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At least 307 records · Page 17

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NASA NTRS1991

Analysis of Fluctuating Static Pressure Measurements in a Large High Reynolds Number Transonic Cryogenic Wind Tunnel

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.

Igoe, William B.↗

NASA NTRS1993

Finite element calculations of transonic flutter in cascades

Aeroelastic stability and response of turbomachinery cascades operating in the transonic regime are explored using a new computational approach. The two-dimensional unsteady Euler equations are used to model the transonic cascade flows, and the cascade blades are modeled as thin plates. To minimize the discretization incompatibility, both the fluid and the solid domains are discretized using finite element procedures. The fluid and the structural models are coupled at the element level by introducing a channel theory, which is analogous to a strip theory in the classical sense. The system of space-discretized equations is marched forward in time using an explicit four-stage Runge-Kutta scheme. Results of free vibration analyses of typical blades are presented and compared to published finite element and Rayleigh-Ritz results. Aeroelastic stability and response characteristics of several transonic cascades are investigated using the new quasi-3D model. These results demonstrate the present 3D finite element blade model is capable of capturing aeroelastic instabilities involving higher modes, including (plate-type) camber bending modes.

Hsiao, Chingteng↗

NASA NTRS1956

On slender-body theory and the area rule at transonic speeds

The basic ideas of the slender-body approximation have been applied to the nonlinear transonic-flow equation for the velocity potential in order to obtain some of the essential features of slender-body theory at transonic speeds. The results of the investigation are presented from a unified point of view which demonstrates the similarity of slender-body solutions in the various Mach number ranges. The transonic area rule and some conditions concerning its validity follow from the analysis.

FLOW, SUPERSONIC↗

NASA NTRS1953

A correlation by means of transonic similarity rules of the experimentally determined characteristics of 18 cambered wings of rectangular plan form

The effects of one type of camber on the aerodynamic characteristics of rectangular wings at high subsonic and transonic speeds have been studied by applying the transonic similarity rules to the correlation of experimental data for a series of 18 cambered wings having NACA 63A2XX and 63A4XX sections, aspect ratios from 1 to 4, and thicknesses from 4 to 8 percent. The data were obtained by use of a transonic bump over a Mach number range of 0.6 to 1.1.

PROFILES - COMPLETE WINGS↗

NASA NTRS1955

Transonic Flow Past Cone Cylinders

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.

Solomon, George E↗

NASA NTRS1957

On Slender-Body Theory at Transonic Speeds

Harder, Keith C↗

NASA NTRS1952

Transonic flow past a wedge profile with detached bow wave

A theoretical study has been made of the aerodynamic characteristics at zero angle of attack of a thin, doubly symmetrical double-wedge profile in the range of supersonic flight speed in which the bow wave is detached. The analysis utilizes the equations of the transonic small-disturbance theory and involves no assumptions beyond those implicit in this theory. The mixed flow about the front half of the profile is calculated by relaxation solution of boundary conditions along the shock polar and sonic line. The purely subsonic flow about the rear of the profile is found by means of the method of characteristics specialized to the transonic small-disturbance theory. Complete calculations were made for four values of the transonic similarity parameter. These were found sufficient to bridge the gap between the previous results of Guderley and Yoshihara at a Mach number of 1 and the results which are readily obtained when the bow wave is attached and the flow is completely supersonic.

Vincenti, Walter G↗

NASA NTRS1952

On a solution of the nonlinear differential equation for transonic flow past a wave-shaped wall

The Prandtl-Busemann small-perturbation method is utilized to obtain the flow of a compressible fluid past an infinitely long wave-shaped wall. When the essential assumption for transonic flow (that all Mach numbers in the region of flow are nearly unity) is introduced, the expression for the velocity potential takes the form of a power series in the transonic similarity parameter. On the basis of this form of the solution, an attempt is made to solve the nonlinear differential equation for transonic flow past the wavy wall. The analysis utilized exhibits clearly the difficulties inherent in nonlinear-flow problems.

Kaplan, Carl↗

NASA NTRS1952

The Langley Annular Transonic Tunnel

Report describes the development of the Langley annular transonic tunnel, a facility in which test Mach numbers from 0.6 to slightly over 1.0 are achieved by rotating the test model in an annular passage between two concentric cylinders. Data obtained for two-dimensional airfoil models in the Langley annular transonic tunnel at subsonic and sonic speeds are shown to be in reasonable agreement with experimental data from other sources and with theory when comparisons are made for nonlifting conditions or for equal normal-force coefficients rather than for equal angles of attack. The trends of pressure distributions obtained from measurements in the Langley annular transonic tunnel are consistent with distributions calculated for Prandtl-Meyer flow.

Habel, Louis W↗

NASA NTRS1992

Results of wind tunnel tests of an ASRM configured 0.03 scale Space Shuttle integrated vehicle model (47-OTS) in the AEDC 16-foot transonic wind tunnel, volume 2

An experimental Aerodynamic and Aero-Acoustic loads data base was obtained at transonic Mach numbers for the Space Shuttle Launch Vehicle configured with the ASRM Solid Rocket Boosters as an increment to the current flight configuration (RSRB). These data were obtained during transonic wind tunnel tests (IA 613A) conducted in the Arnold Engineering Development Center 16-Foot transonic propulsion wind tunnel from March 27, 1991 through April 12, 1991. This test is the first of a series of two tests covering the Mach range from 0.6 to 3.5. Steady state surface static and fluctuating pressure distributions over the Orbiter, External Tank and Solid Rocket Boosters of the Shuttle Integrated Vehicle were measured. Total Orbiter forces, Wing forces and Elevon hinge moments were directly measured as well from force balances. Two configurations of Solid Rocket Boosters were tested, the Redesigned Solid Rocket Booster (RSRB) and the Advanced Solid Rocket Motor (ASRM). The effects of the position (i.e., top, bottom, top and bottom) of the Integrated Electronics Assembly (IEA) box, mounted on the SRB attach ring, were obtained on the ASRM configured model. These data were obtained with and without Solid Plume Simulators which, when used, matched as close as possible the flight derived pressures on the Orbiter and External Tank base. Data were obtained at Mach numbers ranging from 0.6 to 1.55 at a Unit Reynolds Number of 2.5 million per foot through model angles of attack from -8 to +4 degrees at sideslip angles of 0, +4 and -4 degrees.

Marroquin, J.↗

NASA NTRS1992

Results of wind tunnel tests of an ASRM configured 0.03 scale Space Shuttle integrated vehicle model (47-OTS) in the AEDC 16-foot Transonic wind tunnel (IA613A), volume 1

An experimental Aerodynamic and Aero-Acoustic loads data base was obtained at transonic Mach numbers for the Space Shuttle Launch Vehicle configured with the ASRM Solid Rocket Boosters as an increment to the current flight configuration (RSRB). These data were obtained during transonic wind tunnel tests (IA 613A) conducted in the Arnold Engineering Development Center 16-Foot transonic propulsion wind tunnel from March 27, 1991 through April 12, 1991. This test is the first of a series of two tests covering the Mach range from 0.6 to 3.5. Steady state surface static and fluctuating pressure distributions over the Orbiter, External Tank and Solid Rocket Boosters of the Shuttle Integrated Vehicle were measured. Total Orbiter forces, Wing forces and Elevon hinge moments were directly measured as well from force balances. Two configurations of Solid Rocket Boosters were tested, the Redesigned Solid Rocket Booster (RSRB) and the Advanced Solid Rocket Motor (ASRM). The effects of the position (i.e. top, bottom, top and bottom) of the Integrated Electronics Assembly (IEA) box, mounted on the SRB attach ring, were obtained on the ASRM configured model. These data were obtained with and without Solid Plume Simulators which, when used, matched as close as possible the flight derived pressures on the Orbiter and External Tank base. Data were obtained at Mach numbers ranging from 0.6 to 1.55 at a Unit Reynolds Number of 2.5 million per foot through model angles of attack from -8 to +4 degrees at sideslip angles of 0, +4 and -4 degrees.

Marroquin, J.↗

NASA NTRS1993

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

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.

Erickson, Gary E.↗

NASA NTRS1994

Experimental studies of transonic flow field near a longitudinally slotted wind tunnel wall

The results of detailed parametric experiments are presented for the near-wall flow field of a longitudinally slotted transonic wind tunnel. Existing data are reevaluated and new data obtained in the Langley 6- by 19-inch Transonic Wind Tunnel are presented and analyzed. In the experiments, researchers systematically investigate many pertinent wall-geometry variables such as the wall openness and the number of slots along with the free stream Mach number and model angle of attack. Flow field surveys on the plane passing through the centerline of the slot were conducted and are presented. The effects of viscosity on the slot flow are considered in the analysis. The present experiments, combined with those of previous investigations, give a more complete physical characterization of the flow near and through the slotted wall of a transonic wind tunnel.

Everhart, Joel L.↗

NASA NTRS1993

A general theory of two- and three-dimensional rotational flow in subsonic and transonic turbomachines

This report represents a general theory applicable to axial, radial, and mixed flow turbomachines operating at subsonic and supersonic speeds with a finite number of blades of finite thickness. References reflect the evolution of computational methods used, from the inception of the theory in the 50's to the high-speed computer era of the 90's. Two kinds of relative stream surfaces, S(sub 1) and S(sub 2), are introduced for the purpose of obtaining a three-dimensional flow solution through the combination of two-dimensional flow solutions. Nonorthogonal curvilinear coordinates are used for the governing equations. Methods of computing transonic flow along S(sub 1) and S(sub 2) stream surfaces are given for special cases as well as for fully three-dimensional transonic flows. Procedures pertaining to the direct solutions and inverse solutions are presented. Information on shock wave locations and shapes needed for computations are discussed. Experimental data from a Deutsche Forschungs- und Versuchsanstalt fur Luft- und Raumfahrt e.V. (DFVLR) rotor and from a Chinese Academy of Sciences (CAS) transonic compressor rotor are compared with the computed flow properties.

Wu, Chung-Hua↗

NASA NTRS1961

Transonic Aerodynamic Loading Characteristics of a Wing-Body-Tail Combination Having a 52.5 deg. Sweptback Wing of Aspect Ratio 3 With Conical Wing Camber and Body Indentation for a Design Mach Number of Square Root of 2

An investigation has been made of the effects of conical wing camber and body indentation according to the supersonic area rule on the aerodynamic wing loading characteristics of a wing-body-tail configuration at transonic speeds. The wing aspect ratio was 3, taper ratio was 0.1, and quarter-chord-line sweepback was 52.5 deg. with 3-percent-thick airfoil sections. The tests were conducted in the Langley 16-foot transonic tunnel at Mach numbers from 0.80 to 1.05 and at angles of attack from 0 deg. to 14 deg., with Reynolds numbers based on mean aerodynamic chord varying from 7 x 10(exp 6) to 8 x 10(exp 6). Conical camber delayed wing-tip stall and reduced the severity of the accompanying longitudinal instability but did not appreciably affect the spanwise load distribution at angles of attack below tip stall. Body indentation reduced the transonic chordwise center-of-pressure travel from about 8 percent to 5 percent of the mean aerodynamic chord.

Cassetti, Marlowe D.↗

NASA NTRS1961

A Transonic Investigation of Changing Indentation Design Mach Number on the Aeodynamic Characteristics of a 45 deg Sweptback-Wing-Body Combination Designed for High Performance

The effects of changing indentation design Mach number on the aerodynamic characteristics of a 45 deg. sweptback-wing-body combination designed for high performance have been investigated at Mach numbers from 0.80 to 1.13 in the Langley 8-foot transonic tunnel and at a Mach number of 1.43 in the Langley 8-foot transonic pressure tunnel. The Reynolds number of the investigation covered the range from approximately 2.5 x 10 (exp 6) to approximately 3.0 x 10(exp 6) based on the mean aerodynamic chord of the wing. The 45 deg. sweptback wing with camber and a thickened root was tested at 0 deg. angle of incidence on an unindented body and on bodies indented for Mach numbers M of 1.0, 1.2, and 1.4. Transonic and supersonic area rules were used in the design of the indented bodies. Theoretical zero-lift wave drag was calculated for these wing-body combinations. A -2 deg. angle of incidence of the wing, and M = 1.4 revised body indentation, and fixed transition also were investigated. Experimental values of zero-lift wave drag for the indented-body combinations followed closely the area-rule concept in that the lowest zero-lift wave-drag coefficient was obtained at or near the Mach number for which the body of the combination was designed. Theoretical values of zero-lift wave drag were considered to be in good agreement with the experimental results. At a given supersonic Mach number the highest values of maximum lift-drag ratio for the various combinations also were obtained at or near the Mach number for which the body of the combination was designed. At Mach numbers of 1.0, 1.2, and 1.43, the maximum lift-drag ratios were 15.3, 13.0, and 9.2, respectively. The use of an angle of incidence of -2 deg. for the wing in combination with the M = 1.2 body increased the zero-lift wave drag and decreased the maximum lift-drag ratio. All configurations maintained stable characteristics up to the highest lift coefficient of the investigation (C(L) approx. equal to 0.5).

Loving, Donald L.↗

NASA NTRS1960

Transonic Performance Characteristics of Several Jet Noise Suppressors

An investigation of the transonic performance characteristics of several noise-suppressor configurations has been conducted in the Langley 16-foot transonic tunnel. The models were tested statically and over a Mach number range from 0.70 to 1.05 at an angle of attack of 0 deg. The primary jet total-pressure ratio was varied from 1.0 (jet off) to about 4.5. The effect of secondary air flow on the performance of two of the configurations was investigated. A hydrogen peroxide turbojet-engine simulator was used to supply the hot-jet exhaust. An 8-lobe afterbody with centerbody, short shroud, and secondary air had the highest thrust-minus-drag coefficients of the six noise-suppressor configurations tested. The 12-tube and 12-lobe afterbodies had the lowest internal losses. The presence of an ejector shroud partially shields the external pressure distribution of the 8-lobe after-body from the influence of the primary jet. A ring-airfoil shroud increased the static thrust of the annular nozzle but generally decreased the thrust minus drag at transonic Mach numbers.

Schmeer, James W.↗

NASA NTRS1959

Wind-Tunnel Measurements of Effect of Dive-Recovery Flaps at Transonic Speeds on Models of a Seaplane and a Transport

The effects of wing-lower-surface dive-recovery flaps on the aero- dynamic characteristics of a transonic seaplane model and a transonic transport model having 40 deg swept wings have been investigated in the Langley 16-foot transonic tunnel. The seaplane model had a wing with an aspect ratio of 5.26, a taper ratio of 0.333, and NACA 63A series airfoil sections streamwise. The transport model had a wing with an aspect ratio of 8, a taper ratio of 0.3, and NACA 65A series airfoil sections perpendicular to the quarter-chord line. The effects of flap deflection, flap longitudinal location, and flap sweep were generally investigated for both horizontal-tail-on and horizontal-tail-off configurations. Model force and moment measurements were made for model angles of attack from -5 deg to 14 deg in the Mach number range from 0.70 to 1.075 at Reynolds numbers of 2.95 x 10(exp 6) to 4.35 x 10(exp 6). With proper longitudinal location, wing-lower-surface dive-recovery flaps produced lift and pitching-moment increments that increased with flap deflection. For the transport model a flap located aft on the wing proved to be more effective than one located more forward., both flaps having the same span and approximately the same deflection. For the seaplane model a high horizontal tail provided added effectiveness for the deflected-flap configuration.

Heath, Atwood R., Jr.↗