopter rotor airfoil section designed with a viscous transonic analysis code

NASA Technical Reports Server (NTRS)

Noonan, K. W.

1981-01-01

An investigation was conducted in the Langley 6- by 28-Inch Transonic Tunnel to determine the two dimensional aerodynamic characteristics of a 10-percent-thick helicopter rotor airfoil at Mach numbers from 0.33 to 0.87 and respective Reynolds numbers from 4.9 x 10 to the 6th to 9.8 x 10 to the 6th. This airfoil, designated the RC-10(N)-1, was also investigated at Reynolds numbers from 3.0 x 10 to the 6th to 7.3 x 10 to the 6th at respective Mach numbers of 0.33 to 0.83 for comparison wit the SC 1095 (with tab) airfoil. The RC-10(N)-1 airfoil was designed by the use of a viscous transonic analysis code. The results of the investigation indicate that the RC-10(N)-1 airfoil met all the design goals. At a Reynolds number of about 9.4 x 10 to the 6th the drag divergence Mach number at zero normal-force coefficient was 0.815 with a corresponding pitching-moment coefficient of zero. The drag divergence Mach number at a normal-force coefficient of 0.9 and a Reynolds number of about 8.0 x 10 to the 6th was 0.61. The drag divergence Mach number of this new airfoil was higher than that of the SC 1095 airfoil at normal-force coefficients above 0.3. Measurements in the same wind tunnel at comparable Reynolds numbers indicated that the maximum normal-force coefficient of the RC-10(N)-1 airfoil was higher than that of the NACA 0012 airfoil for Mach numbers above about 0.35 and was about the same as that of the SC 1095 airfoil for Mach numbers up to 0.5.

Aerodynamic Loading Characteristics at Mach Numbers from 0.80 to 1.20 of a 1/10-Scale Three-Stage Scout Model

NASA Technical Reports Server (NTRS)

Kelly, Thomas C.

1961-01-01

Aerodynamic loads results have been obtained in the Langley 8-foot transonic pressure tunnel at Mach numbers from 0.80 to 1.20 for a 1/10-scale model of the upper three stages of the Scout vehicle. Tests were conducted through an angle-of-attack range from -8 deg to 8 deg at an average test Reynolds number per foot of about 4.0 x 10(exp 6). Results indicated that the peak negative pressures associated with expansion corners at the nose and transition flare exhibit sizeable variations which occur over a relatively small Mach number range. The magnitude of the variations may cause the critical local loading condition for the full-scale vehicle to occur at a Mach number considerably lower than that at which the maximum dynamic pressure occurs in flight. The addition of protuberances simulating antennas and wiring conduits had slight, localized effects. The lift carryover from the nose and transition flare on the cylindrical portions of the model generally increased with an increase in Mach number.

Evaluation of Flush-Mounted, S-Duct Inlets With Large Amounts of Boundary Layer Ingestion

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Morehouse, Melissa B.

2003-01-01

A new high Reynolds number test capability for boundary layer ingesting inlets has been developed for the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. Using this new capability, an experimental investigation of four S-duct inlet configurations with large amounts of boundary layer ingestion (nominal boundary layer thickness of about 40% of inlet height) was conducted at realistic operating conditions (high subsonic Mach numbers and full-scale Reynolds numbers). The objectives of this investigation were to 1) develop a new high Reynolds number, boundary-layer ingesting inlet test capability, 2) evaluate the performance of several boundary layer ingesting S-duct inlets, 3) provide a database for CFD tool validation, and 4) provide a baseline inlet for future inlet flow-control studies. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on duct exit diameter) from 5.1 million to a fullscale value of 13.9 million, and inlet mass-flow ratios from 0.39 to 1.58 depending on Mach number. Results of this investigation indicate that inlet pressure recovery generally decreased and inlet distortion generally increased with increasing Mach number. Except at low Mach numbers, increasing inlet mass-flow increased pressure recovery and increased distortion. Increasing the amount of boundary layer ingestion (by decreasing inlet throat height and increasing inlet throat width) or ingesting a boundary layer with a distorted profile decreased pressure recovery and increased distortion. Finally, increasing Reynolds number had almost no effect on inlet distortion but increased inlet recovery by about one-half percent at a Mach number near cruise.

Numerical simulation of divergent rocket-based-combined-cycle performances under the flight condition of Mach 3

NASA Astrophysics Data System (ADS)

Cui, Peng; Xu, WanWu; Li, Qinglian

2018-01-01

Currently, the upper operating limit of the turbine engine is Mach 2+, and the lower limit of the dual-mode scramjet is Mach 4. Therefore no single power systems can operate within the range between Mach 2 + and Mach 4. By using ejector rockets, Rocket-based-combined-cycle can work well in the above scope. As the key component of Rocket-based-combined-cycle, the ejector rocket has significant influence on Rocket-based-combined-cycle performance. Research on the influence of rocket parameters on Rocket-based-combined-cycle in the speed range of Mach 2 + to Mach 4 is scarce. In the present study, influences of Mach number and total pressure of the ejector rocket on Rocket-based-combined-cycle were analyzed numerically. Due to the significant effects of the flight conditions and the Rocket-based-combined-cycle configuration on Rocket-based-combined-cycle performances, flight altitude, flight Mach number, and divergence ratio were also considered. The simulation results indicate that matching lower altitude with higher flight Mach numbers can increase Rocket-based-combined-cycle thrust. For another thing, with an increase of the divergent ratio, the effect of the divergent configuration will strengthen and there is a limit on the divergent ratio. When the divergent ratio is greater than the limit, the effect of divergent configuration will gradually exceed that of combustion on supersonic flows. Further increases in the divergent ratio will decrease Rocket-based-combined-cycle thrust.

Columbia: The first five flights entry heating data series. Volume 4: The lower windward wing 50 percent and 80 percent semispans

NASA Technical Reports Server (NTRS)

Williams, S. D.

1983-01-01

Entry heating flight data and wind tunnel data on the lower wing 50% and 80% Semi-Spans are presented for the first five flights of the Space Shuttle Orbiter. The heating rate data is presented in terms of normalized film heat transfer coefficients as a function of angle-of-attack, Mach number, and Normal Shock Reynolds number. The surface heating rates and temperatures were obtained via the JSC NONLIN/INVERSE computer program. Time history plots of the surface heating rates and temperatures are also presented.

Forced Boundary-Layer Transition on X-43 (Hyper-X) in NASA LaRC 20-Inch Mach 6 Air Tunnel

NASA Technical Reports Server (NTRS)

Berry, Scott A.; DiFulvio, Michael; Kowalkowski, Matthew K.

2000-01-01

Aeroheating and boundary layer transition characteristics for the X-43 (Hyper-X) configuration have been experimentally examined in the Langley 20-Inch Mach 6 Air Tunnel. Global surface heat transfer distributions, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. Parametric variations include angles-of-attack of 0-deg, 2-deg, and 4-deg; Reynolds numbers based on model length of 1.2 to 15.4 million; and inlet cowl door both open and closed. The effects of discrete roughness elements on the forebody boundary layer, which included variations in trip configuration and height, were investigated. This document is intended to serve as a release of preliminary data to the Hyper-X program; analysis is limited to observations of the experimental trends in order to expedite dissemination.

Forced Boundary-Layer Transition on X-43 (Hyper-X) in NASA LaRC 31-Inch Mach 10 Air Tunnel

NASA Technical Reports Server (NTRS)

Berry, Scott A.; DiFulvio, Michael; Kowalkowski, Matthew K.

2000-01-01

Aeroheating and boundary layer transition characteristics for the X-43 (Hyper-X) configuration have been experimentally examined in the Langley 31-Inch Mach 10 Air Tunnel. Global surface heat transfer distributions, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. Parametric variations include angles-of-attack of 0-deg, 2-deg, 3-deg, and 4-deg; Reynolds numbers based on model length of 1.2 to 5.1 million; and inlet cowl door both open and closed. The effects of discrete roughness elements on the forebody boundary layer, which included variations in trip configuration and height, were investigated. This document is intended to serve as a release of preliminary data to the Hyper-X program; analysis is limited to observations of the experimental trends in order to expedite dissemination.

Aeroheating Testing and Predictions for Project Orion CEV at Turbulent Conditions

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Berger, Karen T.; Horvath, Thomas J.; Coblish, Joseph J.; Norris, Joseph D.; Lillard, Randolph P.; Kirk, Benjamin S.

2009-01-01

An investigation of the aeroheating environment of the Project Orion Crew Exploration Vehicle was performed in the Arnold Engineering Development Center Hypervelocity Wind Tunnel No. 9 Mach 8 and Mach 10 nozzles and in the NASA Langley Research Center 20 - Inch Mach 6 Air Tunnel. Heating data were obtained using a thermocouple-instrumented approx.0.035-scale model (0.1778-m/7-inch diameter) of the flight vehicle. Runs were performed in the Tunnel 9 Mach 10 nozzle at free stream unit Reynolds numbers of 1x10(exp 6)/ft to 20x10(exp 6)/ft, in the Tunnel 9 Mach 8 nozzle at free stream unit Reynolds numbers of 8 x 10(exp 6)/ft to 48x10(exp 6)/ft, and in the 20-Inch Mach 6 Air Tunnel at free stream unit Reynolds numbers of 1x10(exp 6)/ft to 7x10(exp 6)/ft. In both facilities, enthalpy levels were low and the test gas (N2 in Tunnel 9 and air in the 20-Inch Mach 6) behaved as a perfect-gas. These test conditions produced laminar, transitional and turbulent data in the Tunnel 9 Mach 10 nozzle, transitional and turbulent data in the Tunnel 9 Mach 8 nozzle, and laminar and transitional data in the 20- Inch Mach 6 Air Tunnel. Laminar and turbulent predictions were generated for all wind tunnel test conditions and comparisons were performed with the experimental data to help define the accuracy of computational method. In general, it was found that both laminar data and predictions, and turbulent data and predictions, agreed to within less than the estimated 12% experimental uncertainty estimate. Laminar heating distributions from all three data sets were shown to correlate well and demonstrated Reynolds numbers independence when expressed in terms of the Stanton number based on adiabatic wall-recovery enthalpy. Transition onset locations on the leeside centerline were determined from the data and correlated in terms of boundary-layer parameters. Finally turbulent heating augmentation ratios were determined for several body-point locations and correlated in terms of the boundary-layer momentum Reynolds number.

High-Order Multioperator Compact Schemes for Numerical Simulation of Unsteady Subsonic Airfoil Flow

NASA Astrophysics Data System (ADS)

Savel'ev, A. D.

2018-02-01

On the basis of high-order schemes, the viscous gas flow over the NACA2212 airfoil is numerically simulated at a free-stream Mach number of 0.3 and Reynolds numbers ranging from 103 to 107. Flow regimes sequentially varying due to variations in the free-stream viscosity are considered. Vortex structures developing on the airfoil surface are investigated, and a physical interpretation of this phenomenon is given.

Mach-Number Measurement with Laser and Pressure Probes in Humid Supersonic Flow

NASA Technical Reports Server (NTRS)

Herring, G. C.

2008-01-01

Mach-number measurements using a nonintrusive optical technique, laser-induced thermal acoustics (LITA), are compared to pressure probes in humid supersonic airflow. The two techniques agree well in dry flow (-35 C dew point), but LITA measurements show about five times larger fractional change in Mach number than that of the pressure-probe when water is purposefully introduced into the flow. Possible reasons for this discrepancy are discussed.

High-Speed Wind-Tunnel Investigation of the Longitudinal Stability and Control Characteristics of a 0.10-Scale Model of the Grumman XF9F-2 Airplane, TED No. NACA DE301

NASA Technical Reports Server (NTRS)

Polhamus, Edward C.; King, Thomas J., Jr.

1948-01-01

An investigation was made in the Langley high-speed 7-by 10-foot tunnel to determine the high-speed longitudinal stability end con&o1 characteristics of a 0.01-scale model of the Grumman XF9F-2 airplane in the Mach number range from 0.40 to 0.85. The results indicated that the lift and drag force breaks occurred at a Mach number of about 0.76. The aerodynamic-center position moved rearward after the force break and control position stability was present for all Mach numbers up to a Mach number of 0.80.

Transonic Aerodynamic Characteristics of a Model of a Proposed Six-Engine Hull-Type Seaplane Designed for Supersonic Flight

NASA Technical Reports Server (NTRS)

Wornom, Dewey E.

1960-01-01

Force tests of a model of a proposed six-engine hull-type seaplane were performed in the Langley 8-foot transonic pressure tunnel. The results of these tests have indicated that the model had a subsonic zero-lift drag coefficient of 0.0240 with the highest zero-lift drag coefficient slightly greater than twice the subsonic drag level. Pitchup tendencies were noted for subsonic Mach numbers at relatively high lift coefficients. Wing leading-edge droop increased the maximum lift-drag ratio approximately 8 percent at a Mach number of 0.80 but this effect was negligible at a Mach number of 0.90 and above. The configuration exhibited stable lateral characteristics over the test Mach number range.

Recent National Transonic Facility Test Process Improvements (Invited)

NASA Technical Reports Server (NTRS)

Kilgore, W. A.; Balakrishna, S.; Bobbitt, C. W., Jr.; Adcock, J. B.

2001-01-01

This paper describes the results of two recent process improvements; drag feed-forward Mach number control and simultaneous force/moment and pressure testing, at the National Transonic Facility. These improvements have reduced the duration and cost of testing. The drag feed-forward Mach number control reduces the Mach number settling time by using measured model drag in the Mach number control algorithm. Simultaneous force/moment and pressure testing allows simultaneous collection of force/moment and pressure data without sacrificing data quality thereby reducing the overall testing time. Both improvements can be implemented at any wind tunnel. Additionally the NTF is working to develop and implement continuous pitch as a testing option as an additional method to reduce costs and maintain data quality.

Recent National Transonic Facility Test Process Improvements (Invited)

NASA Technical Reports Server (NTRS)

Kilgore, W. A.; Balakrishna, S.; Bobbitt, C. W., Jr.; Adcock, J. B.

2001-01-01

This paper describes the results of two recent process improvements; drag feed-forward Mach number control and simultaneous force/moment and pressure testing, at the National Transonic Facility. These improvements have reduced the duration and cost of testing. The drag feedforward Mach number control reduces the Mach number settling time by using measured model drag in the Mach number control algorithm. Simultaneous force/moment and pressure testing allows simultaneous collection of force/moment and pressure data without sacrificing data quality thereby reducing the overall testing time. Both improvements can be implemented at any wind tunnel. Additionally the NTF is working to develop and implement continuous pitch as a testing option as an additional method to reduce costs and maintain data quality.

Sweep and Compressibility Effects on Active Separation Control at High Reynolds Numbers

NASA Technical Reports Server (NTRS)

Seifert, Avi; Pack, LaTunia G.

2000-01-01

This paper explores the effects of compressibility, sweep and excitation location on active separation control at high Reynolds numbers. The model, which was tested in a cryogenic pressurized wind tunnel, simulates the upper surface of a 20% thick GlauertGoldschmied type airfoil at zero angle of attack. The flow is fully turbulent since the tunnel sidewall boundary layer flows over the model. Without control, the flow separates at the highly convex area and a large turbulent separation bubble is formed. Periodic excitation is applied to gradually eliminate the separation bubble. Two alternative blowing slot locations as well as the effect of compressibility, sweep and steady suction or blowing were studied. During the test the Reynolds numbers ranged from 2 to 40 million and Mach numbers ranged from 0.2 to 0.7. Sweep angles were 0 and 30 deg. It was found that excitation must be introduced slightly upstream of the separation region regardless of the sweep angle at low Mach number. Introduction of excitation upstream of the shock wave is more effective than at its foot. Compressibility reduces the ability of steady mass transfer and periodic excitation to control the separation bubble but excitation has an effect on the integral parameters, which is similar to that observed in low Mach numbers. The conventional swept flow scaling is valid for fully and even partially attached flow, but different scaling is required for the separated 3D flow. The effectiveness of the active control is not reduced by sweep. Detailed flow field dynamics are described in the accompanying paper.

Sweep and Compressibility Effects on Active Separation Control at High Reynolds Numbers

NASA Technical Reports Server (NTRS)

Seifert, Avi; Pack, LaTunia G.

2000-01-01

This paper explores the effects of compressibility, sweep and excitation location on active separation control at high Reynolds numbers. The model, which was tested in a cryogenic pressurized wind tunnel, simulates the upper surface of a 20% thick Glauert Goldschmied type airfoil at zero angle of attack. The flow is fully turbulent since the tunnel sidewall boundary layer flows over the model. Without control, the flow separates at the highly convex area and a large turbulent separation bubble is formed. Periodic excitation is applied to gradually eliminate the separation bubble. Two alternative blowing slot locations as well as the effect of compressibility, sweep and steady suction or blowing were studied. During the test the Reynolds numbers ranged from 2 to 40 million and Mach numbers ranged from 0.2 to 0.7. Sweep angles were 0 and 30 deg. It was found that excitation must be introduced slightly upstream of the separation region regardless of the sweep angle at low Mach number. Introduction of excitation upstream of the shock wave is more effective than at its foot. Compressibility reduces the ability of steady mass transfer and periodic excitation to control the separation bubble but excitation has an effect on the integral parameters, which is similar to that observed in low Mach numbers. The conventional swept flow scaling is valid for fully and even partially attached flow, but different scaling is required for the separated 3D flow. The effectiveness of the active control is not reduced by sweep. Detailed flow field dynamics are described in the accompanying paper.

DSMC simulations of Mach 20 nitrogen flows about a 70 degree blunted cone and its wake

NASA Technical Reports Server (NTRS)

Moss, James N.; Dogra, Virendra K.; Wilmoth, Richard G.

1993-01-01

Numerical results obtained with the direct simulation Monte Carlo (DSMC) method are presented for Mach 20 nitrogen flow about a 70-deg blunted cone. The flow conditions simulated are those that can be obtained in existing low-density hypersonic wind tunnels. Three sets of flow conditions are simulated with freestream Knudsen numbers ranging from 0.03 to 0.001. The focus is to characterize the wake flow under rarefied conditions. This is accomplished by calculating the influence of rarefaction on wake structure along with the impact that an afterbody has on flow features. This data report presents extensive information concerning flowfield features and surface quantities.

Subsonic Aerodynamic Assessment of Vortex Flow Management Devices on a High-Speed Civil Transport Configuration

NASA Technical Reports Server (NTRS)

Campbell, Bryan A.; Applin, Zachary T.; Kemmerly, Guy T.

1999-01-01

An experimental investigation of the effects of leading-edge vortex management devices on the subsonic performance of a high-speed civil transport (HSCT) configuration was conducted in the Langley 14- by 22-Foot Subsonic Tunnel. Data were obtained over a Mach number range of 0.14 to 0.27, with corresponding chord Reynolds numbers of 3.08 x 10 (sup 6) to 5.47 x 10 (sup 6). The test model was designed for a cruise Mach number of 2.7. During the subsonic high-lift phase of flight, vortical flow dominates the upper surface flow structure, and during vortex breakdown, this flow causes adverse pitch-up and a reduction of usable lift. The experimental results showed that the beneficial effects of small leading-edge vortex management devices located near the model reference center were insufficient to substantially affect the resulting aerodynamic forces and moments. However, devices located at or near the wiring apex region demonstrated potential for pitch control with little effect on overall lift.

Endwall Heat Transfer Measurements in a Transonic Turbine Cascade

NASA Technical Reports Server (NTRS)

Giel, P. W.; Thurman, D. R.; VanFossen, G. J.; Hippensteele, S. A.; Boyle, R. J.

1996-01-01

Turbine blade endwall heat transfer measurements are given for a range of Reynolds and Mach numbers. Data were obtained for Reynolds numbers based on inlet conditions of 0.5 and 1.0 x 106, for isentropic exit Mach numbers of 1.0 and 1.3, and for freestream turbulence intensities of 0.25% and 7.0%. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136' of turning and an axial chord of 12.7 cm. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for computational fluid dynamics (CFD) code and model verification. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet. Endwall heat transfer data were obtained using a steady-state liquid crystal technique.

Surface roughness effects on the subsonic aerodynamics of the Rockwell International 089B-139B orbiter

NASA Technical Reports Server (NTRS)

Ware, G. M.; Spencer, B., Jr.

1973-01-01

An experimental test program was conducted to determine the effects of vehicle surface roughness on the subsonic aerodynamic characteristics of a 0.01875 scale model of a Rockwell International Space Shuttle Configuration. Surface roughness was simulated by applying a sparce coating of carborundum grit to the complete model. Various grit sizes were investigated. Tests were conducted in the Langley Low Turbulence Pressure Tunnel at a constant nominal Mach number of 0.25 with Reynolds number varying from 2 to 12 x 10 to the 6th power per foot. Angle of attack was varied from about -2 to 28 deg at 0 deg and 6 deg angle of sideslip.

Two-dimensional aerodynamic characteristics of several rotorcraft airfoils at Mach numbers from 0.35 to 0.90

NASA Technical Reports Server (NTRS)

Noonan, K. W.; Bingham, G. J.

1977-01-01

An investigation was conducted in the Langley 6- by 28-inch transonic tunnel and the 6- by 19-inch transonic tunnel to determine the two-dimensional aerodynamic characteristics of several rotorcraft airfoils at Mach numbers from 0.35 to 0.90. The airfoils differed in thickness, thickness distribution, and camber. The FX69-H-098, the BHC-540, and the NACA 0012 airfoils were investigated in the 6- by 28-inch tunnel at Reynolds numbers (based on chord) from about 4.7 to 9.3 million at the lowest and highest test Mach numbers respectively. The FX69-H-098, the NLR-1, the BHC-540, and the NACA 23012 airfoils were investigated in the 6- by 19-inch tunnel at Reynolds numbers from about 0.9 to 2.2 million at the lowest and highest test Mach numbers respectively.

Airfoil shape for flight at subsonic speeds. [design analysis and aerodynamic characteristics of the GAW-1 airfoil

NASA Technical Reports Server (NTRS)

Whitcomb, R. T. (Inventor)

1976-01-01

An airfoil is examined that has an upper surface shaped to control flow accelerations and pressure distribution over the upper surface and to prevent separation of the boundary layer due to shock wave formulation at high subsonic speeds well above the critical Mach number. A highly cambered trailing edge section improves overall airfoil lifting efficiency. Diagrams illustrating supersonic flow and shock waves over the airfoil are shown.

Flow analysis for the nacelle of an advanced ducted propeller at high angle-of-attack and at cruise with boundary layer control

NASA Technical Reports Server (NTRS)

Hwang, D. P.; Boldman, D. R.; Hughes, C. E.

1994-01-01

An axisymmetric panel code and a three dimensional Navier-Stokes code (used as an inviscid Euler code) were verified for low speed, high angle of attack flow conditions. A three dimensional Navier-Stokes code (used as an inviscid code), and an axisymmetric Navier-Stokes code (used as both viscous and inviscid code) were also assessed for high Mach number cruise conditions. The boundary layer calculations were made by using the results from the panel code or Euler calculation. The panel method can predict the internal surface pressure distributions very well if no shock exists. However, only Euler and Navier-Stokes calculations can provide a good prediction of the surface static pressure distribution including the pressure rise across the shock. Because of the high CPU time required for a three dimensional Navier-Stokes calculation, only the axisymmetric Navier-Stokes calculation was considered at cruise conditions. The use of suction and tangential blowing boundary layer control to eliminate the flow separation on the internal surface was demonstrated for low free stream Mach number and high angle of attack cases. The calculation also shows that transition from laminar flow to turbulent flow on the external cowl surface can be delayed by using suction boundary layer control at cruise flow conditions. The results were compared with experimental data where possible.

Supersonic, nonlinear, attached-flow wing design for high lift with experimental validation

NASA Technical Reports Server (NTRS)

Pittman, J. L.; Miller, D. S.; Mason, W. H.

1984-01-01

Results of the experimental validation are presented for the three dimensional cambered wing which was designed to achieve attached supercritical cross flow for lifting conditions typical of supersonic maneuver. The design point was a lift coefficient of 0.4 at Mach 1.62 and 12 deg angle of attack. Results from the nonlinear full potential method are presented to show the validity of the design process along with results from linear theory codes. Longitudinal force and moment data and static pressure data were obtained in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.58, 1.62, 1.66, 1.70, and 2.00 over an angle of attack range of 0 to 14 deg at a Reynolds number of 2.0 x 10 to the 6th power per foot. Oil flow photographs of the upper surface were obtained at M = 1.62 for alpha approx. = 8, 10, 12, and 14 deg.

Numerical simulation of interaction of long-wave disturbances with a shock wave on a wedge for the problem of mode decomposition of supersonic flow oscillations

NASA Astrophysics Data System (ADS)

Kirilovskiy, S. V.; Poplavskaya, T. V.; Tsyryulnikov, I. S.

2016-10-01

This work is aimed at obtaining conversion factors of free stream disturbances from shock wave angle φ, angle of acoustic disturbances distribution θ and Mach number M∞ by solving a problem of interaction of long-wave (with the wavelength λ greater than the model length) free-stream disturbances with a shock wave formed in a supersonic flow around the wedge. Conversion factors at x/λ=0.2 as a ration between amplitude of pressure pulsations on the wedge surface and free stream disturbances amplitude were obtained. Factors of conversion were described by the dependence on angle θ of disturbances distribution, shock wave angle φ and Mach number M∞. These dependences are necessary for solving the problem of mode decomposition of disturbances in supersonic flows in wind tunnels.

Historical background and design evolution of the transonic aircraft technology supercritical wing

NASA Technical Reports Server (NTRS)

Ayers, T. G.; Hallissy, J. B.

1981-01-01

Two dimensional wind tunnel test results obtained for supercritical airfoils indicated that substantial improvements in aircraft performance at high subsonic speeds could be achieved by shaping the airfoil to improve the supercritical flow above the upper surface. Significant increases in the drag divergence Mach number, the maximum lift coefficient for buffer onset, and the Mach number for buffet onset at a given lift coefficient were demonstrated for the supercritical airfoil, as compared with a NACA 6 series airfoil of comparable thickness. These trends were corroborated by results from three dimensional wind tunnel and flight tests. Because these indicated extensions of the buffet boundaries could provide significant improvements in the maneuverability of a fighter airplane, an exploratory wind tunnel investigation was initiated which demonstrated that significant aerodynamic improvements could be achieved from the direct substitution of a supercritical airfoil on a variable wing sweep multimission airplane model.

Numerical simulation of supersonic gap flow.

PubMed

Jing, Xu; Haiming, Huang; Guo, Huang; Song, Mo

2015-01-01

Various gaps in the surface of the supersonic aircraft have a significant effect on airflows. In order to predict the effects of attack angle, Mach number and width-to-depth ratio of gap on the local aerodynamic heating environment of supersonic flow, two-dimensional compressible Navier-Stokes equations are solved by the finite volume method, where convective flux of space term adopts the Roe format, and discretization of time term is achieved by 5-step Runge-Kutta algorithm. The numerical results reveal that the heat flux ratio is U-shaped distribution on the gap wall and maximum at the windward corner of the gap. The heat flux ratio decreases as the gap depth and Mach number increase, however, it increases as the attack angle increases. In addition, it is important to find that chamfer in the windward corner can effectively reduce gap effect coefficient. The study will be helpful for the design of the thermal protection system in reentry vehicles.

A Method of Determining Aerodynamic-Influence Coefficients from Wind-Tunnel Data for Wings at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Gainer, Patrick A.

1961-01-01

A method is described for determining aerodynamic-influence coefficients from wind-tunnel data for calculating the steady-state load distribution on a wing with arbitrary angle-of-attack distribution at supersonic speeds. The method combines linearized theory with empirical adjustments in order to give accurate results over a wide range of angles of attack. The experimented data required are pressure distributions measured on a flat wing of the desired planform at the desired Mach number and over the desired range of angles of attack. The method has been tested by applying it to wind-tunnel data measured at Mach numbers of 1.61 and 2.01 on wings of the same planform but of different surface shapes. Influence coefficients adjusted to fit the flat wing gave good predictions of the spanwise and chord-wise distributions of loadings measured on twisted and cambered wings.

Analysis of Tangential Slot Blowing on F/A-18 Isolated Forebody

NASA Technical Reports Server (NTRS)

Gee, Ken; Rizk, Yehia M.; Schiff, Lewis B.

1995-01-01

The generation of significant side forces and yawing moments on an F/A-18 fuselage through tangential slot blowing is analyzed using computational fluid dynamics. The effects of freestream Mach number, jet exit conditions, jet length, and jet location are studied. The effects of over- and underblowing on force and moment production are analyzed. Non-time-accurate solutions are obtained to determine the steady-state side forces, yawing moments, and surface pressure distributions generated by tangential slot blowing. Time-accurate solutions are obtained to study the force onset time lag of tangential slot blowing. Comparison with available experimental data from full-scale wind-tunnel and subscale wind-tunnel tests are made. This computational analysis complements the experimental results and provides a detailed understanding of the effects of tangential slot blowing on the flowfield about the isolated F/A-18 forebody. Additionally, it extends the slot-blowing database to transonic maneuvering Mach numbers.

Hypersonic lateral and directional stability characteristics of aeroassist flight experiment configuration in air and CF4

NASA Technical Reports Server (NTRS)

Micol, John R.; Wells, William L.

1993-01-01

Hypersonic lateral and directional stability characteristics measured on a 60 deg half-angle elliptical cone, which was raked at an angle of 73 deg from the cone centerline and with an ellipsoid nose (ellipticity equal to 2.0 in the symmetry plane), are presented for angles of attack from -10 to 10 deg. The high normal-shock density ratio of a real gas was simulated by tests at a Mach number of 6 in air and CF4 (density ratio equal to 5.25 and 12.0, respectively). Tests were conducted in air at Mach 6 and 10 and in CF4 at Mach 6 to examine the effects of Mach number, Reynolds number, and normal-shock density ratio. Changes in Mach number from 6 to 10 in air or in Reynolds number by a factor of 4 at Mach 6 had a negligible effect on lateral and directional stability characteristics. Variations in normal-shock density ratio had a measurable effect on lateral and directional aerodynamic coefficients, but no significant effect on lateral and directional stability characteristics. Tests in air and CF4 indicated that the configuration was laterally and directionally stable through the test range of angle of attack.

OPACITY BROADENING OF {sup 13}CO LINEWIDTHS AND ITS EFFECT ON THE VARIANCE-SONIC MACH NUMBER RELATION

DOE Office of Scientific and Technical Information (OSTI.GOV)

Correia, C.; De Medeiros, J. R.; Burkhart, B.

2014-04-10

We study how the estimation of the sonic Mach number (M{sub s} ) from {sup 13}CO linewidths relates to the actual three-dimensional sonic Mach number. For this purpose we analyze MHD simulations that include post-processing to take radiative transfer effects into account. As expected, we find very good agreement between the linewidth estimated sonic Mach number and the actual sonic Mach number of the simulations for optically thin tracers. However, we find that opacity broadening causes M{sub s} to be overestimated by a factor of ≈1.16-1.3 when calculated from optically thick {sup 13}CO lines. We also find that there ismore » a dependence on the magnetic field: super-Alfvénic turbulence shows increased line broadening compared with sub-Alfvénic turbulence for all values of optical depth for supersonic turbulence. Our results have implications for the observationally derived sonic Mach number-density standard deviation (σ{sub ρ/(ρ)}) relationship, σ{sub ρ/〈ρ〉}{sup 2}=b{sup 2}M{sub s}{sup 2}, and the related column density standard deviation (σ {sub N/(N)}) sonic Mach number relationship. In particular, we find that the parameter b, as an indicator of solenoidal versus compressive driving, will be underestimated as a result of opacity broadening. We compare the σ {sub N/(N)}-M{sub s} relation derived from synthetic dust extinction maps and {sup 13}CO linewidths with recent observational studies and find that solenoidally driven MHD turbulence simulations have values of σ {sub N/(N)}which are lower than real molecular clouds. This may be due to the influence of self-gravity which should be included in simulations of molecular cloud dynamics.« less

Skin-Friction Measurements at Subsonic and Transonic Mach Numbers with Embedded-Wire Gages

DTIC Science & Technology

1981-01-01

Model ................................... 17 9. Boundary-Layer Rake Installation on EBOR Model...boundary-layer total pressure rake eliminates this bulky mechanism and the long data acquisition time, but it introduces interferences which affect the...its construction. Further, boundary-layer rakes are restricted to measurements in thick boundary layers. Surface pressure probes such as Stanton tubes

Inlet Performance Characteristics from Wind-Tunnel Tests of a 0.10-Scale Air-Induction System Model of the YF-108A Airplane at Mach Numbers of 2.50, 2.76, and 3.00

NASA Technical Reports Server (NTRS)

Blackaby, James R.; Lyman, E. Gene; Altermann, John A., III

1959-01-01

Inlet-performance and external-drag-coefficient characteristics are presented without analysis. Effects are shown of variations of fuselage boundary-layer diverter profile, bleed-surface porosity, bleed-exit area, and inlet ramp, and lip angle.

Parametric investigation of single-expansion-ramp nozzles at Mach numbers from 0.60 to 1.20

NASA Technical Reports Server (NTRS)

Capone, Francis J.; Re, Richard J.; Bare, E. Ann

1992-01-01

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of varying six nozzle geometric parameters on the internal and aeropropulsive performance characteristics of single-expansion-ramp nozzles. This investigation was conducted at Mach numbers from 0.60 to 1.20, nozzle pressure ratios from 1.5 to 12, and angles of attack of 0 deg +/- 6 deg. Maximum aeropropulsive performance at a particular Mach number was highly dependent on the operating nozzle pressure ratio. For example, as the nozzle upper ramp length or angle increased, some nozzles had higher performance at a Mach number of 0.90 because of the nozzle design pressure was the same as the operating pressure ratio. Thus, selection of the various nozzle geometric parameters should be based on the mission requirements of the aircraft. A combination of large upper ramp and large lower flap boattail angles produced greater nozzle drag coefficients at Mach number greater than 0.80, primarily from shock-induced separation on the lower flap of the nozzle. A static conditions, the convergent nozzle had high and nearly constant values of resultant thrust ratio over the entire range of nozzle pressure ratios tested. However, these nozzles had much lower aeropropulsive performance than the convergent-divergent nozzle at Mach number greater than 0.60.

The Effect of Blunt-Trailing-Edge Modifications on the High-Speed Stability and Control Characteristics of a Swept-Wing Fighter Airplane

NASA Technical Reports Server (NTRS)

Sadoff, Melvin; Matteson, Frederick H.; Van Dyke, Rudolph D., Jr.

1954-01-01

An investigation was conducted on a 35 deg swept-wing fighter airplane to determine the effects of several blunt-trailing-edge modifications to the wing and tail on the high-speed stability and control characteristics and tracking performance. The results indicated significant improvement in the pitch-up characteristics for the blunt-aileron configuration at Mach numbers around 0.90. As a result of increased effectiveness of the blunt-trailing-edge aileron, the roll-off, customarily experienced with the unmodified airplane in wings-level flight between Mach numbers of about 0.9 and 1.0 was eliminated, The results also indicated that the increased effectiveness of the blunt aileron more than offset the large associated aileron hinge moment, resulting in significant improvement in the rolling performance at Mach numbers between 0.85 and 1.0. It appeared from these results that the tracking performance with the blunt-aileron configuration in the pitch-up and buffeting flight region at high Mach numbers was considerably improved over that of the unmodified airplane; however, the tracking errors of 8 to 15 mils were definitely unsatisfactory. A drag increment of about O.OOl5 due to the blunt ailerons was noted at Mach numbers to about 0.85. The drag increment was 0 at Mach numbers above 0.90.

Mach number effect on jet impingement heat transfer.

PubMed

Brevet, P; Dorignac, E; Vullierme, J J

2001-05-01

An experimental investigation of heat transfer from a single round free jet, impinging normally on a flat plate is described. Flow at the exit plane of the jet is fully developed and the total temperature of the jet is equal to the ambient temperature. Infrared measurements lead to the characterization of the local and averaged heat transfer coefficients and Nusselt numbers over the impingement plate. The adiabatic wall temperature is introduced as the reference temperature for heat transfer coefficient calculation. Various nozzle diameters from 3 mm to 15 mm are used to make the injection Mach number M vary whereas the Reynolds number Re is kept constant. Thus the Mach number influence on jet impingement heat transfer can be directly evaluated. Experiments have been carried out for 4 nozzle diameters, for 3 different nozzle-to-target distances, with Reynolds number ranging from 7200 to 71,500 and Mach number from 0.02 to 0.69. A correlation is obtained from the data for the average Nusselt number.

Boundary-Layer Edge Conditions and Transition Reynolds Number Data for a Flight Test at Mach 20 (Reentry F)

NASA Technical Reports Server (NTRS)

Johnson, Charles B.; Stainback, P. Calvin; Wicker, Kathleen C.; Boney, Lillian R.

1972-01-01

A flight experiment, designated Reentry F, was conducted to measure heat-transfer rates for laminar, transitional, and turbulent boundary layers on a 5 deg half-angle cone 3.962 m (13 ft) long with a preflight nose radius of 2.54 mm (0.10 in.). Data were obtained over an altitude range from 36.58 to 18.29 km (120 000 to 60 000 ft) at a flight velocity of about 6.096 km/sec (20 000 ft/sec). The nominal values of the free-stream total enthalpy, sharp-cone Mach number, and the wall-to-total enthalpy ratio were 18 MJ/kg (8000 Btu/lb), 15, and 0.03, respectively. Calculated boundary-layer edge conditions that account for effects of the entropy layer and corresponding local transition Reynolds numbers are reported in the present paper. Fully developed turbulent flow occurred with essentially constant boundary-layer edge conditions near the sharp-cone values. Transition data were obtained with local edge Mach numbers ranging from about 5.55 to 15. Transition Reynolds numbers, based on local condition, were as high as 6.6 x 10(exp 7) with an edge Mach number of about 14.4 at an altitude of 24.38 km (80 000 ft). The transition could be correlated with previous flight data taken over a Mach number range from 3 to 12 in terms of parameters including the effects of local unit Reynolds number, boundary-layer wall-to-edge enthalpy ratio, and local Mach number.

Subsonic and transonic pressure measurements on a high-aspect-ratio supercritical-wing model with oscillating control surfaces

NASA Technical Reports Server (NTRS)

Sandford, M. C.; Ricketts, R. H.; Watson, J. J.

1981-01-01

A high aspect ratio supercritical wing with oscillating control surfaces is described. The semispan wing model was instrumented with 252 static orifices and 164 in situ dynamic pressure gases for studying the effects of control surface position and sinusoidal motion on steady and unsteady pressures. Data from the present test (this is the second in a series of tests on this model) were obtained in the Langley Transonic Dynamics Tunnel at Mach numbers of 0.60 and 0.78 and are presented in tabular form.

Analysis of Fluctuating Static Pressure Measurements in a Large High Reynolds Number Transonic Cryogenic Wind Tunnel. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Igoe, William B.

1991-01-01

Dynamic measurements of fluctuating static pressure levels were made using flush mounted high frequency response pressure transducers at eleven locations in the circuit of the National Transonic Facility (NTF) over the complete operating range of this wind tunnel. Measurements were made at test section Mach numbers from 0.2 to 1.2, at pressure from 1 to 8.6 atmospheres and at temperatures from ambient to -250 F, resulting in dynamic flow disturbance measurements at the highest Reynolds numbers available in a transonic ground test facility. Tests were also made independently at variable Mach number, variable Reynolds number, and variable drivepower, each time keeping the other two variables constant thus allowing for the first time, a distinct separation of these three important variables. A description of the NTF emphasizing its flow quality features, details on the calibration of the instrumentation, results of measurements with the test section slots covered, downstream choke, effects of liquid nitrogen injection and gaseous nitrogen venting, comparisons between air and nitrogen, isolation of the effects of Mach number, Reynolds number, and fan drive power, and identification of the sources of significant flow disturbances is included. The results indicate that primary sources of flow disturbance in the NTF may be edge-tones generated by test section sidewall re-entry flaps and the venting of nitrogen gas from the return leg of the tunnel circuit between turns 3 and 4 in the cryogenic mode of operation. The tests to isolate the effects of Mach number, Reynolds number, and drive power indicate that Mach number effects predominate. A comparison with other transonic wind tunnels shows that the NTF has low levels of test section fluctuating static pressure especially in the high subsonic Mach number range from 0.7 to 0.9.

Control of the transition between regular and mach reflection of shock waves

NASA Astrophysics Data System (ADS)

Alekseev, A. K.

2012-06-01

A control problem was considered that makes it possible to switch the flow between stationary Mach and regular reflection of shock waves within the dual solution domain. The sensitivity of the flow was computed by solving adjoint equations. A control disturbance was sought by applying gradient optimization methods. According to the computational results, the transition from regular to Mach reflection can be executed by raising the temperature. The transition from Mach to regular reflection can be achieved by lowering the temperature at moderate Mach numbers and is impossible at large numbers. The reliability of the numerical results was confirmed by verifying them with the help of a posteriori analysis.

Detailed noise measurements on the SR-7A propeller: Tone behavior with helical tip Mach number

NASA Technical Reports Server (NTRS)

Dittmar, James H.; Hall, David G.

1991-01-01

Detailed noise measurements were taken on the SR-7A propeller to investigate the behavior of the noise with helical tip Mach number and then to level off as Mach number was increased further. This behavior was further investigated by obtaining detailed pressure-time histories of data. The pressure-time histories indicate that a portion of the primary pressure pulse is progressively cancelled by a secondary pulse which results in the noise leveling off as the helical tip Mach number is increased. This second pulse appears to originate on the same blade as the primary pulse and is in some way connected to the blade itself. This leaves open the possibility of redesigning the blade to improve the cancellation; thereby, the propeller noise is reduced.

Pfirsch–Schlüter neoclassical heavy impurity transport in a rotating plasma

DOE PAGES

Belli, Emily A.; Candy, Jefferey M.; Angioni, C.

2014-11-07

In this paper, we extend previous analytic theories for the neoclassical transport of a trace heavy impurity in a rotating plasma in the Pfirsch-Schl¨uter regime. The complete diffusive and convective components of the ambipolar particle flux are derived. The solution is valid for arbitrary impurity charge and impurity Mach number and for general geometry. Inclusion of finite main ion temperature gradient effects is shown in the small ion Mach number limit. A simple interpolation formula is derived for the case of high impurity charge and circular geometry. While an enhancement of the diffusion coefficient is found for order one impuritymore » Mach number, a reduction due to the rotation-driven poloidal asymmetry in the density occurs for very large Mach number.« less

Detailed noise measurements on the SR-7A propeller: Tone behavior with helical tip Mach number

NASA Astrophysics Data System (ADS)

Dittmar, James H.; Hall, David G.

1991-12-01

Detailed noise measurements were taken on the SR-7A propeller to investigate the behavior of the noise with helical tip Mach number and then to level off as Mach number was increased further. This behavior was further investigated by obtaining detailed pressure-time histories of data. The pressure-time histories indicate that a portion of the primary pressure pulse is progressively cancelled by a secondary pulse which results in the noise leveling off as the helical tip Mach number is increased. This second pulse appears to originate on the same blade as the primary pulse and is in some way connected to the blade itself. This leaves open the possibility of redesigning the blade to improve the cancellation; thereby, the propeller noise is reduced.

Jet noise modification by the 'whistler nozzle'

NASA Technical Reports Server (NTRS)

Hasan, M. A. Z.; Islam, O.; Hussain, A. K. M. F.

1984-01-01

The farfield noise characteristics of a subsonic whistler nozzle jet are measured as a function of Mach number (0.25, 0.37, and, 0.51), emission angle, and excitation mode. It is shown that a whistler nozzle has greater total and broadband acoustic power than an excited contraction nozzle; and that the intensity of far-field noise is a function of emission angle, Mach number, and whistler excitation stage. The whistler nozzle excitation produces broadband noise amplification with constant spectral shape; the broadband noise amplification (without associated whistler tones and harmonics) increases omnidirectionally with emission angle at all Mach numbers; and the broadband amplification factor decreases as Mach number and emission angle increase. Finally the whistler nozzle is described as a very efficient but inexpensive siren with applications in not only jet excitation but also acoustics.

An experimental investigation of nacelle-pylon installation on an unswept wing at subsonic and transonic speeds

NASA Technical Reports Server (NTRS)

Carlson, J. R.; Compton, W. B., III

1984-01-01

A wind tunnel investigation was conducted to determine the aerodynamic interference associated with the installation of a long duct, flow-through nacelle on a straight unswept untapered supercritical wing. Experimental data was obtained for the verification of computational prediction techniques. The model was tested in the 16-Foot Transonic Tunnel at Mach numbers from 0.20 to 0.875 and at angles of attack from about 0 deg to 5 deg. The results of the investigation show that strong viscous and compressibility effects are present at the transonic Mach numbers. Numerical comparisons show that linear theory is adequate for subsonic Mach number flow prediction, but is inadequate for prediction of the extreme flow conditions that exist at the transonic Mach numbers.

Measurements of the Time-Averaged and Instantaneous Induced Velocities in the Wake of a Helicopter Rotor Hovering at High Tip Speeds

NASA Technical Reports Server (NTRS)

Heyson, Harry H.

1960-01-01

Measurements of the time-averaged induced velocities were obtained for rotor tip speeds as great as 1,100 feet per second (tip Mach number of 0.98) and measurements of the instantaneous induced velocities were obtained for rotor tip speeds as great as 900 feet per second. The results indicate that the small effects on the wake with increasing Mach number are primarily due to the changes in rotor-load distribution resulting from changes in Mach number rather than to compressibility effects on the wake itself. No effect of tip Mach number on the instantaneous velocities was observed. Under conditions for which the blade tip was operated at negative pitch angles, an erratic circulatory flow was observed.

Low Reynolds Number Aerodynamic Characteristics of Several Airplane Configurations Designed to Fly in the Mars Atmosphere at Subsonic Speeds

NASA Technical Reports Server (NTRS)

Re, Richard J.; Pendergraft, Odis C., Jr.; Campbell, Richard L.

2006-01-01

A 1/4-scale wind tunnel model of an airplane configuration developed for short duration flight at subsonic speeds in the Martian atmosphere has been tested in the Langley Research Center Transonic Dynamics Tunnel. The tunnel was pumped down to extremely low pressures to represent Martian Mach/Reynolds number conditions. Aerodynamic data were obtained and upper and lower surface wind pressures were measured at one spanwise station on some configurations. Three unswept wings of the same planform but different airfoil sections were tested. Horizontal tail incidence was varied as was the deflection of plain and split trailing-edge flaps. One unswept wing configuration was tested with the lower part of the fuselage removed and the vertical/horizontal tail assembly inverted and mounted from beneath the fuselage. A sweptback wing was also tested. Tests were conducted at Mach numbers from 0.50 to 0.90. Wing chord Reynolds number was varied from 40,000 to 100,000 and angles of attack and sideslip were varied from -10deg to 20deg and -10deg to 10deg, respectively.

Aerothermal tests of a 12.5 percent cone at Mach 6.7 for various Reynolds numbers, angles of attack and nose shapes. [conducted in Langley 8-foot high temperature tunnel

NASA Technical Reports Server (NTRS)

Nowak, R. J.; Albertson, C. W.; Hunt, L. R.

1984-01-01

The effects of free-stream unit Reynolds number, angle of attack, and nose shape on the aerothermal environment of a 3-ft basediameter, 12.5 deg half-angle cone were investigated in the Langley 8-foot high temperature tunnel at Mach 6.7. The average total temperature was 3300 R, the freestream unit Reynolds number ranged from 400,000 to 1,400,000 per foot, and the angle of attack ranged from 0 deg to 10 deg. Three nose configurations were tested on the cone: a 3-in-radius tip, a 1-in-radius tip on an ogive frustum, and a sharp tip on an ogive frustum. Surface-pressure and cold-wall heating-rate distributions were obtained for laminar, transitional temperature in the shock layer were obtained. The location of the start of transition moved forward both on windward and leeward sides with increasing free-stream Reynolds numbers, increasing angle of attack, and decreasing nose bluntness.

Boundary-Layer Transition on a Group of Blunt Nose Shapes at a Mach Number of 2.20

NASA Technical Reports Server (NTRS)

Jackson, Mary W.; Czarnecki, K. R.

1961-01-01

An investigation has been made to study boundary-layer transition on six axisymmetrical blunt bodies of revolution. Model shapes were selected with respect to the degree of favorable pressure gradient over the model surface. Tests were conducted at a Mach number of 2.20 and over a range of free-stream Reynolds number per foot of about 1.4 x 10(exp 6) to 6.5 x 10(exp 6). The tests were made at an angle of attack of 0 deg. with zero heat transfer. For the hemisphere, the flow remained essentially laminar over the model surface length for the entire pressure range of the tests. For a strong favorable pressure gradient followed by any weak favorable, neutral, or adverse gradient, the tendency was for transition to occur at or immediately behind the shoulder. A single strip of three-dimensional roughness in the region of strong favorable pressure gradient did not fix transition on the models at the roughness location except at the maximum test pressures, whereas a second roughness strip added in a region of neutral or adverse pressure gradient did fix transition. Experimental pressure coefficients agreed closely with modified Newtonian theory except in the shoulder region.

Transition Analysis for the HIFiRE-1 Flight Experiment

NASA Technical Reports Server (NTRS)

Li, Fei; Choudhari, Meelan M.; Chang, Chau-Lyan; Kimmel, Roger; Adamczak, David; Smith, Mark S.

2011-01-01

The HIFiRE-1 flight experiment provided a valuable database pertaining to boundary layer transition over a 7-degree half-angle, circular cone model from supersonic to hypersonic Mach numbers, and a range of Reynolds numbers and angles of incidence. This paper reports the initial findings from the ongoing computational analysis pertaining to the measured in-flight transition behavior. Transition during the ascent phase at nearly zero degree angle of attack is dominated by second mode instabilities except in the vicinity of the cone meridian where a roughness element was placed midway along the length of the cone. The first mode instabilities were found to be weak at all trajectory points analyzed from the ascent phase. For times less than approximately 18.5 seconds into the flight, the peak amplification ratio for second mode disturbances is sufficiently small because of the lower Mach numbers at earlier times, so that the transition behavior inferred from the measurements is attributed to an unknown physical mechanism, potentially related to step discontinuities in surface height near the locations of a change in the surface material. Based on the time histories of temperature and/or heat flux at transducer locations within the aft portion of the cone, the onset of transition correlated with a linear PSE N-factor of approximately 14.

Implementation of Preconditioned Dual-Time Procedures in OVERFLOW

NASA Technical Reports Server (NTRS)

Pandya, Shishir A.; Venkateswaran, Sankaran; Pulliam, Thomas H.; Kwak, Dochan (Technical Monitor)

2003-01-01

Preconditioning methods have become the method of choice for the solution of flowfields involving the simultaneous presence of low Mach and transonic regions. It is well known that these methods are important for insuring accurate numerical discretization as well as convergence efficiency over various operating conditions such as low Mach number, low Reynolds number and high Strouhal numbers. For unsteady problems, the preconditioning is introduced within a dual-time framework wherein the physical time-derivatives are used to march the unsteady equations and the preconditioned time-derivatives are used for purposes of numerical discretization and iterative solution. In this paper, we describe the implementation of the preconditioned dual-time methodology in the OVERFLOW code. To demonstrate the performance of the method, we employ both simple and practical unsteady flowfields, including vortex propagation in a low Mach number flow, flowfield of an impulsively started plate (Stokes' first problem) arid a cylindrical jet in a low Mach number crossflow with ground effect. All the results demonstrate that the preconditioning algorithm is responsible for improvements to both numerical accuracy and convergence efficiency and, thereby, enables low Mach number unsteady computations to be performed at a fraction of the cost of traditional time-marching methods.

Nozzle wall roughness effects on free-stream noise and transition in the pilot low-disturbance tunnel

NASA Technical Reports Server (NTRS)

Creel, T. R., Jr.; Beckwith, I. E.; Chen, F. J.

1985-01-01

An investigation at Mach 3.5 into the effects of nozzle wall roughness on free stream pressure fluctuations and cone transition Reynolds numbers was conducted in the pilot low disturbance tunnel at the Langley Research Center. Nozzle wall roughness caused by either particle deposits or imperfections in surface finish increased free stream noise levels and reduced the transition Reynolds numbers on a cone mounted in the test rhombus.

Transonic transport study: Economics

NASA Technical Reports Server (NTRS)

Smith, C. L.; Wilcox, D. E.

1972-01-01

An economic analysis was performed to evaluate the impact of advanced materials, increased aerodynamic and structural efficiencies, and cruise speed on advanced transport aircraft designed for cruise Mach numbers of .90, .98, and 1.15. A detailed weight statement was generated by an aircraft synthesis computer program called TRANSYN-TST; these weights were used to estimate the cost to develop and manufacture a fleet of aircraft of each configuration. The direct and indirect operating costs were estimated for each aircraft, and an average return on investment was calculated for various operating conditions. There was very little difference between the operating economics of the aircraft designed for Mach numbers .90 and .98. The Mach number 1.15 aircraft was economically marginal in comparison but showed significant improvements with the application of carbon/epoxy structural material. However, the Mach .90 and Mach .98 aircraft are the most economically attractive vehicles in the study.

The Effect of Leading-Edge Sweep and Surface Inclination on the Hypersonic Flow Field Over a Blunt Flat Plate

NASA Technical Reports Server (NTRS)

Creager, Marcus O.

1959-01-01

An investigation of the effects of variation of leading-edge sweep and surface inclination on the flow over blunt flat plates was conducted at Mach numbers of 4 and 5.7 at free-stream Reynolds numbers per inch of 6,600 and 20,000, respectively. Surface pressures were measured on a flat plate blunted by a semicylindrical leading edge over a range of sweep angles from 0 deg to 60 deg and a range of surface inclinations from -10 deg to +10 deg. The surface pressures were predicted within an average error of +/- 8 percent by a combination of blast-wave and boundary-layer theory extended herein to include effects of sweep and surface inclination. This combination applied equally well to similar data of other investigations. The local Reynolds number per inch was found to be lower than the free-stream Reynolds number per inch. The reduction in local Reynolds number was mitigated by increasing the sweep of the leading edge. Boundary-layer thickness and shock-wave shape were changed little by the sweep of the leading edge.

Experimental Effects of Propulsive Jets and Afterbody Configurations on the Zero-lift Drag of Bodies of Revolution at a Mach Number of 1.59

NASA Technical Reports Server (NTRS)

De Moraes, Carlos A; Nowitzky, Albin M

1954-01-01

The present investigation was made at a free-stream Mach number of 1.59 to compare the afterbody drags to a series of conical boattailed models at zero angle of attack. Afterbody drags were obtained for both the power-off and the power-on conditions. Power-on drags were obtained as a function of afterbody fineness ratio, jet pressure ratio and divergence, and jet Mach number.

Aerodynamic and Thermal Performance Characteristics of Supersonic-X Type Decelerators at Mach Number 8

DTIC Science & Technology

1977-02-01

SUPERSONIC -X TYPE DECELERATORS AT MACH NUMBER 8 t’z.r I # I JJ’, o,. VON KARMAN GAS DYNAMICS FACILITY ARNOLD ENGINEERING DEVELOPMENT CENTER AIR FORCE...AERODYNAMIC AND THERMAL PERFORMANCE CHARACTERISTICS OF SUPERSONIC - X TYPE DECELERATORS AT MACH NUMBER 8 ’ 7 AU THORCs,: p ; J . D. Corce , ARO, Inc...pe r fo rmance cha rac t e r i s t i c s of model nylon, Kevlar 29, and Bisbenzimidazobenzophenanthroline Supersonic -X type parachutes behind a

The impact of materials technology and operational constraints on the economics of cruise speed selection

NASA Technical Reports Server (NTRS)

Clauss, J. S., Jr.; Bruckman, F. A.; Horning, D. L.; Johnston, R. H.; Werner, J. V.

1981-01-01

Six material concepts at Mach 2.0 and three material concepts at Mach 2.55 were proposed. The resulting evaluations, based on projected development, production, and operating costs, indicate that aircraft designs with advanced composites as the primary material ingredient have the lowest fare premiums at both Mach 2.0 and 2.55. Designs having advanced metallics as the primary material ingredient are not economical. Advanced titanium, employing advanced manufacturing methods such as SFF/DB, requires a fare premium of about 30 percent at both Mach 2.0 and 2.55. Advanced aluminum, usable only at the lower Mach number, requires a fare premium of 20 percent. Cruise speeds in the Mach 2.0-2.3 regime are preferred because of the better economics and because of the availability of two material concepts to reduce program risk - advanced composites and advanced aluminums. This cruise speed regime also avoids the increase in risk associated with the more complex inlets and airframe systems and higher temperature composite matrices required at the higher Mach numbers typified by Mach 2.55.

Limitations of the method of characteristics when applied to axisymmetric hypersonic nozzle design

NASA Technical Reports Server (NTRS)

Edwards, Anne C.; Perkins, John N.; Benton, James R.

1990-01-01

A design study of axisymmetric hypersonic wind tunnel nozzles was initiated by NASA Langley Research Center with the objective of improving the flow quality of their ground test facilities. Nozzles for Mach 6 air, Mach 13.5 nitrogen, and Mach 17 nitrogen were designed using the Method of Characteristics/Boundary Layer (MOC/BL) approach and were analyzed with a Navier-Stokes solver. Results of the analysis agreed well with design for the Mach 6 case, but revealed oblique shock waves of increasing strength originating from near the inflection point of the Mach 13.5 and Mach 17 nozzles. The findings indicate that the MOC/BL design method has a fundamental limitation that occurs at some Mach number between 6 an 13.5. In order to define the limitation more exactly and attempt to discover the cause, a parametric study of hypersonic ideal air nozzles designed with the current MOC/BL method was done. Results of this study indicate that, while stagnations conditions have a moderate affect on the upper limit of the method, the method fails at Mach numbers above 8.0.

Effect of nose shape and tail length on supersonic stability characteristics of a projectile

NASA Technical Reports Server (NTRS)

Sawyer, W. C.; Collins, I. K.

1973-01-01

The effect of nose shape and tail length on the static stability of a fin-stabilized projectile has been investigated in the Langley Unitary Plan with tunnel at angles of attack to about 12 deg for a Mach number range from 1.5 to 2.5. The tests were made at a constant Reynolds number of 6.56 x 1,000,000 per meter. The results of the investigation showed that nose shape had no effect on the static stability. Increasing the tail length resulted in a progressively stabilizing tendency. However, only the 1.5-caliber-tail-length configuration was stable over the test angle-of-attack range at Mach number 1.5. This configuration was marginally stable or unstable at the higher Mach numbers, and the shorter configurations were unstable at all Mach numbers for either part of or the entire test angle-of-attack range.

An Experimental Evaluation of the Performance of Two Combination Pitot Pressure Probes

NASA Technical Reports Server (NTRS)

Arend, David J.; Saunders, John D.

2009-01-01

Experimental tests have been completed which recorded the ability of two combination steady state and high response time varying Pitot probe designs to accurately measure steady stagnation pressure at a single location in a flow field. Tests were conducted of double-barreled and coannular Prati probes in a 3.5 in. diameter free jet probe calibration facility from Mach 0.1 to 0.9. Geometric symmetry and pitch (-40 deg to 40 deg) and yaw (0 deg to 40 deg) angle actuation were used to fully evaluate the probes. These tests revealed that the double-barreled configuration induced error in its steady state measurement at zero incidence that increased consistently with jet Mach number to 1.1 percent at Mach 0.9. For all Mach numbers, the double-barreled probe nulled at a pitch angle of approximately 7.0 deg and provided inconsistent measurements when yawed. The double-barreled probe provided adequate measurements via both its steady state and high response tubes (within +/- 0.15 percent accuracy) over unacceptable ranges of biased pitch and inconsistent yaw angles which varied with Mach number. By comparison, the coannular probe provided accurate measurements (at zero incidence) for all jet Mach numbers as well as over a flow angularity range which varied from +/- 26.0 deg at Mach 0.3 deg to +/- 14.0 deg at Mach 0.9. Based on these results, the Prati probe is established as the preferred design. Further experimental tests are recommended to document the frequency response characteristics of the Prati probe.

Longitudinal Aerodynamic Characteristics and Effect of Rocket Jet on Drag of Models of the Hermes A-3A and A-3B Missiles in Free Flight at Mach Numbers From 0.6 to 2.0

NASA Technical Reports Server (NTRS)

Jackson, H. Herbert

1955-01-01

A free-flight investigation over a Mach number range from 0.6 to 2.0 has been conducted to determine the longitudinal aerodynamic characteristics and effect of rocket jet on zero-lift drag of 1/5-scale models of two ballistic-type missiles, the Hermes A-3A and A-3B. Models of both types of missiles exhibited very nearly linear normal forces and pitching moments over the angle-of-attack range of 8 deg to -4 deg and Mach number range tested. The centers of pressure for both missiles were not appreciably affected by Mach number over the subsonic range; however, between a Mach number of 1.02 and 1.50 the center of pressure for the A-3A model moved forward 0.34 caliber with increasing Mach number. At a trim angle-of-attack of approximately 30 deg, the A-3A model indicated a total drag coefficient 30% higher than the power-off zero-lift drag over the subsonic Mach number range and 10% higher over the supersonic range. Under the conditions of the present test, and excluding the effect of the jet on base drag, there was no indicated effect of the propulsive jet on the total drag of the A-3A model. The propulsive jet operating at a jet pressure ratio p(sub j)/p(sub o) of 0.8 caused approximately 100% increase in base drag over the Mach number range M = 0.6 to 1.0. This increase in base drag amounts to 15% of the total drag. An underexpanded jet operating at jet pressure ratios corresponding approximately to those of the full-scale missile caused a 22% reduction in base drag at M = 1.55 (p(sub j)/p(sub o) = 1.76) but indicated no change at M = 1.30 (p(sub j)/p(sub o) = 1.43). At M = 1.1 and p(sub j)/p(sub o) = 1.55, the jet caused a 50% increase in base drag.

Computational and experimental investigation of two-dimensional scramjet inlets and hypersonic flow over a sharp flat plate

NASA Astrophysics Data System (ADS)

Messitt, Donald G.

1999-11-01

The WIND code was employed to compute the hypersonic flow in the shock wave boundary layer merged region near the leading edge of a sharp flat plate. Solutions were obtained at Mach numbers from 9.86 to 15.0 and free stream Reynolds numbers of 3,467 to 346,700 in-1 (1.365 · 105 to 1.365 · 107 m-1) for perfect gas conditions. The numerical results indicated a merged shock wave and viscous layer near the leading edge. The merged region grew in size with increasing free stream Mach number, proportional to Minfinity 2/Reinfinity. Profiles of the static pressure in the merged region indicated a strong normal pressure gradient (∂p/∂y). The normal pressure gradient has been neglected in previous analyses which used the boundary layer equations. The shock wave near the leading edge was thick, as has been experimentally observed. Computed shock wave locations and surface pressures agreed well within experimental error for values of the rarefaction parameter, chi/M infinity2 < 0.3. A preliminary analysis using kinetic theory indicated that rarefied flow effects became important above this value. In particular, the WIND solution agreed well in the transition region between the merged flow, which was predicted well by the theory of Li and Nagamatsu, and the downstream region where the strong interaction theory applied. Additional computations with the NPARC code, WIND's predecessor, demonstrated the ability of the code to compute hypersonic inlet flows at free stream Mach numbers up to 20. Good qualitative agreement with measured pressure data indicated that the code captured the important physical features of the shock wave - boundary layer interactions. The computed surface and pitot pressures fell within the combined experimental and numerical error bounds for most points. The calculations demonstrated the need for extremely fine grids when computing hypersonic interaction flows.

Measurements of surface-pressure and wake-flow fluctuations in the flow field of a whitcomb supercritical airfoil

NASA Technical Reports Server (NTRS)

Roos, F. W.; Riddle, D. W.

1977-01-01

Measurements of surface pressure and wake flow fluctuations were made as part of a transonic wind tunnel investigation into the nature of a supercritical airfoil flow field. Emphasis was on a range of high subsonic Mach numbers and moderate lift coefficients corresponding to the development of drag divergence and buffeting. Fluctuation data were analyzed statistically for intensity, frequency content, and spatial coherence. Variations in these parameters were correlated with changes in the mean airfoil flow field.

Influence of Surface Roughness on the Second Order Transport of Turbulence in Non-Equilibrium Boundary Layers

DTIC Science & Technology

2006-10-08

FINAL REPORT to Air Force Office of Scientific Research (AFOSR) Project Title Influence of Surface Roughness on the Second Order Transport of...large amount of research has been performed to quantify the effects of Mach number, roughness, and wall curvature on turbulent boundary layers. However...18 a) b) c) Figure 3: a) A. D. Smith high pressure storage tank. b) Morin B series actuator controlling Virgo Engineers Trunion Mounted Ball Valve. c

Exploratory studies of the cruise performance of upper surface blown configuration: Experimental program, high-speed force tests

NASA Technical Reports Server (NTRS)

Braden, J. A.; Hancock, J. P.; Burdges, K. P.; Hackett, J. E.

1979-01-01

The work to develop a wing-nacelle arrangement to accommodate a wide range of upper surface blown configuration is reported. Pertinent model and installation details are described. Data of the effects of a wide range of nozzle geometric variations are presented. Nozzle aspect ratio, boattail angle, and chordwise position are among the parameters investigated. Straight and swept wing configurations were tested across a range of nozzle pressure ratios, lift coefficients, and Mach numbers.

A New Finite Element Supersonic Kernel Function Method in Lifting Surface Theory. Volume 2. User’s Manual

DTIC Science & Technology

1976-04-01

node. A schematic flow chart of the program is shown i& Fig. 1. Description of Variables BETA COEF IANGLE 1BUF ICHECK IMAX INFO JMAX KMAX ß...MAXINT DEL IMAX JMAX XLAMDA NMODE NP NELEM ICHECK Mach number Reduced frequency Mesh spacing as measured by the length of the side of the...Number of nodes Number of elements Option parameter used to check the mesh correctness. For ICHECK = 1, a quick run is performed to print out the

High Reynolds Number Investigation of a Flush-Mounted, S-Duct Inlet With Large Amounts of Boundary Layer Ingestion

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Carter, Melissa B.; Allan, Brian G.

2005-01-01

An experimental investigation of a flush-mounted, S-duct inlet with large amounts of boundary layer ingestion has been conducted at Reynolds numbers up to full scale. The study was conducted in the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. In addition, a supplemental computational study on one of the inlet configurations was conducted using the Navier-Stokes flow solver, OVERFLOW. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on aerodynamic interface plane diameter) from 5.1 million to 13.9 million (full-scale value), and inlet mass-flow ratios from 0.29 to 1.22, depending on Mach number. Results of the study indicated that increasing Mach number, increasing boundary layer thickness (relative to inlet height) or ingesting a boundary layer with a distorted profile decreased inlet performance. At Mach numbers above 0.4, increasing inlet airflow increased inlet pressure recovery but also increased distortion. Finally, inlet distortion was found to be relatively insensitive to Reynolds number, but pressure recovery increased slightly with increasing Reynolds number.

Pressure Fluctuations Induced by a Hypersonic Turbulent Boundary Layer

NASA Technical Reports Server (NTRS)

Duan, Lian; Choudhari, Meelan M.; Zhang, Chao

2016-01-01

Direct numerical simulations (DNS) are used to examine the pressure fluctuations generated by a spatially-developed Mach 5.86 turbulent boundary layer. The unsteady pressure field is analyzed at multiple wall-normal locations, including those at the wall, within the boundary layer (including inner layer, the log layer, and the outer layer), and in the free stream. The statistical and structural variations of pressure fluctuations as a function of wall-normal distance are highlighted. Computational predictions for mean velocity pro les and surface pressure spectrum are in good agreement with experimental measurements, providing a first ever comparison of this type at hypersonic Mach numbers. The simulation shows that the dominant frequency of boundary-layer-induced pressure fluctuations shifts to lower frequencies as the location of interest moves away from the wall. The pressure wave propagates with a speed nearly equal to the local mean velocity within the boundary layer (except in the immediate vicinity of the wall) while the propagation speed deviates from the Taylor's hypothesis in the free stream. Compared with the surface pressure fluctuations, which are primarily vortical, the acoustic pressure fluctuations in the free stream exhibit a significantly lower dominant frequency, a greater spatial extent, and a smaller bulk propagation speed. The freestream pressure structures are found to have similar Lagrangian time and spatial scales as the acoustic sources near the wall. As the Mach number increases, the freestream acoustic fluctuations exhibit increased radiation intensity, enhanced energy content at high frequencies, shallower orientation of wave fronts with respect to the flow direction, and larger propagation velocity.

Heat-Transfer Measurements in Free Flight at Mach Numbers up to 14.6 on a Flat-Faced Conical Nose with a Total Angle of 29 Degrees

NASA Technical Reports Server (NTRS)

Rumsey, Charles B; Lee, Dorothy B

1958-01-01

Skin-temperature measurements have been made at several locations on a flat-faced cone-cylinder nose which was flight tested on a fivestage rocket-propeller model to a Mach number of 14.64 and a free-stream Reynolds number of 2.0 x 10(exp 6), based on flat-face diameter, at an altitude of 66,300 feet. The copper nose had a 29 deg total-angle conical section which was 1.6 flat-face diameters long. The aerodynamic-heating rates determined from the temperature measurements reached 1,440 Btu/( sec) (sq ft) on the flat face. The heating rates near the center of the flat face agreed well at Mach numbers up to 13.6 with those obtained by a theory for laminar stagnation-point heating in equilibrium dissociated air (Avco Res. Rep. 1). At Mach numbers above 13.6, the heating rates at locations near the center of the flat face became progressively lower than stagnation-point theory and. were 29 percent lower at Mach number 14.6 at the end. of the test. The reason for this behavior of the heating on the central part of the flat face was not determined. Excluding the relatively low heating rates that occurred on the central part of the nose at the highest Mach numbers, the distribution of experimental heating along the innermost 0.79 of the flat-face radius, expressed as a percentage of stagnation-point heating, was in fair agreement with the distribution predicted by laminar theory. At a location of 0.71 radii from the stagnation point, the experimental heating was very near 130 percent of the theoretical stagnation-point rate at Mach numbers from 11 to 14.5. The experimental beating rates on the conical section of the nose were in good agreement with laminar-cone theory using the assumption of theoretical sharp-cone static pressure on the conical section.

Real-time aerodynamic heating and surface temperature calculations for hypersonic flight simulation

NASA Technical Reports Server (NTRS)

Quinn, Robert D.; Gong, Leslie

1990-01-01

A real-time heating algorithm was derived and installed on the Ames Research Center Dryden Flight Research Facility real-time flight simulator. This program can calculate two- and three-dimensional stagnation point surface heating rates and surface temperatures. The two-dimensional calculations can be made with or without leading-edge sweep. In addition, upper and lower surface heating rates and surface temperatures for flat plates, wedges, and cones can be calculated. Laminar or turbulent heating can be calculated, with boundary-layer transition made a function of free-stream Reynolds number and free-stream Mach number. Real-time heating rates and surface temperatures calculated for a generic hypersonic vehicle are presented and compared with more exact values computed by a batch aeroheating program. As these comparisons show, the heating algorithm used on the flight simulator calculates surface heating rates and temperatures well within the accuracy required to evaluate flight profiles for acceptable heating trajectories.

Space shuttle: Aerodynamic stability, control effectiveness and drag characteristics of a shuttle orbiter configuration at Mach numbers from 0.6 to 4.96

NASA Technical Reports Server (NTRS)

Ramsey, P. E.

1972-01-01

Experimental aerodynamic investigations were conducted in the NASA/MSFC 14-inch Trisonic Wind Tunnel from Sept. 27 to Oct. 7, 1972 on a 0.004 scale model of the NR ATP baseline shuttle orbiter configuration. Six component aerodynamic force and moment data were recorded at 0 deg sideslip angle over an angle of attack range from 0 to 20 deg for Mach numbers of 0.6 to 4.96, 20 to 40 deg for Mach numbers of 0.6, 0.9, 2.99, and 4.96, and 40 to 60 deg for Mach numbers of 2.99 and 4.96. Data were obtained over a sideslip range of -10 to 10 deg at 0, 10, and 20 deg angles of attack over the Mach range and 30 and 50 deg at Mach numbers of 2.99 and 4.96. The purpose of the test was to define the buildup, performance, stability, and control characteristics of the orbiter configuration. The model parameters, were: body alone; body-wing; body-wing-tail; elevon deflections of 0, 10, -20, and -40 deg both full and split); aileron deflections of plus or minus 10 deg (full and split); rudder flares of 10 and 40 deg, and a rudder deflection of 15 deg about the 10 and 40 deg flare positions.

Flight investigation of XB-70 structural response to oscillatory aerodynamic shaker excitation and correlation with analytical results

NASA Technical Reports Server (NTRS)

Mckay, J. M.; Kordes, E. E.; Wykes, J. H.

1973-01-01

The low frequency symmetric structural response and damping characteristics of the XB-70 airplane were measured at four flight conditions: heavyweight at a Mach number of 0.87 at an altitude of 7620 meters (25,000 feet); lightweight at a Mach number of 0.86 at an altitude of 7620 meters (25,000 feet); a Mach number of 1.59 at an altitude of 11,918 meters (39.100 feet); and a Mach number of 2.38 and an altitude of 18,898 meters (62,000 feet). The flight data are compared with the response calculated by using early XB-70 design data and with the response calculated with mass, structural, and aerodynamic data updated to reflect as closely as possible the airplane characteristics at three of the flight conditions actually flown.

Effect of Mach number on the efficiency of microwave energy deposition in supersonic flow

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lashkov, V. A., E-mail: valerial180150@gmail.com; Karpenko, A. G., E-mail: aspera.2003.ru@mail.ru; Khoronzhuk, R. S.

The article is devoted to experimental and numerical studies of the efficiency of microwave energy deposition into a supersonic flow around the blunt cylinder at different Mach numbers. Identical conditions for energy deposition have been kept in the experiments, thus allowing to evaluate the pure effect of varying Mach number on the pressure drop. Euler equations are solved numerically to model the corresponding unsteady flow compressed gas. The results of numerical simulations are compared to the data obtained from the physical experiments. It is shown that the momentum, which the body receives during interaction of the gas domain modified bymore » microwave discharge with a shock layer before the body, increases almost linearly with rising of Mach number and the efficiency of energy deposition also rises.« less

An experimental and theoretical study of the aerodynamic characteristics of some generic missile concepts at Mach numbers from 2 to 6.8

NASA Technical Reports Server (NTRS)

Spearman, M. Leroy; Braswell, Dorothy O.

1994-01-01

A study has been made of the experimental and theoretical aerodynamic characteristics for some generic high-speed missile concepts at Mach numbers from 2 to 6.8. The basic body for this study had a length-to-diameter ratio of 10 with the forward half being a modified blunted ogive and the rear half being a cylinder. Modifications made to the basic body included the addition of an after body flare, the addition of highly swept cruciform wings and the addition of highly swept aft tails. The effects of some controls were also investigated with all-moving wing controls on the flared body and trailing-edge flap controls on the winged body. The results indicated that the addition of a flare, wings, or tails to the basic body all provided static longitudinal stability with varying amounts of increased axial force. The control arrangements were effective in producing increments of normal-force and pitching-moment at the lower Mach numbers. At the highest Mach number, the flap control on the winged body was ineffective in producing normal-force or pitching-moment but the all-moving wing control on the flared body, while losing pitch effectiveness, still provided normal-force increments. Calculated results obtained through the use of hypersonic impact theory were in generally good agreement with experiment at the higher Mach numbers but were not accurate at the lower Mach numbers.

Studies on shock interactions with moving cylinders using immersed boundary method

NASA Astrophysics Data System (ADS)

Luo, Kun; Luo, Yujuan; Jin, Tai; Fan, Jianren

2017-06-01

The process of shock interaction with a rigid cylinder is studied using a compressible immersed boundary method combined with a high-order weighted essentially nonoscillatory scheme. Movement of the cylinder is coupled to the flow field. First, the accuracy of the numerical scheme is validated. Then the influences of the incident shock Mach number and the cylinder diameter are discussed. The results are compared with those from cases with stationary cylinders. It is found that variation of either the incident shock Mach number or the cylinder diameter can cause different schlieren images. At a given dimensionless time, the trajectory of the upper triple point varies nonmonotonically with the incident shock Mach number while the primary reflected shock gets closer to the cylinder with increasing incident shock Mach number. For any moving case with a given incident shock Mach number and cylinder diameter, the trajectory of the upper triple point, the time evolution of the normalized vertical distance from the highest point of the primary reflected shock to the centerline of the cylinder, and the time evolution of the normalized shock detachment distance can all be predicted by linear correlation. As for the time evolution of the force exerted on the cylinder, the peak of the moving cylinder appears earlier than the stationary one in dimensionless time, with much lower value. Correlations to predict the occurrence of the peak drag and its value under different shock Mach numbers and cylinder diameters are proposed. The resulting cylinder movement is also briefly discussed.

Columbia: The first 5 flights entry heating data series. Volume 5: The side fuselage and payload bay door

NASA Technical Reports Server (NTRS)

Williams, S. D.

1984-01-01

Entry heating flight data and wind tunnel data on the side fuselage and payload bay door, Z = 400 and 440 trace aft of X/L=0.2, for the first five flights of the Space Shuttle Orbiter are presented. The heating rate data are reviewed in terms of normalized film heat transfer coefficients as a function of angle of attack, Mach number, and normal shock Reynolds number. The surface heatings rates and temperatures were obtained by the JSC NONLIN/INVERSE computer program. Time history plots of the surface heating rates and temperatures are outlined.

A study of sonic boom overpressure trends with respect to weight, altitude, Mach number, and vehicle shaping

NASA Technical Reports Server (NTRS)

Needleman, Kathy E.; Mack, Robert J.

1990-01-01

This paper presents and discusses trends in nose shock overpressure generated by two conceptual Mach 2.0 configurations. One configuration was designed for high aerodynamic efficiency, while the other was designed to produce a low boom, shaped-overpressure signature. Aerodynamic lift, sonic boom minimization, and Mach-sliced/area-rule codes were used to analyze and compute the sonic boom characteristics of both configurations with respect to cruise Mach number, weight, and altitude. The influence of these parameters on the overpressure and the overpressure trends are discussed and conclusions are given.

Subsonic static and dynamic stability characteristics of the test technique demonstrator NASP configuration

NASA Technical Reports Server (NTRS)

Boyden, Richmond P.; Dress, David A.; Fox, Charles H., Jr.; Huffman, Jarrett K.; Cruz, Christopher I.

1993-01-01

The paper describes the procedure used for and the results obtained of wind-tunnel tests of the National Aerospace Plane (NASP) configuration, which were conducted in the NASA Langley Research Center High Speed Tunnel using a blended body NASP configuration designed by the research center. Static and dynamic stability characteristics were measured at Mach numbers 0.3, 0.6, and 0.8. In addition to tests of the baseline configuration, component buildup tests with a canard surface and with a body flap were carried out. Results demonstrated a positive static stability of the baseline configuration, except at the higher angles of attack at Mach 0.8. A good agreement was found between the inphase dynamic parameters and the corresponding static data.

Space shuttle abort separation pressure investigation. Volume 1, Part A: Booster data at Mach 5

NASA Technical Reports Server (NTRS)

Trimmer, L. L.; Love, D. A.; Rampy, J. M.; Decker, J. P.; Blackwell, K. L.; Strike, W. T.

1972-01-01

Pressure data obtained from a joint Langley Research Center (LaRC)/Marshall Space Flight Center (MSFC) Space Shuttle about stage separation wind tunnel test are presented. The .00556 scale models of the McDonnell-Douglas orbiter and booster configurations were tested in proximity in Tunnel A of the Von Karman Facility (VKF), Arnold Engineering Development Center (AEDC). Mach numbers were 5.0, 3.0, and 2.0 and nominal Reynolds numbers were 1.09, 1.60, and 1.74 million per foot, respectively. Pressure data were obtained for the booster upper surface and orbiter lower surface at angles of attack of -10 deg, -5, 0, 5, and 10 deg for zero degrees sideslip. The models were tested at incidence angles of 0 and 5 deg for several separation distances and power conditions. Plug nozzles utilizing air were used to simulate booster and orbiter plumes at various altitudes along a nominal ascent trajectory. Powered conditions were 100, 50, and 0 percent of full power for the orbiter and 100, 50 and 0 percent of full power for the booster. Data were also obtained with the booster canard off in close proximity.

Second-order small-disturbance solutions for hypersonic flow over power-law bodies

NASA Technical Reports Server (NTRS)

Townsend, J. C.

1975-01-01

Similarity solutions were found which give the adiabatic flow of an ideal gas about two-dimensional and axisymmetric power-law bodies at infinite Mach number to second order in the body slenderness parameter. The flow variables were expressed as a sum of zero-order and perturbation similarity functions for which the axial variations in the flow equations separated out. The resulting similarity equations were integrated numerically. The solutions, which are universal functions, are presented in graphic and tabular form. To avoid a singularity in the calculations, the results are limited to body power-law exponents greater than about 0.85 for the two-dimensional case and 0.75 for the axisymmetric case. Because of the entropy layer induced by the nose bluntness (for power-law bodies other than cones and wedges), only the pressure function is valid at the body surface. The similarity results give excellent agreement with the exact solutions for inviscid flow over wedges and cones having half-angles up to about 20 deg. They give good agreement with experimental shock-wave shapes and surface-pressure distributions for 3/4-power axisymmetric bodies, considering that Mach number and boundary-layer displacement effects are not included in the theory.

Effect of Tail Dihedral on Lateral Control Effectiveness at High Subsonic Speeds of Differentially Deflected Horizontal-Tail Surfaces on a Configuration having a Thin Highly Tapered Wing

NASA Technical Reports Server (NTRS)

Fournier, Paul G.

1959-01-01

Tests have been conducted in the Langley high-speed 7- by 10-foot tunnel to determine the effect of tail dihedral on lateral control effectiveness of a complete-model configuration having differentially deflected horizontal-tail surfaces. Limited tests were made to determine the lateral characteristics as well as the longitudinal characteristics in sideslip. The wing had an aspect ratio of 3, a taper ratio of 0.14, 28.80 deg sweep of the quarter-chord line with zero sweep at the 80-percent-chord line, and NACA 65A004 airfoil sections. The test Mach number range extended from 0.60 to 0.92. There are only small variations in the roll effectiveness parameter C(sub iota delta) with negative tail dihedral angle. The tail size used on the test model, however, is perhaps inadequate for providing the roll rates specified by current military requirements at subsonic speeds. The lateral aerodynamic characteristics were essentially constant throughout the range of sideslip angle from 12 deg to -12 deg. A general increase in yawing moment was noted with increased negative dihedral throughout the Mach number range.

Planar shock reflection on a wedged concave reflector

NASA Astrophysics Data System (ADS)

Yu, Fan-Ming; Sheu, Kuen-Dong

2001-04-01

The investigation of shock reflection and shock diffraction phenomena upon a wedged concave reflector produced by a planar incident shock wave has been done in the shock tube facility of Institute of Aeronautics and Astronautics, National Cheng- Kung University. The experiment proceeds upon three wedged concave reflectors models the upper and lower wedge angles arrangement of them are (50 degrees, 50 degrees) - 35 degrees, 35 degrees) and (50 degrees, 35 degrees), respectively. They were tested at Mach numbers of 1.2 - 1.65 and 2.0. On the first reflector, following the regular reflection on the 50 degree-wedged surface by the incident shock wave, a Mach shock diffraction behavior has been observed as shock moves outward from the apex of the reflector. On the apex of the reflector, it behaviors as a sector of the blast shock moving on a diverging channel. On the shadowgraph pictures it has been observed there exists a pattern of gas dynamics focus upon the second reflector. The Mach reflection from the 35 degree- wedged surface as being generated by the planar incident shock wave, on which the overlapping of the two triple points from both wedged surface offers the focusing mechanism. The shock interference, which proceeds by the Mach shock reflection and the regular shock diffraction from the reflector, generates a very complicate rolling-up of slip lines system. On the third reflector, the mixed shock interference behavior has been observed of which two diffraction shocks from concave 50 degree-wedged surface and 35 degree-wedged surface interfere with each other. The measurement of the peak pressure along a ray from the model apex parallel to incident shock direction indicates that the measured maximum pressure rising is larger near the apex of the reflector. Considering the measured maximum pressure increment due to the reflection shocks indicate that the wave strength upon large apex angle reflector is greater than it is upon small apex angle reflector. However, as considering the measured maximum pressure increment following the diffraction shocks, the results show that due to the focusing process upon (35 degree, 35 degree) reflector, it is of the largest increment.

Investigation of Dive Brakes and a Dive-Recovery Flap on a High-Aspect-Ratio Wing in the Langley 8-Foot High-Speed Tunnel

NASA Technical Reports Server (NTRS)

Mattson, Axel T.

1946-01-01

The results of tests made to determine the aerodynamic characteristics of a solid brake, a slotted brake, and a dive-recovery flap mounted on a high aspect ratio wing at high Mach numbers are presented. The data were obtained in the Langley 8-foot high-speed tunnel for corrected Mach numbers up to 0.940. The results have been analyzed with regard to the suitability of dive-control devices for a proposed high-speed airplane in limiting the airplane terminal Mach number by the use of dive brakes and in achieving favorable dive-recovery characteristics by the use of a dive-recovery flap. The analysis of the results indicated that the slotted brake would limit the proposed airplane terminal Mach number to values below 0.880 for altitudes up to 35,000 feet and a wing loading of 80 pounds per square foot and the dive-recovery flap would produce trim changes required for controlled pull-outs at 25,000 feet for a Mach number range from 0.800 to 0.900. Basic changes in spanwise loading are presented to aid in the evaluation of the wing strength requirements.

Inviscid Flow Computations of the Orbital Sciences X-34 Over a Mach Number Range of 1.25 to 6.0

NASA Technical Reports Server (NTRS)

Prabhu, Ramadas K.

2001-01-01

This report documents the results of an inviscid computational study conducted on the Orbital Sciences X-34 vehicle to compute its inviscid longitudinal aerodynamic characteristics over a Mach number range of 1.25 to 6.0. The unstructured grid software FELISA was used and th e aerodynamic characteristics were computed at Mach numbers 1.25, 1.6, 2.5, 4.0, 4.63, and 6.0, and an angle of attack range of -4 to 32 degrees. These results were compared with available aerodynamic data from wind tunnel test on X-34 models. The comparison showed excellent agreement in C(sub N). The computed pitching moment compared well at Mach numbers 2.5 and higher, and at angles of attack of up to 12 deg. The agreement was not good at higher angles of attack possibly due to viscous effects. At lower Mach numbers there were significant differences between computed and measured C(sub m) values. This could not be explained. Since the present computations are inviscid, the computed C(sub A) was consistently lower than the measured values as expected.

Lee-side flow over delta wings at supersonic speeds

NASA Technical Reports Server (NTRS)

Miller, D. S.; Wood, R. M.

1985-01-01

An experimental investigation of the lee-side flow on sharp leading-edge delta wings at supersonic speeds has been conducted. Pressure data were obtained at Mach numbers from 1.5 to 2.8, and three types of flow-visualization data (oil-flow, tuft, and vapor-screen) were obtained at Mach numbers from 1.7 to 2.8 for wing leading-edge sweep angles from 52.5 deg to 75 deg. From the flow-visualization data, the lee-side flows were classified into seven distinct types and a chart was developed that defines the flow mechanism as a function of the conditions normal to the wing leading edge, specifically, angle of attack and Mach number. Pressure data obtained experimentally and by a semiempirical prediction method were employed to investigate the effects of angle of attack, leading-edge sweep, and Mach number on vortex strength and vortex position. In general, the predicted and measured values of vortex-induced normal force and vortex position obtained from experimental data have the same trends with angle of attack, Mach number, and leading-edge sweep; however, the vortex-induced normal force is underpredicted by 15 to 30 percent, and the vortex spanwise location is overpredicted by approximately 15 percent.

Uncertainty Due to Unsteady Fluid/Structure Interaction for the Ares I Vehicle Traversing the Transonic Regime

NASA Technical Reports Server (NTRS)

Bartels, Robert E.

2012-01-01

Rapid reduced-order numerical models are being investigated as candidates to simulate the dynamics of a flexible launch vehicle during atmospheric ascent. There has also been the extension of these new approaches to include gust response. These methods are used to perform aeroelastic and gust response analyses at isolated Mach numbers. Such models require a method to time march through a succession of ascent Mach numbers. An approach is presented for interpolating reduced-order models of the unsteady aerodynamics at successive Mach numbers. The transonic Mach number range is considered here since launch vehicles can suffer the highest dynamic loads through this range. Realistic simulations of the flexible vehicle behavior as it traverses this Mach number range are presented. The response of the vehicle due to gusts is computed. Uncertainties in root mean square and maximum bending moment and crew module accelerations are presented due to assumed probability distributions in design parameters, ascent flight conditions, gusts. The primary focus is on the uncertainty introduced by modeling fidelity. It is found that an unsteady reduced order model produces larger excursions in the root mean square loading and accelerations than does a quasi-steady reduced order model.

An Analysis of the Applicability of the Hypersonic Similarity Law to the Study of Flow About Bodies of Revolution at Zero Angle of Attack

NASA Technical Reports Server (NTRS)

Ehret, Dorris M.; Rossow, Vernon J.; Stevens, Victor I.

1950-01-01

The hypersonic similarity law as derived by Tsien has been investigated by comparing the pressure distributions along bodies of revolution at zero angle of attack. In making these comparisons, particular attention was given to determining the limits of Mach number and fineness ratio for which the similarity law applies. For the purpose of this investigation, pressure distributions determined by the method of characteristics for ogive cylinders for values of Mach numbers and fineness ratios varying from 1.5 to 12 were compared. Pressures on various cones and on cone cylinders were also compared in this study. The pressure distributions presented demonstrate that the hypersonic similarity law is applicable over a wider range of values of Mach numbers and fineness ratios than might be expected from the assumptions made in the derivation. This is significant since within the range of applicability of the law a single pressure distribution exists for all similarly shaped bodies for which the ratio of free-stream Mach number to fineness ratio is constant. Charts are presented for rapid determination of pressure distributions over ogive cylinders for any combination of Mach number and fineness ratio within defined limits.

Flow Separation Ahead of a Blunt Axially Symmetric Body at Mach Numbers 1.76 to 2.10

NASA Technical Reports Server (NTRS)

Moeckel, W E

1951-01-01

The pressure distribution and drag were determined for a spherical-nosed axially symmetric body with thin projecting rods at Mach numbers of 1.76, 1.93, and 2.10. The upstream projection distance of the rods was varied over a wide range to study changes in the character of the flow separation and to determine the variation of drag and pressure distribution with tip projection. Drag coefficients between 0.18 and 0.30 were obtained for most tip projections at each Mach number.

Wind-tunnel investigation at Mach numbers from 0.25 to 1.01 of a transport configuration designed to cruise at near-sonic speeds. [conducted in langley 8-foot transonic pressure tunnel

NASA Technical Reports Server (NTRS)

Langhans, R. A.; Flechner, S. G.

1972-01-01

The results of the investigation showed that the configuration exhibits a sufficiently high drag divergence Mach number to cruise at near sonic speeds. The configuration is longitudinally stable through the cruise Mach number and lift coefficient range, but at higher lift coefficients displays pitchup and becomes unstable. The configuration was directionally stable at all test conditions and laterally stable in the angle of attack range required for cruise.

Performance Characteristics of Flush and Shielded Auxiliary Exits at Mach Numbers of 1.5 to 2.0

NASA Technical Reports Server (NTRS)

Abdalla, Kaleel L.

1959-01-01

The performance characteristics of several flush and shielded auxiliary exits were investigated at Mach numbers of 1.5 to 2.0, and jet pressure ratios from jet off to 10. The results indicate that the shielded configurations produced better overall performance than the corresponding flush exits over the Mach-number and pressure-ratio ranges investigated. Furthermore, the full-length shielded exit was highest in performance of all the configurations. The flat-exit nozzle block provided considerably improved performance compared with the curved-exit nozzle block.

Analysis of gas turbine engines using water and oxygen injection to achieve high Mach numbers and high thrust

NASA Technical Reports Server (NTRS)

Henneberry, Hugh M.; Snyder, Christopher A.

1993-01-01

An analysis of gas turbine engines using water and oxygen injection to enhance performance by increasing Mach number capability and by increasing thrust is described. The liquids are injected, either separately or together, into the subsonic diffuser ahead of the engine compressor. A turbojet engine and a mixed-flow turbofan engine (MFTF) are examined, and in pursuit of maximum thrust, both engines are fitted with afterburners. The results indicate that water injection alone can extend the performance envelope of both engine types by one and one-half Mach numbers at which point water-air ratios reach 17 or 18 percent and liquid specific impulse is reduced to some 390 to 470 seconds, a level about equal to the impulse of a high energy rocket engine. The envelope can be further extended, but only with increasing sacrifices in liquid specific impulse. Oxygen-airflow ratios as high as 15 percent were investigated for increasing thrust. Using 15 percent oxygen in combination with water injection at high supersonic Mach numbers resulted in thrust augmentation as high as 76 percent without any significant decrease in liquid specific impulse. The stoichiometric afterburner exit temperature increased with increasing oxygen flow, reaching 4822 deg R in the turbojet engine at a Mach number of 3.5. At the transonic Mach number of 0.95 where no water injection is needed, an oxygen-air ratio of 15 percent increased thrust by some 55 percent in both engines, along with a decrease in liquid specific impulse of 62 percent. Afterburner temperature was approximately 4700 deg R at this high thrust condition. Water and/or oxygen injection are simple and straightforward strategies to improve engine performance and they will add little to engine weight. However, if large Mach number and thrust increases are required, liquid flows become significant, so that operation at these conditions will necessarily be of short duration.

Parametric study of shock-induced combustion in a hydrogen air system

NASA Technical Reports Server (NTRS)

Ahuja, J. K.; Tiwari, Surendra N.

1994-01-01

A numerical parametric study is conducted to simulate shock-induced combustion under various free-stream conditions and varying blunt body diameter. A steady combustion front is established if the free-stream Mach number is above the Chapman-Jouguet speed of the mixture, whereas an unsteady reaction front is established if the free-stream Mach number is below or at the Chapman-Jouguet speed of the mixture. The above two cases have been simulated for Mach 5.11 and Mach 6.46 with a projectile diameter of 15 mm. Mach 5.11, which is an underdriven case, shows an unsteady reaction front, whereas Mach 6.46, which is an overdriven case, shows a steady reaction front. Next for Mach 5. 11 reducing the diameter to 2.5 mm causes the instabilities to disappear, whereas, for Mach 6.46 increasing the diameter of the projectile to 225 mm causes the instabilities to reappear, indicating that Chapman-Jouguet speed is not the only deciding factor for these instabilities to trigger. The other key parameters are the projectile diameter, induction time, activation energy and the heat release. The appearance and disappearance of the instabilities have been explained by the one-dimensional wave interaction model.

Aerodynamic characteristics of three helicopter rotor airfoil sections at Reynolds number from model scale to full scale at Mach numbers from 0.35 to 0.90. [conducted in Langley 6 by 28 inch transonic tunnel

NASA Technical Reports Server (NTRS)

Noonan, K. W.; Bingham, G. J.

1980-01-01

An investigation was conducted in the Langely 6 by 28 inch transonic tunnel to determine the two dimensional aerodynamic characteristics of three helicopter rotor airfoils at Reynolds numbers from typical model scale to full scale at Mach numbers from about 0.35 to 0.90. The model scale Reynolds numbers ranged from about 700,00 to 1,500,000 and the full scale Reynolds numbers ranged from about 3,000,000 to 6,600,000. The airfoils tested were the NACA 0012 (0 deg Tab), the SC 1095 R8, and the SC 1095. Both the SC 1095 and the SC 1095 R8 airfoils had trailing edge tabs. The results of this investigation indicate that Reynolds number effects can be significant on the maximum normal force coefficient and all drag related parameters; namely, drag at zero normal force, maximum normal force drag ratio, and drag divergence Mach number. The increments in these parameters at a given Mach number owing to the model scale to full scale Reynolds number change are different for each of the airfoils.

Numerical Modeling of Active Flow Control in a Boundary Layer Ingesting Offset Inlet

NASA Technical Reports Server (NTRS)

Allan, Brian G.; Owens, Lewis R.; Berrier, Bobby L.

2004-01-01

This investigation evaluates the numerical prediction of flow distortion and pressure recovery for a boundary layer ingesting offset inlet with active flow control devices. The numerical simulations are computed using a Reynolds averaged Navier-Stokes code developed at NASA. The numerical results are validated by comparison to experimental wind tunnel tests conducted at NASA Langley Research Center at both low and high Mach numbers. Baseline comparisons showed good agreement between numerical and experimental results. Numerical simulations for the inlet with passive and active flow control also showed good agreement at low Mach numbers where experimental data has already been acquired. Numerical simulations of the inlet at high Mach numbers with flow control jets showed an improvement of the flow distortion. Studies on the location of the jet actuators, for the high Mach number case, were conducted to provide guidance for the design of a future experimental wind tunnel test.

Cruise noise of the SR-2 propeller model in a wind tunnel

NASA Technical Reports Server (NTRS)

Dittmar, James H.

1989-01-01

Noise data on the SR-2 model propeller were taken in the NASA Lewis Research Center 8- by 6-Foot Wind Tunnel. The maximum blade passing tone rises with increasing helical tip Mach number to a peak level at a helical tip Mach number of about 1.05; then it remains the same or decreases at higher helical tip Mach numbers. This behavior, which has been observed with other propeller models, points to the possibility of using higher propeller tip speeds to limit airplane cabin noise while maintaining high flight speed and efficiency. Noise comparisons of the straight-blade SR-2 propeller and the swept-blade SR-7A propeller showed that the tailored sweep of the SR-7A appears to be the cause of both lower peak noise levels and a slower noise increase with increasing helical tip Mach number.

Generation and Evolution of High-Mach-Number Laser-Driven Magnetized Collisionless Shocks in the Laboratory.

PubMed

Schaeffer, D B; Fox, W; Haberberger, D; Fiksel, G; Bhattacharjee, A; Barnak, D H; Hu, S X; Germaschewski, K

2017-07-14

We present the first laboratory generation of high-Mach-number magnetized collisionless shocks created through the interaction of an expanding laser-driven plasma with a magnetized ambient plasma. Time-resolved, two-dimensional imaging of plasma density and magnetic fields shows the formation and evolution of a supercritical shock propagating at magnetosonic Mach number M_{ms}≈12. Particle-in-cell simulations constrained by experimental data further detail the shock formation and separate dynamics of the multi-ion-species ambient plasma. The results show that the shocks form on time scales as fast as one gyroperiod, aided by the efficient coupling of energy, and the generation of a magnetic barrier between the piston and ambient ions. The development of this experimental platform complements present remote sensing and spacecraft observations, and opens the way for controlled laboratory investigations of high-Mach number collisionless shocks, including the mechanisms and efficiency of particle acceleration.

Farfield inflight measurements of high-speed turboprop noise

NASA Technical Reports Server (NTRS)

Balombin, J. R.; Loeffler, I. J.

1983-01-01

A flight program was carried out to determine the variation of noise level with distance from a model high-speed propeller. Noise measurements were obtained at different distances from a SR-3 propeller mounted on a JetStar aircraft, with the test instrumentation mounted on a Learjet flown in formation. The propeller was operated at 0.8 m flight Mach number, 1.12 helical tip Mach number and at 0.7 flight Mach number, 1.0 helical tip Mach number. The instantaneous pressure from individual blades was observed to rise faster at the 0.8 flight speed, than at the 0.7 M flight speed. The measured levels appeared to decrease in good agreement with a 6 dB/doubling of distance decay, over the measurement range of approximately 16 m to 100 m distance. Further extrapolation, to the distances represented by a community, would suggest that the propagated levels during cruise would not cause a serious community annoyance.

The stability of a trailing-line vortex in compressible flow

NASA Technical Reports Server (NTRS)

Stott, Jillian A. K.; Duck, Peter W.

1992-01-01

We consider the inviscid stability of the Batchelor (1964) vortex in a compressible flow. The problem is tackled numerically and also asymptotically, in the limit of large (aximuthal and streamwise) wavenumbers, together with large Mach numbers. The nature of the solution passes through different regimes as the Mach number increases, relative to the wavenumbers. At very high wavenumbers and Mach numbers, the mode which is present in the incompressible case ceases to be unstable, while new 'center mode' forms, whose stability characteristics, are determined primarily by conditions close to the vortex axis. We find that generally the flow becomes less unstable as the Mach number increases, and that the regime of instability appears generally confined to disturbances in a direction counter to the direction of the rotation of the swirl of the vortex. Throughout the paper, comparison is made between our numerical results and results obtained from the various asymptotic theories.

A simplified Mach number scaling law for helicopter rotor noise

NASA Technical Reports Server (NTRS)

Aravamudan, K. S.; Lee, A.; Harris, W. L.

1978-01-01

Mach number scaling laws are derived for the rotational and the high-frequency broadband noise from helicopter rotors. The rotational scaling law is obtained directly from the theory of Lowson and Ollerhead (1969) by exploiting the properties of the dominant terms in the expression for the complex Fourier coefficients of sound radiation from a point source. The scaling law for the high-frequency broadband noise is obtained by assuming that the noise sources are acoustically compact and computing the instantaneous pressure due to an element on an airfoil where vortices are shed. Experimental results on the correlation lengths for stationary airfoils are extended to rotating airfoils. On the assumption that the correlation length varies as the boundary layer displacement thickness, it is found that the Mach number scaling law contains a factor of Mach number raised to the exponent 5.8. Both scaling laws were verified by model tests.

Experimental aerodynamic characteristics at Mach numbers from 0.60 to 2.70 of two supersonic cruise fighter configurations

NASA Technical Reports Server (NTRS)

Dollyhigh, S. M.

1979-01-01

Two 0.085-scale full span wind-tunnel models of a Mach 1.60 design supercruiser configuration were tested at Mach numbers from 0.60 to 2.70. One model incorporated a varying dihedral (swept-up) wing to obtain the desired lateral-directional characteristics; the other incorporated more conventional twin vertical tails. The data from the wind-tunnel tests are presented without analysis.

Influence of orbital-maneuvering-system fairings and rudder flare on the transonic aerodynamic characteristics of a space shuttle orbiter

NASA Technical Reports Server (NTRS)

Ellison, J. C.

1975-01-01

An investigation was conducted in the Langley 8-foot transonic pressure tunnel to determine the influence of orbital-maneuvering-system fairings and a flared rudder on the aerodynamic characteristics of a space shuttle-orbiter configuration. Tests were made at Mach numbers from 0.4 to 1.2, at angles of attack from -1 deg to 24 deg, at angles of sideslip of 0 deg and 5 deg, and at a Reynolds number, based on model length, of 4 million. The model with the orbital-maneuvering-system fairings had a minimum untrimmed lift-drag ratio from 7.4 to 3.4 at Mach numbers from 0.4 to 1.2 and a maximum trimmed lift-drag ratio of about 3.55 at Mach 0.8 with the rudder flared 30 deg. The directional stability was increased at Mach 0.8 and 1.2 by addition of the orbital-maneuvering-system fairings and at Mach 1.2 by flaring the rudder.

Design of a variable area diffuser for a 15-inch Mach 6 open-jet tunnel

NASA Technical Reports Server (NTRS)

Loney, Norman W.

1994-01-01

The Langley 15-inch Mach 6 High Temperature Tunnel was recently converted from a Mach 10 Hypersonic Flow Apparatus. This conversion was effected to improve the capability of testing in Mach 6 air at relatively high reservoir temperatures not previously possible at Langley. Elevated temperatures allow the matching of the Mach numbers, Reynolds numbers, and ratio of wall-to-adiabatic-wall temperatures (TW/Taw) between this and the Langley 20-inch Mach 6 CF4 Tunnel. This ratio is also matched for Langley's 31-inch Mach 10 Tunnel and is an important parameter useful in the simulation of slender bodies such as National Aerospace Plane (NASP) configurations currently being studied. Having established the nozzle's operating characteristics, the decision was made to install another test section to provide model injection capability. This test section is an open-jet type, with an injection system capable of injecting a model from retracted position to nozzle centerline between 0.5 and 2 seconds. Preliminary calibrations with the new test section resulted in Tunnel blockage. This blockage phenomenon was eliminated when the conical center body in the diffuser was replaced. The issue then, is to provide a new and more efficient variable area diffuser configuration with the capability to withstand testing of larger models without sending the Tunnel into an unstart condition. Use of the 1-dimensional steady flow equation with due regard to friction and heat transfer was employed to estimate the required area ratios (exit area / throat area) in a variable area diffuser. Correlations between diffuser exit Mach number and area ratios, relative to the stagnation pressure ratios and diffuser inlet Mach number were derived. From these correlations, one can set upper and lower operating pressures and temperatures for a given diffuser throat area. In addition, they will provide appropriate input conditions for the full 3-dimensional computational fluid dynamics (CFD) code for further simulation studies.

Effects of independent variation of Mach and Reynolds numbers on the low-speed aerodynamic characteristics of the NACA 0012 airfoil section

NASA Technical Reports Server (NTRS)

Ladson, Charles L.

1988-01-01

A comprehensive data base is given for the low speed aerodynamic characteristics of the NACA 0012 airfoil section. The Langley low-turbulence pressure tunnel is the facility used to obtain the data. Included in the report are the effects of Mach number and Reynolds number and transition fixing on the aerodynamic characteristics. Presented are also comparisons of some of the results with previously published data and with theoretical estimates. The Mach number varied from 0.05 to 0.36. The Reynolds number, based on model chord, varied from 3 x 10 to the 6th to 12 x 10 to the 6th power.

Convective heat transfer studies at high temperatures with pressure gradient for inlet flow Mach number of 0.45

NASA Technical Reports Server (NTRS)

Pedrosa, A. C. F.; Nagamatsu, H. T.; Hinckel, J. A.

1984-01-01

Heat transfer measurements were determined for a flat plate with and without pressure gradient for various free stream temperatures, wall temperature ratios, and Reynolds numbers for an inlet flow Mach number of 0.45, which is a representative inlet Mach number for gas turbine rotor blades. A shock tube generated the high temperature and pressure air flow, and a variable geometry test section was used to produce inlet flow Mach number of 0.45 and accelerate the flow over the plate to sonic velocity. Thin-film platinum heat gages recorded the local heat flux for laminar, transition, and turbulent boundary layers. The free stream temperatures varied from 611 R (339 K) to 3840 R (2133 K) for a T(w)/T(r,g) temperature ratio of 0.87 to 0.14. The Reynolds number over the heat gages varied from 3000 to 690,000. The experimental heat transfer data were correlated with laminar and turbulent boundary layer theories for the range of temperatures and Reynolds numbers and the transition phenomenon was examined.

Steady- and unsteady-pressure measurements on a supercritical-wing model with oscillating control surfaces at subsonic and transonic speeds

NASA Technical Reports Server (NTRS)

Sandford, M. C.; Ricketts, R. H.

1983-01-01

A high aspect ratio supercritical wing with oscillating control surfaces is described. The semispan wing model was instrumented with 252 static pressure orifices and 164 in situ dynamic pressure gages for studying the effects of control surface position and sinusoidal motion on steady and unsteady pressures. Results from the present test (the third in a series of tests on this model) were obtained in the Langley Transonic Dynamics Tunnel at Mach numbers of 0.60, 0.78, and 0.86 and are presented in tabular form.

Flight-determined lag of angle-of-attack and angle-of-sideslip sensors in the YF-12A airplane from analysis of dynamic maneuvers

NASA Technical Reports Server (NTRS)

Gilyard, G. B.; Belte, D.

1974-01-01

Magnitudes of lags in the pneumatic angle-of-attack and angle-of-sideslip sensor systems of the YF-12A airplane were determined for a variety of flight conditions by analyzing stability and control data. The three analysis techniques used are described. An apparent trend with Mach number for measurements from both of the differential-pressure sensors showed that the lag ranged from approximately 0.15 second at subsonic speed to 0.4 second at Mach 3. Because Mach number was closely related to altitude for the available flight data, the individual effects of Mach number and altitude on the lag could not be separated clearly. However, the results indicated the influence of factors other than simple pneumatic lag.

Measurements of Aerodynamic Heat Transfer and Boundary-Layer Transition on a 10 deg Cone in Free Flight at Supersonic Mach Numbers up to 5.9

NASA Technical Reports Server (NTRS)

Rumsey, Charles B.; Lee, Dorothy B.

1961-01-01

Measurements of aerodynamic heat transfer have been made at six stations on the 40-inch-long 10 deg. total-angle conical nose of a rocket- propelled model which was flight tested at Mach numbers up to 5.9. are presented for a range of local Mach number just outside the bound- ary layer on the cone from 1.57 to 5.50, and a range of local Reynolds number from 6.6 x 10(exp 6) to 55.2 x 10(exp 6) based on length from the nose tip.

Investigation of the Development of Laminar Boundary-Layer Instabilities Along a Cooled-Wall Hollow Cylinder at Mach Number 8

DTIC Science & Technology

1990-01-01

CVE M- a f~ r *C0Wn~fINAM SNOnuary 90re t WtaS aci Rep00ortfon 3-4I~af .au"guast89 f 4. TrIAftOSWTITL S. FUNDING NUMBERS InvestWiaton of the Development...over a range of pressure from 20 to 300 psia at Mach number 6, and 50 to 900 psia at Mach number 8, 6 with air supplied by the VKF main compressor ...plant. Stagnation temperatures sufficient to avoid air liquefaction in the test section (up to 1350° R ) are obtained through the use of a natural gas

Use of source distributions for evaluating theoretical aerodynamics of thin finite wings at supersonic speeds

NASA Technical Reports Server (NTRS)

Evvard, John C

1950-01-01

A series of publications on the source-distribution methods for evaluating the aerodynamics of thin wings at supersonic speeds is summarized, extended, and unified. Included in the first part are the deviations of: (a) the linearized partial-differential equation for unsteady flow at a substantially constant Mach number. b) The source-distribution solution for the perturbation-velocity potential that satisfies the boundary conditions of tangential flow at the surface and in the plane of the wing; and (c) the integral equation for determining the strength and the location of sources to describe the interaction effects (as represented by upwash) of the bottom and top wing surfaces through the region between the finite wing boundary and the foremost Mach wave. The second part deals with steady-state thin-wing problems. The third part of the report approximates the integral equation for unsteady upwash and includes a solution of approximate equation. Expressions are then derived to evaluate the load distributions for time-dependent finite-wing motions.

Measurements of fluctuating pressure in a rectangular cavity in transonic flow at high Reynolds numbers

NASA Technical Reports Server (NTRS)

Tracy, M. B.; Plentovich, E. B.; Chu, Julio

1992-01-01

An experiment was performed in the Langley 0.3 meter Transonic Cryogenic Tunnel to study the internal acoustic field generated by rectangular cavities in transonic and subsonic flows and to determine the effect of Reynolds number and angle of yaw on the field. The cavity was 11.25 in. long and 2.50 in. wide. The cavity depth was varied to obtain length-to-height (l/h) ratios of 4.40, 6.70, 12.67, and 20.00. Data were obtained for a free stream Mach number range from 0.20 to 0.90, a Reynolds number range from 2 x 10(exp 6) to 100 x 10(exp 6) per foot with a nearly constant boundary layer thickness, and for two angles of yaw of 0 and 15 degs. Results show that Reynolds number has little effect on the acoustic field in rectangular cavities at angle of yaw of 0 deg. Cavities with l/h = 4.40 and 6.70 generated tones at transonic speeds, whereas those with l/h = 20.00 did not. This trend agrees with data obtained previously at supersonic speeds. As Mach number decreased, the amplitude, and bandwidth of the tones changed. No tones appeared for Mach number = 0.20. For a cavity with l/h = 12.67, tones appeared at Mach number = 0.60, indicating a possible change in flow field type. Changes in acoustic spectra with angle of yaw varied with Reynolds number, Mach number, l/h ratios, and acoustic mode number.

Experiments on a smooth wall hypersonic boundary layer at Mach 6

NASA Astrophysics Data System (ADS)

Neeb, Dominik; Saile, Dominik; Gülhan, Ali

2018-04-01

The turbulent boundary layer along the surface of high-speed vehicles drives shear stress and heat flux. Although essential to the vehicle design, the understanding of compressible turbulent boundary layers at high Mach numbers is limited due to the lack of available data. This is particularly true if the surface is rough, which is typically the case for all technical surfaces. To validate a methodological approach, as initial step, smooth wall experiments were performed. A hypersonic turbulent boundary layer at Ma = 6 (Ma_e=5.4) along a 7{}° sharp cone model at low Reynolds numbers Re_{θ } ≈ 3000 was characterized. The mean velocities in the boundary layer were acquired by means of Pitot pressure and particle image velocimetry (PIV) measurements. Furthermore, the PIV data were used to extract turbulent intensities along the profile. The mean velocities in the boundary layer agree with numerical data, independent of the measurement technique. Based on the profile data, three different approaches to extract the skin friction velocity were applied and show favorable comparison to literature and numerical data. The extracted values were used for inner and outer scaling of the van Driest transformed velocity profiles which are in good agreement to incompressible theoretical data. Morkovin scaled turbulent intensities show ambiguous results compared to literature data which may be influenced by inflow turbulence level, particle lag and other measurement uncertainties.

Least-squares/parabolized Navier-Stokes procedure for optimizing hypersonic wind tunnel nozzles

NASA Technical Reports Server (NTRS)

Korte, John J.; Kumar, Ajay; Singh, D. J.; Grossman, B.

1991-01-01

A new procedure is demonstrated for optimizing hypersonic wind-tunnel-nozzle contours. The procedure couples a CFD computer code to an optimization algorithm, and is applied to both conical and contoured hypersonic nozzles for the purpose of determining an optimal set of parameters to describe the surface geometry. A design-objective function is specified based on the deviation from the desired test-section flow-field conditions. The objective function is minimized by optimizing the parameters used to describe the nozzle contour based on the solution to a nonlinear least-squares problem. The effect of the changes in the nozzle wall parameters are evaluated by computing the nozzle flow using the parabolized Navier-Stokes equations. The advantage of the new procedure is that it directly takes into account the displacement effect of the boundary layer on the wall contour. The new procedure provides a method for optimizing hypersonic nozzles of high Mach numbers which have been designed by classical procedures, but are shown to produce poor flow quality due to the large boundary layers present in the test section. The procedure is demonstrated by finding the optimum design parameters for a Mach 10 conical nozzle and a Mach 6 and a Mach 15 contoured nozzle.

Revisiting Turbulence Model Validation for High-Mach Number Axisymmetric Compression Corner Flows

NASA Technical Reports Server (NTRS)

Georgiadis, Nicholas J.; Rumsey, Christopher L.; Huang, George P.

2015-01-01

Two axisymmetric shock-wave/boundary-layer interaction (SWBLI) cases are used to benchmark one- and two-equation Reynolds-averaged Navier-Stokes (RANS) turbulence models. This validation exercise was executed in the philosophy of the NASA Turbulence Modeling Resource and the AIAA Turbulence Model Benchmarking Working Group. Both SWBLI cases are from the experiments of Kussoy and Horstman for axisymmetric compression corner geometries with SWBLI inducing flares of 20 and 30 degrees, respectively. The freestream Mach number was approximately 7. The RANS closures examined are the Spalart-Allmaras one-equation model and the Menter family of kappa - omega two equation models including the Baseline and Shear Stress Transport formulations. The Wind-US and CFL3D RANS solvers are employed to simulate the SWBLI cases. Comparisons of RANS solutions to experimental data are made for a boundary layer survey plane just upstream of the SWBLI region. In the SWBLI region, comparisons of surface pressure and heat transfer are made. The effects of inflow modeling strategy, grid resolution, grid orthogonality, turbulent Prandtl number, and code-to-code variations are also addressed.

Free-Flight Tests of 0.11-Scale North American F-100 Airplane Wings to Investigate the Possibility of Flutter in Transonic Speed Range at Varying Angles of Attack

NASA Technical Reports Server (NTRS)

O'Kelly, Burke R.

1954-01-01

Free-flight tests in the transonic speed range utilizing rocketpropelled models have been made on three pairs of 0.11-scale North American F-100 airplane wings having an aspect ratio of 3.47, a taper ratio of 0.308, 45 degree sweepback at the quarter-chord line, and thickness ratios of 31 and 5 percent to investigate the possibility of flutte r. Data from tests of two other rocket-propelled models which accidentally fluttered during a drag investigation of the North American F-100 airplane are also presented. The first set of wings (5 percent thick) was tested on a model which was disturbed in pitch by a moving tail and reached a maximum Mach number of 0.85. The wings encountered mild oscillations near the first - bending frequency at high lift coefficients. The second set of wings 9 percent thick was tested up to a maximum Mach number of 0.95 at (2) angles of attack provided by small rocket motors installed in the nose of the model. No oscillations resembling flutter were encountered during the coasting flight between separation from the booster and sustainer firing (Mach numbers from 0.86 to 0.82) or during the sustainer firing at accelerations of about 8g up to the maximum Mach number of the test (0.95). The third set of wings was similar to the first set and was tested up to a maximum Mach number of 1.24. A mild flutter at frequencies near the first-bending frequency of the wings was encountered between a Mach number of 1.15 and a Mach number of 1.06 during both accelerating and coasting flight. The two drag models, which were 0.ll-scale models of the North American F-100 airplane configuration, reached a maximum Mach number of 1.77. The wings of these models had bending and torsional frequencies which were 40 and 89 percent, respectively, of the calculated scaled frequencies of the full-scale 7-percent-thick wing. Both models experienced flutter of the same type as that experienced-by the third set of wings.

A Transonic Wind-Tunnel Investigation of the Performance and of the Static Stability and Control Characteristics of a Model of a Fighter-Type Airplane which Embodies Partial Body Indentation

NASA Technical Reports Server (NTRS)

Bielat, Ralph P.

1959-01-01

An investigation was conducted to obtain the aerodynamic characteristics of a model of a fighter-type airplane embodying partial body indentation. The wing had an aspect ratio of 4, taper ratio of 0.5, 35 deg sweepback of the 0.25-chord line, and a modified NACA 65A006 airfoil section at the root and a modified NACA 65A004 airfoil section at the tip. The fuselage has been indented in the region of the wing in order to obtain a favorable area distribution. The results reported herein consist of the performance and of the static longitudinal and lateral stability and control characteristics of the complete model. The Mach number range extended from 0.60 to 1.13, and the corresponding Reynolds number based on the wing mean aerodynamic chord varied from 1.77 x 10(exp 6) to 2.15 x 10(exp 6). The drag rise for both the cambered leading edge and symmetrical wing sections occurred at a Mach number of 0.95. Certain local modifications to the body which further improved the distribution of cross-sectional area gave additional reductions in drag at a Mach number of 1.00. The basic configuration indicated a mild pitch-up tendency at lift coefficients near 0.70 for the Mach number range from 0.80 to 0.90; however, the pitch-up instability may not be too objectionable on the basis of dynamic-stability considerations. The basic configuration indicated positive directional stability and positive effective dihedral through the angle-of-attack range and Mach number range with the exception of a region of negative effective dihedral at low lifts at Mach numbers of 1.00 and slightly above.

Aerodynamic pressure and heating-rate distributions in tile gaps around chine regions with pressure gradients at a Mach number of 6.6

NASA Technical Reports Server (NTRS)

Hunt, L. Roane; Notestine, Kristopher K.

1990-01-01

Surface and gap pressures and heating-rate distributions were obtained for simulated Thermal Protection System (TPS) tile arrays on the curved surface test apparatus of the Langley 8-Foot High Temperature Tunnel at Mach 6.6. The results indicated that the chine gap pressures varied inversely with gap width because larger gap widths allowed greater venting from the gap to the lower model side pressures. Lower gap pressures caused greater flow ingress from the surface and increased gap heating. Generally, gap heating was greater in the longitudinal gaps than in the circumferential gaps. Gap heating decreased with increasing gap depth. Circumferential gap heating at the mid-depth was generally less than about 10 percent of the external surface value. Gap heating was most severe at local T-gap junctions and tile-to-tile forward-facing steps that caused the greatest heating from flow impingement. The use of flow stoppers at discrete locations reduced heating from flow impingement. The use of flow stoppers at discrete locations reduced heating in most gaps but increased heating in others. Limited use of flow stoppers or gap filler in longitudinal gaps could reduce gap heating in open circumferential gaps in regions of high surface pressure gradients.

3-D Wizardry: Design in Papier-Mache, Plaster, and Foam.

ERIC Educational Resources Information Center

Wolfe, George

Papier-mache, plaster, and foam are inexpensive and versatile media for 3-dimensional classroom and studio art experiences. They can be used equally well by elementary, high school, or college students. Each medium has its own characteristic. Papier-mache is pliable but dries into a hard, firm surface that can be waterproofed. Plaster can be…

An investigation of the effects of the high maximum-thickness-to-chord ratio on the performance of nozzle guide vanes in a transonic planar cascade

NASA Astrophysics Data System (ADS)

Radmard, Rama

1993-03-01

The performance of turbine airfoils is usually predicted by empirical correlations, which however are inadequate for the case of airfoils with maximum thickness to chord ratio (MTCR) higher than 25 percent. Studies were conducted to create a data base from which the performance of turbine airfoils with a MTCR higher than 25 percent could be predicted. A planar cascade consisting of four airfoils was constructed to allow the investigation of the effect of the MTCR on the airfoil performance. Three airfoil sets with MTCR of 15.2 percent (baseline), 26.6 percent, and 48.2 percent were used. Measurements included surface Mach number distributions for the baseline airfoil, total pressure loss coefficients, and deviation angles for isentropic exit Mach numbers of 0.7 (design), 0.9, and 1.1. The effect of varying the inlet boundary layer thickness and free-stream turbulence level was also examined. The results showed that the 26.6 percent airfoil produced lower losses as predicted by the Kacker and Okapuu (1982) correlation. The introduction of turbulence produced a significant redistribution of losses in the exit plane. The secondary loss decreased as the leading edge diameter was increased. Except for the baseline blade where high under-turning in exit flow angle was observed, the airfoils showed a decrease in over-turning with increasing exit Mach number, as predicted by Ainley and Mathieson (1951).

Impact of Shock Front Rippling and Self-reformation on the Electron Dynamics at Low-Mach-number Shocks

NASA Astrophysics Data System (ADS)

Yang, Zhongwei; Lu, Quanming; Liu, Ying D.; Wang, Rui

2018-04-01

Electron dynamics at low-Mach-number collisionless shocks are investigated by using two-dimensional electromagnetic particle-in-cell simulations with various shock normal angles. We found: (1) The reflected ions and incident electrons at the shock front provide an effective mechanism for the quasi-electrostatic wave generation due to the charge-separation. A fraction of incident electrons can be effectively trapped and accelerated at the leading edge of the shock foot. (2) At quasi-perpendicular shocks, the electron trapping and reflection is nonuniform due to the shock rippling along the shock surface and is more likely to take place at some locations accompanied by intense reflected ion-beams. The electron trapping process has a periodical evolution over time due to the shock front self-reformation, which is controlled by ion dynamics. Thus, this is a cross-scale coupling phenomenon. (3) At quasi-parallel shocks, reflected ions can travel far back upstream. Consequently, quasi-electrostatic waves can be excited in the shock transition and the foreshock region. The electron trajectory analysis shows these waves can trap electrons at the foot region and reflect a fraction of them far back upstream. Simulation runs in this paper indicate that the micro-turbulence at the shock foot can provide a possible scenario for producing the reflected electron beam, which is a basic condition for the type II radio burst emission at low-Mach-number interplanetary shocks driven by Coronal Mass Ejections (CMEs).

The Influence of Sweep on the Aerodynamic Loading of an Oscillating NACA0012 Airfoil. Volume 2: Data Report

NASA Technical Reports Server (NTRS)

St.hilaire, A. O.; Carta, F. O.

1979-01-01

The effect of sweep on the dynamic response of the NACA 0012 airfoil was investigated. Unsteady chordwise distributed pressure data were obtained from a tunnel spanning wing equipped with 21 single surface transducers (13 on the suction side and 8 on the pressure side of the airfoil). The pressure data were obtained at pitching amplitudes of 8 and 10 degrees over a tunnel Mach number range of 0.10 to 0.46 and a pitching frequency range of 2.5 to 10.6 cycles per second. The wing was oscillated in the unswept and swept positions about the quarter-chord pivot axis relative to mean incidence angle settings of 0, 9, 12, and 15 degrees. A compilation of all the response data obtained during the test program is presented. These data are in the form of normal force, chord force, lift force, pressure drag, and moment hysteresis loops derived from chordwise integrations of the unsteady pressure distributions. The hysteresis loops are organized in two main sections. In the first section, the loop data are arranged to show the effect of sweep (lambda = 0 and 30 deg) for all available combinations of mean incidence angle, pitching amplitude, reduced frequency, and chordwise Mach number. The second section shows the effect of chordwise Mach number (MC = 0.30 and MC = 0.40) on the swept wing response for all available combinations of mean incidence angle, pitching amplitude, and reduced frequency.

Flutter of a Low-Aspect-Ratio Rectangular Wing

NASA Technical Reports Server (NTRS)

Cole, Stanley R.

1989-01-01

A flutter test of a low-aspect-ratio rectangular wing was conducted in the Langley Transonic Dynamics Tunnel (TDT). The model used in this flutter test consisted of a rigid wing mounted to the wind-tunnel wall by a flexible, rectangular beam. The flexible support shaft was connected to the wing root and was cantilever mounted to the wind-tunnel wall. The wing had an aspect ratio of 1.5 based on the wing semispan and an NACA 64A010 airfoil shape. The flutter boundary of the model was determined for a Mach number range of 0.5 to 0.97. The shape of the transonic flutter boundary was determined. Actual flutter points were obtained on both the subsonic and supersonic sides of the flutter bucket. The model exhibited a deep transonic flutter bucket over a narrow range of Mach number. At some Mach numbers, the flutter conditions were extrapolated using a subcritical response technique. In addition to the basic configuration, modifications were made to the model structure such that the first bending frequency was changed without significantly affecting the first torsion frequency. The experiment showed that increasing the bending stiffness of the model support shaft through these modifications lowered the flutter dynamic pressure. Flutter analysis was conducted for the basic model as a comparison with the experimental results. This flutter analysis was conducted with subsonic lifting-surface (kernel function) aerodynamics using the k method for the flutter solution.

Wind-Tunnel Results of Advanced High-Speed Propellers at Takeoff, Climb, and Landing Mach Numbers

NASA Technical Reports Server (NTRS)

Stefko, George L.; Jeracki, Robert J.

1985-01-01

Low-speed wind-tunnel performance tests of two advanced propellers have been completed at the NASA Lewis Research Center as part of the NASA Advanced Turboprop Program. The 62.2 cm (24.5 in.) diameter adjustable-pitch models were tested at Mach numbers typical of takeoff, initial climbout, and landing speeds (i.e., from Mach 0.10 to 0.34) at zero angle of attack in the NASA Lewis 10 by 10 Foot Supersonic Wind Tunnel. Both models had eight blades and a cruise-design-point operating condition of Mach 0.80, and 10.668 km (35,000 ft) I.S.A. altitude, a 243.8 m/s (800 ft/sec) tip speed, and a high power loading of 301 kW/sq m (37.5 shp/sq ft). Each model had its own integrally designed area-ruled spinner, but used the same specially contoured nacelle. These features reduced blade-section Mach numbers and relieved blade-root choking at the cruise condition. No adverse or unusual low-speed operating conditions were found during the test with either the straight blade SR-2 or the 45 deg swept SR-3 propeller. Typical efficiencies of the straight and 45 deg swept propellers were 50.2 and 54.9 percent, respectively, at a takeoff condition of Mach 0.20 and 53.7 and 59.1 percent, respectively, at a climb condition of Mach 0.34.

Investigation at Mach Numbers of 0.20 to 3.50 of Blended Wing-Body Combinations of Sonic Design with Diamond, Delta, and Arrow Plan Forms

NASA Technical Reports Server (NTRS)

Holdaway, George H.; Mellenthin, Jack A.

1960-01-01

The models had aspect-ratio-2 diamond, delta, and arrow wings with the leading edges swept 45.00 deg, 59.04 deg, and 70.82 deg, respectively. The wing sections were computed by varying the section shape along with the body radii (blending process) to match the prescribed area distribution and wing plan form. The wing sections had an average value of maximum thickness ratio of about 4 percent of the local chords in a streamwise direction. The models were tested with transition fixed at Reynolds numbers of about 4,000,000 to 9,000,0000, based on the mean aerodynamic chord of the wings. The effect of varying Reynolds number was checked at both subsonic and supersonic speeds. The diamond model was superior to the other plan forms at transonic speeds ((L/D)max = 11.00 to 9.52) because of its higher lift-curve slope and near optimum wave drag due to the blending process. For the wing thickness tested with the diamond model, the marked body and wing contouring required for transonic conditions resulted in a large wave-drag penalty at the higher supersonic Mach numbers where the leading and trailing edges of the wing were supersonic. Because of the low sweep of the trailing edge of the delta model, this configuration was less adaptable to the blending process. Removing a body bump prescribed by the Mach number 1.00 design resulted in a good supersonic design. This delta model with 10 percent less volume was superior to the other plan forms at Mach numbers of 1.55 to 2.35 ((L/D)max = 8.65 to 7.24), but it and the arrow model were equally good at Mach numbers of 2.50 to 3.50 ((L/D)max - 6.85 to O.39). At transonic speeds the arrow model was inferior because of the reduced lift-curve slope associated with its increased sweep and also because of the wing base drag. The wing base-drag coefficients of the arrow model based on the wing planform area decreased from a peak value of 0.0029 at Mach number 1.55 to 0.0003 at Mach number 3.50. Linear supersonic theory was satisfactory for predicting the aerodynamic trends at Mach numbers from 1.55 to 3.50 of lift-curve slope, wave drag, drag due to lift, aerodynamic-center location, and maximum lift-drag ratios for each of the models.

Experimental Hypersonic Aerodynamic Characteristics of the Space Shuttle Orbiter for a Range of Damage Scenarios

NASA Technical Reports Server (NTRS)

Brauckman, Gregory J.; Scallion, William I.

2003-01-01

Aerodynamic tests in support of the Columbia accident investigation were conducted in two hypersonic wind tunnels at the NASA Langley Research Center, the 20-Inch Mach 6 Air Tunnel and the 20-Inch Mach 6 CF4 Tunnel. The primary purpose of these tests was to measure the forces and moments generated by a variety of outer mold line alterations (damage scenarios) using 0.0075-scale models of the Space Shuttle Orbiter (approximately 10 inches in length). Simultaneously acquired global heat transfer mappings were obtained for a majority of the configurations tested. Test parameters include angles of attack from 38 to 42 deg, unit Reynolds numbers from 0.26 to 3.0 x10^6 per foot, and normal shock density ratios of 5 (Mach 6 air) and 12 (Mach 6 CF4). The damage scenarios evaluated included asymmetric boundary layer transition, gouges in the windward surface acreage thermal protection system tiles, wing leading edge damage (partially and fully missing reinforced carbon-carbon (RCC) panels), holes through the wing from the windward surface to the leeside, deformation of the wing windward surface, and main landing gear door and/or gear deployment. The aerodynamic data were compared to the magnitudes and directions observed in flight, and the heating images were evaluated in terms of the location of the generated disturbances and how these disturbance might relate to the response of discrete gages on the Columbia Orbiter vehicle during entry. The measured aerodynamic increments were generally small in magnitude, as were the flight-derived values during most of the entry. Asymmetric boundary layer transition (ABLT) results were consistent with the flight-derived Shuttle ABLT model, but not with the observed flight trends for STS-107. The partially missing leading edge panel results best matched both the early aerodynamic and heating trends observed in flight. A progressive damage scenario is presented that qualitatively matches the flight observations for the full entry.

Altitude Performance Characteristics of Tail-pipe Burner with Variable-area Exhaust Nozzle

NASA Technical Reports Server (NTRS)

Jansen, Emmert T; Thorman, H Carl

1950-01-01

An investigation was conducted in the NACA Lewis altitude wind tunnel to determine effect of altitude and flight Mach number on performance of tail-pipe burner equipped with variable-area exhaust nozzle and installed on full-scale turbojet engine. At a given flight Mach number, with constant exhaust-gas and turbine-outlet temperatures, increasing altitude lowered the tail-pipe combustion efficiency and raised the specific fuel consumption while the augmented thrust ratio remained approximately constant. At a given altitude, increasing flight Mach number raised the combustion efficiency and augmented thrust ratio and lowered the specific fuel consumption.

Vehicle integration effects on hypersonic waveriders. M.S. Thesis - George Washington Univ.

NASA Technical Reports Server (NTRS)

Cockrell, Charles Edward, Jr.

1994-01-01

The integration of a class of hypersonic high-lift configurations known as waveriders into hypersonic cruise vehicles was evaluated. Waveriders offer advantages in aerodynamic performance and propulsion/airframe integration (PAI) characteristics over conventional hypersonic shapes. A wind-tunnel model was developed which integrates realistic vehicle components with two waverider shapes, referred to as the 'straight-wing' and 'cranked-wing' shapes. Both shapes were conical-flow-derived waveriders at a design Mach number of 4.0. The cranked-wing shape was designed to provide advantages in subsonic performance and directional stability over conventional waveriders. Experimental data and limited computational fluid dynamics (CFD) predictions were obtained over a Mach number range of 2.3 to 4.63 at a Reynolds number of 2.0x10(exp 6) per foot. The CFD predictions and flow visualization data confirmed the shock attachment characteristics of the baseline waverider shapes and illustrated the waverider flow-field properties. Both CFD predictions and experimental data showed that no significant performance degradations occur at off-design Mach numbers for the waverider shapes and the integrated configurations. The experimental data showed that the effects of adding a realistic canopy were minimal. The effects of adding engine components were to increase the drag and thus degrade the aerodynamic performance of the configuration. A significant degradation in aerodynamic performance was observed when 0 degree control surfaces were added to close the blunt base of the waverider to a sharp trailing edge. A comparison of the fully-integrated waverider models to the baseline shapes showed that the performance was significantly degraded when all of the components were added to the waveriders. The fully-integrated configurations studied here do not offer significant performance advantages over conventional hypersonic vehicles, but still offer advantages in air-breathing propulsion integration. Additionally, areas are identified in this study where improvements could be made to enhance the performance. Both fully-integrated configurations are longitudinally unstable over the Mach number range studied for unpowered conditions. The cranked-wing fully-integrated configuration provided significantly better lateral-directional stability characteristics than the straight-wing configuration.

A critical shock mach number for particle acceleration in the absence of pre-existing cosmic rays: M=√5

DOE Office of Scientific and Technical Information (OSTI.GOV)

Vink, Jacco; Yamazaki, Ryo, E-mail: j.vink@uva.nl

2014-01-10

It is shown that, under some generic assumptions, shocks cannot accelerate particles unless the overall shock Mach number exceeds a critical value M>√5. The reason is that for M≤√5 the work done to compress the flow in a particle precursor requires more enthalpy flux than the system can sustain. This lower limit applies to situations without significant magnetic field pressure. In case that the magnetic field pressure dominates the pressure in the unshocked medium, i.e., for low plasma beta, the resistivity of the magnetic field makes it even more difficult to fulfill the energetic requirements for the formation of shockmore » with an accelerated particle precursor and associated compression of the upstream plasma. We illustrate the effects of magnetic fields for the extreme situation of a purely perpendicular magnetic field configuration with plasma beta β = 0, which gives a minimum Mach number of M = 5/2. The situation becomes more complex, if we incorporate the effects of pre-existing cosmic rays, indicating that the additional degree of freedom allows for less strict Mach number limits on acceleration. We discuss the implications of this result for low Mach number shock acceleration as found in solar system shocks, and shocks in clusters of galaxies.« less

The Aerodynamic Characteristics in Pitch of a 1/15-Scale Model of the Grumman F11F-1 Airplane at Mach Numbers of 1.41, 1.61, and 2.01, TED No. NACA DE 390

NASA Technical Reports Server (NTRS)

Driver, Cornelius

1956-01-01

Tests have been made in the Langley 4- by 4-foot supersonic pressure tunnel at Mach numbers of 1.41, 1.61, and 2.01 to determine the static longitudinal stability and control characteristics of various arrangements of the Grumman F11F-1 airplane. Tests were made of the complete model and various combinations of its component parts and, in addition, the effects of various body modifications, a revised vertical tail, and wing fences on the longitudinal characteristics were determined. The results indicate that for a horizontal-tail incidence of -10 deg the trim lift coefficient varied from 0.29 at a Mach number of 1.61 to 0.23 at a Mach number of 2.01 with a corresponding decrease in lift-drag trim from 3.72 to 3.15. Stick-position instability was indicated in the low-supersonic-speed range. A photographic-type nose modification resulted in slightly higher values of minimum drag coefficient but did not significantly affect the static stability or lift-curve slope. The minimum drag coefficient for the complete model with the production nose remained essentially constant at 0.047 throughout the Mach number range investigated.

Design and Navier-Stokes analysis of hypersonic wind tunnel nozzles. M.S. Thesis

NASA Technical Reports Server (NTRS)

Benton, James R.

1989-01-01

Four hypersonic wind tunnel nozzles ranging in Mach number from 6 to 17 are designed with the method of characteristics and boundary layer approach (MOC/BL) and analyzed with a Navier-Stokes solver. Limitations of the MOC/BL approach when applied to thick high speed boundary layers with non-zero normal pressure gradients are investigated. Working gases include ideal air, thermally perfect nitrogen and virial CF4. Agreement between the design conditions and Navier-Stokes solutions for ideal air at Mach 6 is good. Thermally perfect nitrogen showed poor agreement at Mach 13.5 and Mach 17. Navier-Stokes solutions for CF4 are not obtained, but comparison of the effects of low gamma to those of high Mach number suggests that the Navier-Stokes solution would not compare well with design.

Design of a very-low-bleed Mach 2.5 mixed-compression inlet with 45 percent internal contraction

NASA Technical Reports Server (NTRS)

Wasserbauer, J. F.; Shaw, R. J.; Neumann, H. E.

1975-01-01

A full-scale, mixed-compression inlet was designed for operation with the TF30-P-3 turbofan engine and tested at Mach numbers of 2.5 and 2.0. The two-cone axisymmetric inlet had minimum internal contraction consistent with high total pressure recovery and low cowl drag. At Mach 2.5, inlet recovery was 0.906 with only 0.021 centerbody bleed mass-flow ratio and no cowl bleed. Increased centerbody bleed gave a maximum inlet unstart angle of attack of 6.85 deg. At Mach 2.0, inlet recovery was 0.94 with only 0.014 centerbody bleed mass-flow ratio and no cowl bleed. Inlet performance and angle-of-attack tolerance is presented for operation at Mach numbers of 2.5 and 2.0.

In-flight Compressible Turbulent Boundary Layer Measurements on a Hollow Cylinder at a Mach Number of 3.0

NASA Technical Reports Server (NTRS)

Quinn, R. D.; Gong, L.

1978-01-01

Skin temperatures, shearing forces, surface static pressures, and boundary layer pitot pressures and total temperatures were measured on a hollow cylinder 3.04 meters long and 0.437 meter in diameter mounted beneath the fuselage of the YF-12A airplane. The data were obtained at a nominal free stream Mach number of 3.0 and at wall-to-recovery temperature ratios of 0.66 to 0.91. The free stream Reynolds number had a minimal value of 4.2 million per meter. Heat transfer coefficients and skin friction coefficients were derived from skin temperature time histories and shear force measurements, respectively. Boundary layer velocity profiles were derived from pitot pressure measurements, and a Reynolds analogy factor of 1.11 was obtained from the measured heat transfer and skin friction data. The skin friction coefficients predicted by the theory of van Driest were in excellent agreement with the measurements. Theoretical heat transfer coefficients, in the form of Stanton numbers calculated by using a modified Reynolds analogy between skin friction and heat transfer, were compared with measured values. The measured velocity profiles were compared to Coles' incompressible law-of-the-wall profile.

An evaluation of three two-dimensional computational fluid dynamics codes including low Reynolds numbers and transonic Mach numbers

NASA Technical Reports Server (NTRS)

Hicks, Raymond M.; Cliff, Susan E.

1991-01-01

Full-potential, Euler, and Navier-Stokes computational fluid dynamics (CFD) codes were evaluated for use in analyzing the flow field about airfoils sections operating at Mach numbers from 0.20 to 0.60 and Reynolds numbers from 500,000 to 2,000,000. The potential code (LBAUER) includes weakly coupled integral boundary layer equations for laminar and turbulent flow with simple transition and separation models. The Navier-Stokes code (ARC2D) uses the thin-layer formulation of the Reynolds-averaged equations with an algebraic turbulence model. The Euler code (ISES) includes strongly coupled integral boundary layer equations and advanced transition and separation calculations with the capability to model laminar separation bubbles and limited zones of turbulent separation. The best experiment/CFD correlation was obtained with the Euler code because its boundary layer equations model the physics of the flow better than the other two codes. An unusual reversal of boundary layer separation with increasing angle of attack, following initial shock formation on the upper surface of the airfoil, was found in the experiment data. This phenomenon was not predicted by the CFD codes evaluated.

Aerodynamic tests and analysis of a turbojet-boosted launch vehicle concept (spacejet) over a Mach number range of 1.50 to 2.86. [Langley Unitary Plan Wind Tunnel Tests

NASA Technical Reports Server (NTRS)

Riebe, G. D.; Small, W. J.; Morris, O. A.

1981-01-01

Results from analytical and experimental studies of the aerodynamic characteristics of a turbojet-boosted launch vehicle concept through a Mach number range of 1.50 to 2.86 are presented. The vehicle consists of a winged orbiter utilizing an area-ruled axisymmetric body and two winged turbojet boosters mounted underneath the orbiter wing. Drag characteristics near zero lift were of prime interest. Force measurements and flow visualization techniques were employed. Estimates from wave drag theory, supersonic lifting surface theory, and impact theory are compared with data and indicate the ability of these theories to adequately predict the aerodynamic characteristics of the vehicle. Despite the existence of multiple wings and bodies in close proximity to each other, no large scale effects of boundary layer separation on drag or lift could be discerned. Total drag levels were, however, sensitive to booster locations.

Results of tests OA26 and IA16 in the NASA/ARC 3.5-foot hypersonic wind tunnel on an 0.015-scale model (36-OTS) of the space shuttle configuration 140A/B to obtain pressures for venting analysis

NASA Technical Reports Server (NTRS)

Spangler, R. H.; Thornton, D. E.; Polek, T. E.

1974-01-01

Tests were conducted, from November 15 to December 4, 1973, to obtain surface pressure data on an 0.015-scale replica of the Space Shuttle Vehicle 4. Data were obtained at Mach numbers of 5.3, 7.4, and 10.3, to support the venting analysis for both launch and entry conditions. These tests were the final tests in a series covering a Mach number range from 0.6 to 10.3. The model was instrumented with pressure orifices in the vicinity of the cargo bay door hinge and parting lines, and on the side of the fuselage at the crew compartment, and below the orbital maneuvering system pods at the aft compartment. The model was tested at angles of attack and sideslip consistent with expected divergencies from the nominal trajectory.

2-D Air-Breathing Lightcraft Engine Experiments in Hypersonic Conditions

NASA Astrophysics Data System (ADS)

Salvador, Israel I.; Myrabo, Leik N.; Minucci, Marco A. S.; de Oliveira, Antonio C.; Toro, Paulo G. P.; Chanes, José B.; Rego, Israel S.

2011-11-01

Experiments were performed with a 2-D, repetitively-pulsed (RP) laser Lightcraft model in hypersonic flow conditions. The main objective was the feasibility analysis for impulse generation with repetitively-pulsed air-breathing laser Lightcraft engines at hypersonic speeds. The future application of interest for this basic research endeavor is the laser launch of pico-, nano-, and micro-satellites (i.e., 0.1-100 kg payloads) into Low-Earth-Orbit, at low-cost and on-demand. The laser propulsion experiments employed a Hypersonic Shock Tunnel integrated with twin gigawatt pulsed Lumonics 620-TEA CO2 lasers (˜ 1 μs pulses), to produce the required test conditions. This hypersonic campaign was carried out at nominal Mach numbers ranging from 6 to 10. Time-dependent surface pressure distributions were recorded together with Schlieren movies of the flow field structure resulting from laser energy deposition. Results indicated laser-induced pressure increases of 0.7-0.9 bar with laser pulse energies of ˜ 170 J, on off-shroud induced breakdown condition, and Mach number of 7.

Numerical Simulation of Supersonic Gap Flow

PubMed Central

Jing, Xu; Haiming, Huang; Guo, Huang; Song, Mo

2015-01-01

Various gaps in the surface of the supersonic aircraft have a significant effect on airflows. In order to predict the effects of attack angle, Mach number and width-to-depth ratio of gap on the local aerodynamic heating environment of supersonic flow, two-dimensional compressible Navier-Stokes equations are solved by the finite volume method, where convective flux of space term adopts the Roe format, and discretization of time term is achieved by 5-step Runge-Kutta algorithm. The numerical results reveal that the heat flux ratio is U-shaped distribution on the gap wall and maximum at the windward corner of the gap. The heat flux ratio decreases as the gap depth and Mach number increase, however, it increases as the attack angle increases. In addition, it is important to find that chamfer in the windward corner can effectively reduce gap effect coefficient. The study will be helpful for the design of the thermal protection system in reentry vehicles. PMID:25635395

The dynamic-response characteristics of a 35 degree swept-wing airplane as determined from flight measurements

NASA Technical Reports Server (NTRS)

Triplett, William C; Brown, Stuart C; Smith, G Allan

1955-01-01

The longitudinal and lateral-directional dynamic-response characteristics of a 35 degree swept-wing fighter-type airplane determined from flight measurements are presented and compared with predictions based on theoretical studies and wind-tunnel data. Flights were made at an altitude of 35,000 feet covering the Mach number range of 0.50 to 1.04. A limited amount of lateral-directional data were also obtained at 10,000 feet. The flight consisted essentially of recording transient responses to pilot-applied pulsed motions of each of the three primary control surfaces. These transient data were converted into frequency-response form by means of the Fourier transformation and compared with predicted responses calculated from the basic equations. Experimentally determined transfer functions were used for the evaluation of the stability derivatives that have the greatest effect on the dynamic response of the airplane. The values of these derivatives, in most cases, agreed favorably with predictions over the Mach number range of the test.

Space shuttle orbiter rear mounted reaction control system jet interaction study. [hypersonic wind tunnel tests

NASA Technical Reports Server (NTRS)

Rausch, J. R.

1977-01-01

The effect of interaction between the reaction control system (RCS) jets and the flow over the space shuttle orbiter in the atmosphere was investigated in the NASA Langley 31-inch continuous flow hypersonic tunnel at a nominal Mach number of 10.3 and in the AEDC continuous flow hypersonic tunnel B at a nominal Mach number of 6, using 0.01 and .0125 scale force models with aft RCS nozzles mounted both on the model and on the sting of the force model balance. The data show that RCS nozzle exit momentum ratio is the primary correlating parameter for effects where the plume impinges on an adjacent surface and mass flow ratio is the parameter when the plume interaction is primarily with the external stream. An analytic model of aft mounted RCS units was developed in which the total reaction control moments are the sum of thrust, impingement, interaction, and cross-coupling terms.

Longitudinal aerodynamic characteristics of a wing-winglet model designed at M = 0.8, C sub L = 0.4 using linear aerodynamic theory

NASA Technical Reports Server (NTRS)

Kuhlman, J. M.

1983-01-01

Wind tunnel test results have been presented herein for a subsonic transport type wing fitted with winglets. Wind planform was chosen to be representative of wings used on current jet transport aircraft, while wing and winglet camber surfaces were designed using two different linear aerodynamic design methods. The purpose of the wind tunnel investigation was to determine the effectiveness of these linear aerodynamic design computer codes in designing a non-planar transport configuration which would cruise efficiently. The design lift coefficient was chosen to be 0.4, at a design Mach number of 0.8. Force and limited pressure data were obtained for the basic wing, and for the wing fitted with the two different winglet designs, at Mach numbers of 0.60, 0.70, 0.75 and 0.80 over an angle of attack range of -2 to +6 degrees, at zero sideslip. The data have been presented without analysis to expedite publication.