Supersonic Wind Tunnel Tests Research Articles

布製超音速パラシュートの空力特性

Supersonic wind tunnel tests of flexible parachutes whose canopies had a hemispherical shell configuration were conducted. This study was performed to explore the mechanism of flow field oscillation and vibration of canopy movement that were caused by the interaction between the supersonic flow and flexible fabric of the canopies. Effects of the flexibility and gas permeability of fabrics to the drag characteristics of the parachutes were investigated. Three kinds of fabrics (i.e. thin nylon, thick nylon and polyester) were used to produce the canopies. Sizes of canopies were also changed in three steps (nominal diameter D0 = 30, 40, 50[mm]) to investigate the variation of the drag. A load cell was installed in the capsule to measure unsteady and time-averaged drag values of the parachute. Schlieren images by a high speed camera provided us with visualized flow fields showing shock deformations. Moreover, a three-dimensional motion capturing system was used to evaluate the movement of the canopy. The process of parachute ejection from the capsule was also examined. As a result, it was indicated that canopy fabrics with higher gas permeability induce lower mean drag values. On the other hand, the lower flexibility, the larger drag fluctuation is. Furthermore, the spectrum of the canopy movement showed the similar profiles to the spectrum that of the drag fluctuation indicating the direct link between the canopy movement and its drag fluctuation.

Videogrammetry and its Application on Attack Angle Measuring in Wind Tunnel Testing

The videogrammetric measurement is a favorite technology in wind tunnel testing, because it neither destroys the model’s shape, nor changes the stiffness or strength. In this paper, a videogrammetric measurement system(VMS) is introduced, and it is used to measure model’s attack angle. Three engineering videogrammetric experiments of attack angle in 2 meter supersonic wind tunnel testing have demonstrated that the precision of VMS established by this paper is ≤0.0094°, so it is useful and effective.

Videogrammetry of the Model’s Attitude in Wind Tunnel Testing

The advanced precision of drag coefficient is 0.0001 for the high speed wind tunnel test of measuring forces, the model’s angle of attack precision is ≤0.01°following errors distribution. A videogrammetric method of model’s attitude is therefore proposed, its uncertainty is investigated, and a compensation method of its systematic error is also presented by this paper. The three engineering videogrammetric experiments of attack angle in 2 meter supersonic wind tunnel testing have demonstrated that measuring standard deviation of videogrammetric measurement system established by this paper is ≤0.0094°, in addition it neither destroys the model’s shape, nor changes the stiffness or strength, so it is useful and effective.

Uncertainty Identification of Supersonic Wind-Tunnel Testing

In this study, a methodology for the identification of uncertainty of supersonic wind tunnel testing is developed. Instead of conventional multi-tunnel testing, multi-position testing is performed across the characteristic net of supersonic flow. From the variations in aerodynamic coefficients in body axes, the uncertainty caused by flow quality is identified about models of interest. Then, after the consideration of practically independent parameters, uncertainties of aerodynamic coefficients in stability axes are combined. Their validities are confirmed by correlation results of tunnel-to-tunnel comparison tests.

風洞試験模型静的空力弾性変形評価における簡易空力モデル構築

Effects of static aeroelastic deformation of a wind-tunnel test model on the aerodynamic characteristics are discussed in supersonic wind-tunnel tests. The static aeroelastic deformation of the main wing of a supersonic airplane model in the JAXA 1m×1m supersonic wind-tunnel is estimated by the iteration of the full-potential aerodynamic analysis and the FEM structural analysis. The estimated deformation corrections improve the agreement of the measured and predicted aerodynamic performance of the model. Some simple aerodynamic load estimation models are constructed with a few aerodynamic calculation results in order to improve the efficiency of the static aeroelastic deformation estimation for practical applications.

SHAPE MEMORY ALLOY ACTUATED MICRO-FLOW EFFECTORS FOR VORTEX MANIPULATION

Smart structures are seen as an enabling technology for designing innovative control actuation systems for future missiles. In this study, the feasibility of employing shape memory alloy (SMA)-actuated micro-flow effectors to control the vortex shedding behaviour that produces side forces on slender body missiles is examined. Supersonic wind tunnel tests were performed on a slender finless missile equipped with static micro-flow effectors on a conical nose to determine suitable configurations that could generate significant side forces. Shape memory alloy actuators for the flow effector were developed using numerical techniques and validated experimentally. Matching the force-displacement characteristics of the SMA actuator to the micro-flow effector force-displacement requirement was accomplished by a compliant transmission mechanism. The dynamic performance of the micro-flow effector was assured with a two-step variable structure control law. Closed-loop test results showed that the control law was capable of providing effective displacement control up to 1.0 Hz.

Study of Direct-Measuring Skin-Friction Gauge with Rubber Sheet for Damping

This study concerns the direct measuring technique for skin-friction determination. Such a device measures the force on a small, movable surface element as a shear e ow passes over the surface. The typical direct-measuring skin-friction gauge uses a viscous liquid in the gap between the movable surface piece and the casing to dampen vibrations,tocreateanevensurface,tominimizetheeffectsofpressuregradients,andfortemperaturestabilization. In testing, the liquid slowly leaks out. Therefore, we have considered gauges with rubber to e ll all or some of the gap. This led to the development of a gauge with a thin rubber sheet to cover the face of the gauge instead of a previousdesign withrubbere lling theentireinternalvolume.First, ae niteelementmethodmodelwasemployed to fully understand the strain e eld involved and to e nalize the design. The resulting design consisted of a plastic skin frictiongaugewitha12.7-mm-diam e oatingelementonacantileverbeame exure,anapproximately0.51-mm-thick rubber sheet, a 1.6-mm-wide gap around the e oating element on a cantilever beam e exure, and semiconductor strain gauges at the beam base. Vibration tests were performed to determine if this design produced the required damping. These tests were successful. Supersonic wind-tunnel tests at Mach 2.4 with a total pressure of 3 atm and ambient total temperature demonstrated that the rubber sheet survived repeated tests and provided adequate damping. The skin-friction data obtained compared well with theory and other measurements.

Flutter Margin Evaluation for Discrete-Time Systems

A flutter prediction method using a new parameter is proposed, and its effectiveness is examined by numerical and experimental analysis. In this method, the aeroelastic system is identified as the autoregressive moving average (ARMA) model from the random response of a wing excited by flow turbulence, so that no special excitation device is needed. The new parameter is based on a stability test for discrete-time systems and is calculated algebraically from the autoregressive coefficients of the estimated ARMA model. Numerical calculation using a two-dimensional wing model shows that the parameter decreases almost linearly toward zero with increasing dynamic pressure. This is a superior property as a flutter predictor in comparison with damping coefficients and the stability parameter used conventionally. The method is applied to supersonic wind-tunnel flutter test data to examine the validity in actual data analysis. The results obtained under different test conditions demonstrate the advantage of the present method for an accurate and reliable flutter prediction. Moreover, an analytical consideration reveals that the calculated values of the proposed parameter are nearly equal to that of the flutter margin introduced by Zimmerman and Weissenburger (Zimmerman, N. H., and Weissenburger, J. T., Prediction of Flutter Onset Speed Based on Flight Testing at Subcritical Speeds, Journal of Aircraft, Vol. 1, No. 4, 1964, pp. 190-202).

101 薄膜センサーを用いた物体回りの高速流れの計測(その2)

In the super sonic flow through the cascade of a gas-turbine engine, flow separation and shock waves are important phenomena impacting the performance. In previous report1) flow separation can be detected using hot film sensors. And we show the possibility that shock waves can be detected using these sensors. This report describe additional experiment to improve the detection of shock waves. In super sonic wind tunnel test, two kinds of shock waves, that are differed in strength, are indused to occur on the surface of the cone-like models at Mach 2.5. Data are collected from hot film sensors located on the surface. Schlieren flow visualization method is also applied to check the flow. As the results, the shock wave can be detected by the differences between the sensor output.

102 薄膜センサーを用いた物体回りの高速流れの計測

In the super sonic flow through the cascade of a gas-turbine engine, flow separation and shock waves are important phenomena impacting the performance. This report describes an basic experiment to detect these phenomena using hot film sensors. In the super sonic wind tunnel test, flow separation or a shock wave is induced to occur on the surface of the cone-like model at Mach 2.5. Data are collected from hot film sensors located on the surface. Schlieren and oil flow visualization method are also applied to check the flow. As the results, the separation and the shock wave can be detected by the differences between the sensor output.

An optical angle of attack sensor

A major source of transonic and supersonic wind-tunnel test data uncertainty are angle-of-attack-measurement errors caused by unknown sting and balance deflections under load. Since dynamic loads in pressurized or cryogenic wind tunnels generally exceed those in conventional low-speed atmospheric wind tunnels, the need to account for these distortions during model testing is even more acute. A novel laser-based instrument for the in-situ measurement of wind-tunnel-model angle of attack that enables continuous, time-dependent measurements to be made without signal dropout is described. Proof-of-concept experiments, along with the results of recent measurements conducted at the NASA Ames 9-ft*7-ft supersonic wind tunnel, are presented. Experiments were also conducted to determine the reliable range, sensitivity, and long-term stability of the instrument. The results show that 0.01 degrees sensitivity can be achieved and that optical and detector packaging requirements are less stringent than those for current tilt-sensor or accelerometer model installations.< <ETX xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">&gt;</ETX>

Stability coefficients of a missile at angles of attack

Introduction T current trend in maneuverable missile development has created an interest in the nonlinearity of the stability coefficients with angle of attack for high-fineness ratio, finned missiles. New wind-tunnel testing techniques and flowfield analyses are being undertaken. To validate these new procedures, comparisons are being made with the Basic Finner configuration. The Basic Finner, Fig. 1, has been selected by the Supersonic Tunnel Association and AGARD as a standard research configuration for the evaluation of new testing and analysis techniques. In this regard the static and damping stability coefficients for a Basic Finner model are presented as obtained from supersonic wind-tunnel tests. These coefficients will supplement the currently available The stability coefficients presented in this note were determined for angles of attack from 0 to 20 degrees, at a Mach number of 1.5. The fineness ratio (length/diameter) was 10. A standard free-oscillation wind tunnel testing technique was used to obtain the angular performance data. This technique employs a strut or side mount, and the model was free to oscillate in the pitch plane. A nonlinear data analysis procedure was used to extract the variation of the pitching and damping stability coefficients, Cma and Cmq + Cm&9 respectively, with angle of attack.

Influence of Probe Geometry on Pitot-Probe Displacement in Supersonic Turbulent Flow

An experiment was conducted to determine the varying effects of six different probe-tip and support-shaft configurations on pitot tube displacement. The study was stimulated by discrepancies between supersonic wind-tunnel tests conducted by Wilson and Young (1949) and Allen (1972). Wilson (1973) had concluded that these discrepancies were caused by differences in probe geometry. It is shown that in fact, no major differences in profiles of streamwise velocity over streamwise velocity at boundary-layer edge vs normal coordinate over boundary-layer total thickness result from geometry. The true cause of the discrepancies, however, remains to be discovered.

Supersonic wind-tunnel tests of wavy-walled cylinders.

Supersonic wind-tunnel tests of wavy-walled cylinders.

Some Factors Involved in the Supersonic Wind Tunnel Testing of Models

The use of models in the determination of aircraft behaviour is as old as flying itself, and wind tunnel models in particular have played a very important part. Now that supersonic flight of aircraft and missiles is a matter of common experience, it seems appropriate to look at some of the difficulties involved in the supersonic tunnel testing of models of such machines.

First USAF “SMART” test

The transient dynamic behaviour of supersonic disk-gap-band parachutes in a Mars entry environment involving fluid structure interactions is studied. Based on the multi-material Arbitrary Lagrange–Euler method, the coupling dynamic model between a viscous compressible fluid and a flexible large deformation structure of the parachute is solved. The inflation performance of a parachute with a fixed forebody under different flow conditions is analysed. The decelerating parameters of the parachute, including drag area, opening loads, and coefficients, are obtained from the supersonic wind tunnel test data from NASA. Meanwhile, the evolution of the three-dimensional shape of the disk-gap-band parachute during supersonic inflation is presented, and the structural dynamic behaviour of the parachute is predicted. Then, the influence of the presence of the capsule on the flow field of the parachute is investigated, and the wake of unsteady fluid and the distribution of shock wave around the supersonic parachute are presented. Finally, the structural dynamic response of the canopy fabric under high-pressure conditions is comparatively analysed. The results show that the disk-gap-band parachute is well inflated without serious collapse. As the Mach numbers increase from 2.0 to 2.5, the drag coefficients gradually decrease, along with a small decrease in inflation time, which corresponds with test results, and proves the validity of the method proposed in this paper.

A Flexible Nozzle for a Small Supersonic Wind Tunnel

The design of a small supersonic wind-tunnel test section (4 by 10 in.) incorporating a flexible nozzle is outlined. The flexible nozzle consists of a high-strength stepped steel plate. Two screw jacks provide an easy means of continuously changing the nozzle's shape according to the aerodynamic requirements. The boundary-layer compensation can also be varied during operation. Pressure surveys, together with schlieren and interferometric analysis of the test section, show the flow to be uniform over the operating range (M = 1.1 to 1.5).