High-speed Wind Tunnel Research Articles

Steady and Unsteady Temperature Measurements in the Wake of a Linear Cascade With Heated Airfoils

Abstract Understanding basic aerodynamic and thermodynamic processes in engine components is critical to achieving higher efficiencies and lower fuel consumption in aircraft engines. To aid in this process, a linear compressor cascade was investigated in the high-speed cascade wind tunnel of the Institute of Jet Propulsion to quantify the influence of heat transfer on the temperature distribution in the wake and, finally, the profile loss. For this purpose, a patented five-hole probe with an integrated thermocouple was developed and applied for steady measurements. Additionally, a hot-wire measurement setup was implemented to receive temperature fluctuations via the constant current (CC) mode as well as velocity fluctuations via the constant temperature (CT) mode. A novel method for a two-way temperature and velocity correction for the two types of hot-wire measurement is presented. Good agreement between the measurement data of the five-hole probe and averaged data from hot-wire anemometry was found. The temperature distribution indicates the occurrence of energy separation which in some cases is overlain with the effects of heat transfer. In addition, the analysis of unsteady fluctuations of temperature and velocity give more detailed information about the vortex shedding in the wake, including the size of the vortices. Finally, this is the first discussion of energy separation at a compressor cascade combined with overlain effects of heat transfer on the blade surface.

The Effects of Inlet Boundary Layer Condition and Periodically Incoming Wakes on Secondary Flow in a Low Pressure Turbine Cascade

AbstractA particular turbine cascade design is presented with the goal of providing a basis for high quality investigations of endwall flow under high-speed conditions with unsteady inflow. The key feature of the design is an integrated two-part flat plate serving as a cascade endwall at part-span, which enables a variation of the inlet endwall boundary layer conditions. The new design is applied to the T106A low pressure turbine cascade for endwall flow investigations in the High-Speed Cascade Wind Tunnel of the Institute of Jet Propulsion at the Bundeswehr University Munich. Measurements are conducted under realistic flow conditions (M2th = 0.59, Re2th = 2 · 105) in three cases of varying endwall boundary layer conditions with and without periodically incoming wakes. The endwall boundary layer is characterized by 1D-CTA measurements upstream of the blade passage. Secondary flow is evaluated by five-hole-probe measurements in the turbine exit flow. A strong similarity is found between the time-averaged effects of unsteady inflow conditions and the effects of changing inlet endwall boundary layer conditions regarding the attenuation of secondary flow. Furthermore, the experimental investigations show that all design goals for the improved T106A cascade are met.

High-Speed Wind Tunnel Twin-sting Adaptor Design and Verification for Commercial Aircraft

Rear sting, Z sting, and fin sting supports are generally used to measure force and moment during high-speed wind tunnel tests for commercial aircraft, It is necessary to apply another support to obtain the sting interference. Twin sting support is the most appropriate way to get the sting interference due to less secondary interference, and the twin-sting adaptor is core part of twin-sting support, which directly affect the test data accuracy and test envelope. This work is to introduce design and verification of the twin-sting adaptor based on a commercial aircraft model, CAE was applied to analyze and evaluate the aerodynamics, strength and stiffness during the design of adaptor. The adaptor was finally verified in the high-speed wind tunnel testing, and testing data demonstrate that the design have less aerodynamics interference on the wing, and the accurate sting interference could be obtained. The design of twin-sting adaptor could offer a solution for commercial aircraft to carry out wind tunnel testing of sting interference and rear fuselage characteristics.

On High-Resolution Pressure Amplitude and Phase Measurements Comparing Fast-Response Pressure Transducers and Unsteady Pressure-Sensitive Paint

AbstractThis paper presents the simultaneous application of fast-response pressure transducers and unsteady pressure-sensitive paint (unsteady PSP) for the precise determination of pressure amplitudes and phases up to 3000 Hz. These experiments have been carried out on a low-pressure turbine blade cascade under engine-relevant conditions (Re, Ma, and Tu) in the high-speed cascade wind tunnel. Periodic blade/vane interactions were simulated at the inlet to the cascade using a wake generator operating at a constant perturbation frequency of 500 Hz. The main goal of this paper is the detailed comparison of amplitude and phase distributions between both flow sensing techniques at least up to the second harmonic of the wake generator’s fundamental perturbation frequency (i.e., 1000 Hz). Therefore, a careful assessment of the key drivers for relative deviations between measurement results as well as a detailed discussion of the data processing is presented for both measurement techniques. This discussion outlines the mandatory steps which were essential to achieve the quality as presented down to pressure amplitudes of several Pascal even under challenging experimental conditions. Apart from the remarkable consistency of the results, this paper reveals the potential of (unsteady) PSP as a future key flow sensing technique in turbomachinery research, especially for cascade testing. The results demonstrate that PSP was able to successfully sense pressure dynamics with very low fluctuation amplitudes down to 8 Pa.

Application of colour fluorescent oil flow visualization for a high speed cavity

For studying the complex flow structure of a high speed cavity, including shear layer, separation, reattachment and vortex, the single-color fluorescent oil flow visualization system in the 0.6m×0.6m trisonic wind tunnel of China Aerodynamic Research Development Centre was enhanced. Different portions of the cavity such as its bottom wall, left and right side walls, front and rear walls and area outside the cavity utilized different oil films with different fluorescent particles to visualize the flow mixing more effectively in the cavity. Specialized ultraviolet light sources are used to enhance the oil flow image contrast. Additionally, to ensure the oil film follows the surface streamlines and indicates the skin friction lines more accurately, the thickness and viscosity of the oil film are well controlled. A high-speed cavity wind tunnel test was conducted to validate the improved oil flow visualization system at Mach number ranging from 0.6 to 2.0. The experimental result shows that the improved oil flow system can effectively visualize complex flow patterns of the cavity flow at all Mach numbers. Colour fluorescent oil flow visualization system established in the present study provides a capability to visualize the flow mixing phenomenon in cavity flows with a high image contrast.

Analysis of aircraft-store compatibility for internal weapons separation

An evaluation criterion of aircraft-store compatibility for internal weapons separation is put forward and the drop tests are conducted in a high-speed wind tunnel with the Mach 1.5. The dynamic store models are designed by light Mach scaling (LMS) and are released to the freestream by an ejection mechanism driven by the high-pressure air from two rectangular cavities with a length-to-depth ratio of 7.76 and 11.86. The three affected factors of initial pitch angle, angle-of-attack of aircraft and length-to-depth ratio are considered and analysed. The result of the drop tests shows the aircraft-store compatibility of the internal weapons decreases along with increasing AOA of carrier aircraft and even changes from compatibility to incompatibility. The flow characteristics of the cavity have a great influence on the aircraft-store compatibility of internal weapons separation. The compatibility of aircraft-store separation tends to obviously decrease when the cavity change from open to transitional cavity flow. The initial pitch angle of the store is the key factor affecting the aircraft-store compatibility. With the increase of initial pitch angle, the aircraft-store compatibility of is obviously reduced, even unsafe.

Experimental Study on High Speed Two-dimensional Airfoil Based on Conventional High Speed Wind Tunnel Test

The wind tunnel test of two-dimensional airfoil is usually carried out in the special wind tunnel for the airfoil test, but this kind of wind tunnel is usually small in size and low in Reynolds number. On the contrary, the conventional wind tunnel test is generally large in size, in which higher test Reynolds number can be obtained, but there is wall interference and other problems to overcome at the same time. In this study, the RAE2822 airfoil was used to carry out the wind tunnel test of high-speed two-dimensional airfoil in FL-3, which is a large-scale conventional high-speed wind tunnel and the section pressure coefficient was measured. Experimental data after calculation and correction comparing with the standard model data shows that it is practically possible to carry out airfoil wind tunnel test in conventional high-speed wind tunnel.

Study on temperature characteristics and correction of strain gauge balance in actual airflow environment

Strain gauge balances are main force measuring instruments used in high speed wind tunnels. As one of main performance of the strain gauge balance, accuracy plays a vital role in wind-tunnel test technologies. However, spatial temperature gradients across the strain gauge balance caused by the inner deformation of the balance due to temperature changes of the actual airflow environment, has some adverse influence upon the measurement accuracy of the strain gauge balance, especially for the long-duration testing in continuous wind tunnels. In order to better understand the temperature characteristics of strain gauge balances and improve their measurement accuracy, a study has been performed with the balance 30G and calibration model GBM-04A in the 0.6 m×0.6 m continuous transonic wind tunnel at China Aerodynamics Research & Development Center (CARDC), and a correction method for the bridge zero drifts of the strain gauge balance due to temperature gradients across the strain gauge balance for long-duration testing was developed. The study case shows that the deviation of measurement accuracy due to the non-uniform temperature gradient across the strain gauge balance body is about 0.0003 for the drag coefficient at M=0.6 on the calibration model GBM-04A, which can be corrected in a large extent with the method in this paper. In some extent, the correction method can be of great benefit to those long-duration tests in the continuous wind tunnel for its simple operation and high universality.

Loss determination at a linear cascade under consideration of thermal effects

ABSTRACTTargeting higher efficiencies and lower fuel consumption of turbomachines, heat transfer and profile loss are research topics of particular interest. In contrast to that, the interaction of both was, so far, rarely investigated, but gains in importance in recent research activities. The profile loss of engine components can be characterised by the airfoil wakes at the blade rows utilising established measurement and evaluation methods for which an adiabatic flow is typically supposed. To enable the investigation of the influence of heat transfer at the blade on the loss characteristics, a novel evaluation procedure was set up. In addition to the pneumatic data, the total temperature in the airfoil wake at a linear cascade was measured by means of a five-hole probe with an integrated thermocouple. For the evaluation and analysis of these data, different definitions of the loss coefficient were investigated and, finally, extended to account for thermal aspects. Furthermore, established techniques to average the local wake data were applied and compared with special focus to their suitability for non-adiabatic cases. Moreover, an extended version of the mixed-out average as defined by Amecke was utilised applying not only a far-reaching consideration of a temperature gradient but also the inclusion of the third spatial dimension to enable the evaluation of field traverses in addition to single wake traverses. These techniques were applied to wake measurement data from a linear compressor cascade gained in a special test set-up in the high-speed cascade wind tunnel for different operating points and different blade temperatures. The suitability of the new methods could be proven, and initial steps of the aerodynamic analysis of the resulting data are presented. Thereby, the acquired techniques turned out as powerful methods for the evaluation of wake traverses on compressor and turbine cascades under non-adiabatic conditions.

Experimental research on collector effect on operating characteristics of high-speed free-jet wind tunnel

Collector is one of the key parts of the high-speed free-jet wind tunnel, which directly affects the performance of wind tunnel. In this paper, a small high-speed free-jet wind tunnel is used as the test platform to carry out the start-up and operation characteristics test of wind tunnel. Start-up pressure characteristics were obtained at various Mach number, and the influence of the collector throat heights and collector positions on start-up pressure and matching pressure was preliminarily studied. The results show that the high collector throat height can meet the requirements of wind tunnel start-up and pressure matching. The closer the collector is to the nozzle, the smaller the corresponding starting pressure and matching pressure. However, this trend weakens with the decrease of Mach number, which lays a foundation for the development of large high-speed free jet equipment.

Study on Calculation of the Buoyancy Drag Coefficient in the High-Speed Wind Tunnel Test

In order to obtain the drag coefficient in accuracy, it needs to correct the model’s buoyancy drag coefficient in the high-speed wind tunnel test. However, there are several methods on calculating the buoyancy drag coefficient, each will generate a result differently. Three types of methods are studied in this paper, the formulas are derived, the buoyancy drag coefficients are calculated, and the results are compared in the end. It shows that the most exact method is the one based on pressure fitting function, it’s suitable for correcting the model’s drag coefficient with high accuracy requirement.

Rotor airfoil aerodynamic design method and wind tunnel test verification

Well-designed airfoil is very important for high-performance rotor. This paper developed an efficient multi-objective and multi-constraint optimization design system for rotor airfoils based on RANS analysis, and verified the performance of the optimized airfoil. Using CRA09-A as the baseline rotor airfoil, the CRA09-B optimized rotor airfoil was designed successfully. Combined with the foundation of high-precision rotor airfoil stationary test technology, the CRA09-B and CRA09-A rotor airfoils were tested in the S3MA high-speed wind tunnel of ONERA. In order to correct the aerodynamic data, a single parameter linear wall pressure method is used to consider the tunnel effects. The results indicate that multi-objective and multi-constraint optimization design method developed in this study is reliable, and that CRA09-B optimized airfoil provides better stationary performance than CRA09-A airfoil in terms of maximum lift coefficient and lift-to-drag ratio.

A simplified design approach for high-speed wind tunnels. Part I: Table of inclination

This study aims to develop a simplified approach for the design of high-speed wind tunnels, operating within the range of Mach 1-4. The study is conducted in two stages; design stage-I consists of the basic parts of a high-speed wind tunnel, i.e. test section and nozzle throat sizes along with the calculations of the minimum pressure ratios necessary for the starting and running of the wind tunnel. Design stage-II is the aerodynamic design of the inlet nozzle contour with the focus on the divergent nozzle section only. In this stage, the fundamental method of characteristics (MOC) is employed by using region-to-region solution approach. Due to the long and complicated procedure of classical MOC, a simpler approach of “table of inclination” is proposed. This inclination table approach is simple in use, highly precise, and efficient which reduces more than 75 % of the calculation time of the MOC. Validation of the nozzle contour design is performed with the Area-Mach ratio relation followed by the results verification obtained from the MATLAB programming. The contour design results of the study are highly accurate (less than 1 % error) which manifests the future possible significance of the inclination table approach for supersonic wind tunnel designs.

Effect of leading-edge plate controlling on subsonic cavity

Cavity noise caused by the high-speed airflow has been paid much attention in the field of aerospace, and the study of cavity noise suppression has an important significance on improving the safety of aircraft. Effects of the leading-edge plate on the flow and the noise of the cavity model (the ratio of length to depth is 6) at a Mach number (Ma) of 0.6 are investigated with high-speed wind tunnel experiment. The acoustic and the flow field information in the cavity are obtained with the dynamic/static pressure measurement and oil flow experiment. The result shows that the leading-edge plate can greatly lift the shear layer, raise the impact position on the back-wall area, and reduce the flux rate and intensity in the cavity. With the controlling of leading-edge plate, the static pressure and backflow area are effectively suppressed and the SPL and peak noise also fall down significantly. The leading-edge plate provides a new method for cavity noise suppression in subsonic flow condition, which can be effectively applied to flow/noise controlling of cavity structure on aircraft.

High speed wind tunnel force test with traditional pressure-test-model

Usually, force tests and pressure tests are conducted with different model in wind tunnel which could cause trouble for aircraft design as the pressure test results always could not match the force test results. So attempts to do force tests along with pressure tests are conducted. Results show that it is possible to do force test based on pressure-test-model when some improvements are done with the model and sting. Both force and pressure test results can be measured at the same time. The amount of pressure test results need not be reduced, and yet precision and accuracy of the force tests are acceptable for aircraft design, and this technology can be popularized with low cost.

Geometric Effects Analysis and Verification of V-Shaped Support Interference on Blended Wing Body Aircraft

A special V-shaped support for blended wing body aircraft was designed and applied in high-speed wind tunnel tests. In order to reduce the support interference and explore the design criteria of the V-shaped support, interference characteristics and geometric parameter effects of V-shaped support on blended wing body aircraft were numerically studied. According to the numerical results, the corresponding dummy V-shaped supports were designed and manufactured, and verification tests was conducted in a 2.4 m × 2.4 m transonic wind tunnel. The test results were in good agreement with the numerical simulation. Results indicated that pitching moment of blended wing body aircraft is quite sensitive to the V-shaped support geometric parameters, and the influence of the inflection angle is the most serious. To minimize the pitching moment interference, the straight-section diameter and inflection angle should be increased while the straight-section length should be shortened. The results could be used to design special V-shaped support for blended wing body aircraft in wind tunnel tests, reduce support interference, and improve the accuracy of test results.

Aerodynamic noise evaluation for Aeolian tone and acoustical insulating plate by similarity law

The objective of this research is to assess aerodynamic noise sources and the wind velocity conversion of the similarity law in an increasing wind velocity. We measured the aerodynamic noise that radiates from cylinders with and without periodic holes and a pantograph model with and without a sound insulating plate in a low-noise high-speed wind tunnel. The noise level of the cylinders was in proportion to the sixth power of wind velocity because the cylinders acted as a dipole sound source. The effect of noise reduction caused by the change in flow structure, that is, the periodic holes, also continued. However, the second peak noise radiating from the cylinder was in proportion to the sum of the sixth power and the eighth power of wind velocity and was underestimated by wind velocity conversion based on a dipole sound source. The noise level of the pantograph model with the sound insulating plate was in proportion to lower than the sixth power of wind velocity in contrast to the pantograph model itself. This is because the noise reduction effect of the sound insulating plate had a frequency characteristic. In this research, when increasing the wind velocity, the dominant pantograph noise frequency region shifted to a higher level, and the dominant noise level began to be affected by the noise reduction of the sound insulating plate. This caused the reduction effect of the sound insulating plate to be underestimated by 2.5 dB due to the noise level conversion of wind velocity. To solve this problem, the plate’s theoretical reduced noise level, which was calculated by using the Fresnel number, was subtracted from the estimated noise level of the pantograph itself. This lead to pantograph noise with the sound insulating plate being estimated within a 0.8-dB difference.

Optimization and verification of free flight separation similarity law in high-speed wind tunnel

Based on the similarity of separation time, a similarity law optimization method for high-speed weapon delivery test is derived. The typical separation state under wind load is simulated by the numerical method. The real separation data of aircraft, separation data of previous test methods, separation data of ideal wind tunnel test of previous methods, and simulation data of the proposed optimization method are obtained. A comparison of the data shows that the method proposed can improve the performance of tracking. Similarity law optimization starts with the development of motion equations and dynamic equations in the windless state to address the problems of mismatching between vertical and horizontal displacement, and to address the problems of separation trajectory distortion caused by insufficient gravity acceleration of the scaling model of existing light model. The ejection velocity of the model is taken as a factor/vector, and is adjusted reasonably to compensate the linear displacement insufficiency caused by the insufficient vertical acceleration of the light model method, so as to ensure the matching of the vertical and horizontal displacement of the projectile, and to improve the consistency between the test results of high-speed projection and the actual separation trajectory. The optimized similarity law is applicable to many existing free-throwing modes of high-speed wind tunnels. The optimized similarity law is not affected by the ejection velocity and hanging mode of the projectile. The optimized similarity law is suitable not only for the launching of the buried ammunition compartment and external stores, but also for the test design of projectile launching and gravity separation.

Research on the influence of different data distributions on the fitting results of subsonic compressible flow turbulence level and its uncertainty

On the basis of results of subsonic compressible flow turbulence level in high speed wind tunnels by hyperbolic fitting, the uncertainty of turbulence level is quantitatively studied by using Monte Carlo simulation method. In order to generate stochastic sample data for Monte Carlo simulation, two distribution types, uniform distribution and normal distribution, are used to generate the data points to be fitted. The effect of changing characteristic parameters of two distribution types on the turbulence fitting results and its uncertainty is studied. The results show that: (1) for uniform distribution, the larger the data distribution bandwidth is, the larger turbulence level and its uncertainty is; (2) for normal distribution, when the mean of the distribution is fixed and the standard deviation is changed, turbulence level results are basically the same, and the uncertainty and the standard deviation have the same trend; (3) comparing the two distribution types, turbulence level and its uncertainty can both be accurately solved, as long as the characteristic parameters are appropriate.

Flange balance application in high speed wind tunnel test of BWB aircraft

In order to improve quality and efficiency of Blended Wing-Body (BWB) aircraft’s high-speed wind tunnel test, while verify the reliability and practicability of flange balance in the test of such a layout aircraft. The specific flange-connected balance of BWB aircraft was designed and processed, which was designed under the limits of aerodynamic load and model interspace size. The comparison test results between the flange balance and the traditional cone balance were obtained by conducting a comparison test in the 2.4-meters transonic wind tunnel. The results show that the flange balance has high measurement accuracy and good correlation with the traditional cone balance. Flange balance meets the requirement of BWB layout aircraft high speed wind tunnel test and has a broad application prospect, so that it could be popularized in the future.