High Speed Wind Tunnel Testing Research Articles

Experimental study of the influence factors of the droplet size of rotary nozzles during manned helicopter spraying operations

The droplet size significantly affects the performance of aerial spraying operations. Thus, it is crucial to develop rotary nozzles for helicopter spraying operations and analyze the influence of droplet size. We conducted a calibration test of the nozzle rotation speed to analyze the effects of air velocity, blade installation angle, hub diameter, and blade length on the rotation speed. The results showed that the rotation speed increased with an increase in air velocity and nozzle diameter and decreased with an increase in the blade angle. The rotation speed of three types of rotary nozzles was measured in the range of 576–15 358 r/min. A high-speed wind tunnel test was conducted to determine the droplet size of the rotary nozzles at different rotation speeds, spray flow rates, and hub mesh numbers. The influence degree of the three factors on the droplet size was rotation speed > spray flow rate > mesh number. The multiple linear regression equation of the volume median diameter (VMD) of the droplet size was obtained. The droplet size obtained from wind tunnel and flight tests at 5000 and 7000 r/min was compared. The droplet size was in the range of 148.56–259.39 μm and was slightly larger under wind tunnel conditions than during the flight test. The relative width of the droplet spectrum was small for both test methods, indicating that the droplet size distribution of the three rotary nozzles was uniform. Finally, the spray nozzles were used for insect control in forest areas and achieved a control efficiency of 94.6. This study provides a reference for optimizing the operating parameters of spraying nozzles and improving the performance of helicopter-based low-volume spraying operations.

Uncertainty analysis method of internal drag measurement for flow-through model in high-speed wind tunnel

For aircraft with complex air intake and exhaust systems, the flow-through model simulating the internal flow passage is generally used to reduce the interference of internal and external flow fields when carrying out high-speed wind tunnel tests. The drag generated by the internal flow passage needs to be accurately measured and deducted from the full-model aerodynamic force, so as to obtain the aerodynamic characteristics of the aircraft to be tested. However, the existing internal drag measurement methods do not evaluate the uncertainty of the measurement results, which leads to the failure to evaluate the reliability of the flow-through model high-speed wind tunnel test results. This paper introduces the device and measurement method of internal drag of flow-through model. Based on GUM uncertainty analysis method, the analysis method of internal drag measurement uncertainty is established through strict mathematical derivation, which provides technical support for the evaluation of internal drag measurement uncertainty.

Refined Aerodynamic Test of Wide-Bodied Aircraft and Its Application

The large dual-channel wide-bodied aircraft has a long range and a high cruise Mach number. Therefore, its aerodynamic design requires a high level of wind tunnel test refinement. Based on the requirements of aerodynamic design for the future wide-bodied aircraft and the characteristics of high-speed wind tunnel tests, the error theory is introduced to analyze the factors affecting the accuracy of the test data. This study carries out a series of research on the improvement of refined aerodynamic test technology in an FL-26 wind tunnel, including design and optimization of the support system of wide-bodied aircraft, model attitude angle measurement, Mach number control accuracy, measurement and control system stability, test data correction and perfection, high-precision force balance and standard model development. In addition, the effect of the standard specification of the refined aerodynamic test is investigated to improve the data quality. The research findings have been applied in the standard model test and subsequent models of wide-bodied aircraft. The results show that when Mach numbers are less than 0.9, the control accuracy of Mach numbers in the FL-26 wind tunnel is smaller than 0.001 and the measurement error of attack angle is smaller than 0.01°. Therefore, it has the ability to correct the data influenced by factors, such as support/wall interference, model deformation, floating resistance and airflow deflection angle. The repeatability accuracy of the standard model’s comparison test shows that the lift coefficient is less than or equal to 0.0012, the drag coefficient is less than or equal to 0.00004, pitching moment coefficient is less than or equal to 0.0004. The bending resolution of the model’s deformation measurement is less than 0.2 mm, and the tensional deformation is smaller than 0.04°. The revised aerodynamic data and model deformation measurement results have good agreement with that of the ETW wind tunnel. The results demonstrate that the improved technology presented in this paper can significantly enhance the refined aerodynamic test of wide-bodied aircraft.

Steady and Unsteady Hover Test of a Lift Compounding Rotorcraft

Compound rotorcraft can be a viable option to achieve moderately high speeds of 240 knots. However, there is no systematic wind tunnel test data in the public domain to understand the complex interactional aeromechanics involved. This work is part of a test campaign to conduct multiple high-speed wind tunnel tests of various compound configurations. The compound configuration consists of a single main rotor, a single wing on the retreating side for lift augmentation, and a single rear propeller for propulsive augmentation. In this paper, hover test results are presented. Hover tests were conducted for various configurations and at different tip Mach numbers with the maximum value of 0.5. Rotor and wing performances are measured separately using individual balances. Rotor blades are equipped with strain gauges for measuring flap, lag, and torsional moments at radial stations. The wing location is below the main rotor, and the measured wing download is only 2.8% of the rotor thrust. The wing provides a partial ground effect, making the rotor more efficient and almost compensates for the download with less than 1% reduction in net figure of merit. Blade structural loads correlate satisfactorily between the analysis and measurement except for lag bending moment. This discrepancy is a large 2/rev lag bending moment measured in the test, which needs further investigation. Propeller hover test was conducted up to 2500 revolutions per minute (tip Mach number of 0.25) at a single collective. Prediction correlates well and captures the low-Reynolds-number effect for lower revolutions per minute.

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.

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.

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 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.

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.

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.

Experimental Investigation of the Boundary Layer Transition on a Laminar Airfoil Using Infrared Thermography

The transition in the boundary layer is investigated using infrared thermography (IRT). The study is carried out on a laminar airfoil in the transonic intermittent in-draft wind tunnel. The transition in the boundary layer is evocated using transition-generator strips of different thicknesses at two Mach numbers: 0.4 and 0.8. The tested transition-generators thickness to boundary layer displacement thickness ratio was from 0.42 to 1.25. The Reynolds number respect to the airfoil chord is: Re = 1.0 – 1.4·106. The six cases for different transition-generators thickness ratios were compared. The behaviours of laminar and turbulent boundary layers are discussed. The use of IRT has been proven to be an appropriate tool for detecting the transition of the boundary layer in high-speed wind tunnel testing.

Preparation and execution of the NICETRIP low- and high-speed wind tunnel tests

The present paper reports the preparation and execution of high- and low-speed wind tunnel tests of the tilt-rotor project NICETRIP. Within the NICETRIP project (Novel Innovative Competitive Effective Tilt Rotor Integrated Project), co-funded by the EU, wind tunnel tests were conducted for a heavily instrumented 1:5 scale model of a tilt-rotor aircraft (3.5 m wingspan, 1.48 m diameter rotors) based on the ERICA concept (Enhanced Rotorcraft Innovative Concept Achievement). The current paper provides an overview of the contributions of DLR to the low- and high-speed wind tunnel tests performed at DNW-LLF (German-Dutch Wind Tunnels Large Low-speed Facility) in 2013 and at ONERA-S1MA in 2014. It gives detailed information about the preparation activities, the model hardware and its control, the necessary adaptations for the model preparation at DLR, and results of the ground vibration tests. Additionally, the data acquisition and monitoring during the wind tunnel experiments, the trimming of the model in the wind tunnel and the conducted tests are described, including encountered problems and how they were solved. In the last section, some representative test results are shown.

Multidisciplinary Assessment of the Control of the Propellers of a Pusher Geared Open Rotor—Part I: Zero-Dimensional Performance Model for Counter-Rotating Propellers

Due to their high propulsive efficiency, counter-rotating open rotors (CRORs) have the potential to significantly reduce fuel consumption and emissions relative to conventional high bypass ratio turbofans. However, this novel engine architecture presents many design and operational challenges both at engine and aircraft level. The assessment of the impact of the main low-pressure preliminary design and control parameters of CRORs on mission fuel burn, certification noise, and emissions is necessary at preliminary design stages in order to identify optimum design regions. These assessments may also aid the development process when compromises need to be performed as a consequence of design, operational, or regulatory constraints. The required preliminary design simulation tools should ideally be 0D or 1D (for computational purposes) and should capture the impact of the independent variation of the main low-pressure system design and control variables, such as the number of blades, diameter and rotational speed of each propeller, the spacing between the propellers, and the torque ratio (TR) of the gearbox or the counter-rotating turbine (CRT), among others. From a performance point of view, counter-rotating propellers (CRPs) have historically been modeled as single propellers. Such a performance model does not provide the required flexibility for a detailed design and control study. Part I of this two-part publication presents a novel 0D performance model for CRPs allowing an independent definition of the design and operation of each of the propellers. It is based on the classical low-speed performance model for individual propellers, the interactions between them, and a compressibility correction which is applied to both propellers. The proposed model was verified with publicly available wind tunnel test data from NASA and was judged to be suitable for preliminary design studies of geared and direct drive open rotors. The model has to be further verified with high-speed wind tunnel test data of highly loaded propellers, which was not found in the public domain. In Part II, the novel CRP model is used to produce a performance model of a geared open rotor (GOR) engine with a 10% clipped propeller designed for a 160 PAX and 5700 NM aircraft. This engine model is first used to study the impact of the control of the propellers on the engine specific fuel consumption (SFC). Subsequently, it was integrated in a multidisciplinary simulation platform to study the impact of the control of the propellers on engine weight, certification noise, and NOx emission.

Influence of aeroelastic flow induced oscillations on fatigue life of an airplane wing structure

In the field of aeroelasticity, the phenomenon of dynamic vibrations and its prevention is a primary challenge that is imposed on an airplane designer. The fully coupled/partly coupled fluid-structure interaction (FSI) analysis is one of the widely used techniques implemented by design industries to assess the characteristics of limit cycle oscillations (LCO). It requires detailed computational modelling capabilities including a high speed wind tunnel testing environment. Hence, the partly coupled (or) moderate loosely coupled (MLC) FSI techniques are preferred to approximate the real phenomenon better than the quasi-steady models. The turbulence characteristics of a viscous flow field are computed efficiently by a reduced-order modelling approach that offers tractable solutions. This novel approach is used to determine the consequences of LCO on the fatigue life of an airplane wing structure. Further, the numerical fatigue simulations are used to demonstrate the effectiveness of MLC algorithm against experimental techniques.

Advanced open rotor noise prediction

AbstractThe purpose of this paper is to describe the current status of open rotor noise prediction methods and to highlight future challenges in this area. A number of analytic and numerical methods are described which can be used for predicting ‘isolated’ and ‘installed’ open rotor tonal noise. Broadband noise prediction methods are also described and it is noted that further development and validation of the current models is required. The paper concludes with a discussion of the analytical methods which are used to assess the acoustic data collected during the high-speed wind-tunnel testing of a model scale advanced open rotor rig.

Reynolds number sensitivity to aerodynamic forces of twin box bridge girder

This paper presents the experimental results for a streamlined twin box girder of 1545m span suspension bridge in order to investigate the Reynolds number sensitivity to aerodynamic force and pressure coefficients. High speed wind tunnel testing was carried out on a 1:30 scale sectional model at an aeronautical wind tunnel. The drag and lift coefficients revealed significant decreases at a critical Reynolds number that was in the range of ordinary wind tunnel tests. The safety rail reduced the Reynolds number dependency of aerodynamic force coefficients at negative angles of attack. Similarly, the boundary layer trip strip attached at the bottom surface of the girder reduced the Reynolds number dependency by fixing the flow separation. The pressure coefficients near the gap between the twin boxes changed from negative to positive at the critical Reynolds number, which represented smooth ventilation of wind flow through the gap at the supercritical region. The imbalance of the drag and lift forces acting on each box in the subcritical region gradually resolved and both boxes were almost equally loaded by the forces at the supercritical region.

A Case Study of Application of Modern Design of Experiment Methods in High Speed Wind Tunnel Test

For the development trend of advanced aircraft wind tunnel tests which require shorter cycle time and lower cost, a new wind tunnel testing process called Modern Design of Experiments (MDOE) is under development at NASA Langley Research Center. This paper presents a case study of application of Modern Design of Experiment methods in high speed wind tunnel tests based on the aerodynamic data of certain type of aircrafts. The full process including experiment design, execution, analysis of variance, and block design is described. Comparisons between MDOE predictions and OFAT measurements were made in the present study. The results show that the MDOE methods have well predictive capability to meet specific test objectives with less data volume, less cost and less time compared to the conventional OFAT methods, which means the MDOE approach established and used in this study is feasible enough to meet specific aerodynamic wind tunnel testing objectives in place of conventional OFAT methods.

A Case Study of Application of Modern Design of Experiment Methods in High Speed Wind Tunnel Test

A Case Study of Application of Modern Design of Experiment Methods in High Speed Wind Tunnel Test

An Investigation on Manufacture Method of High-Speed Wind-Tunnel Test Models Based on Stereo-Lithography Technique

The paper presents a novel manufacture method of high-speed wind-tunnel test models with internal metal frame and surface photopolymer resin based on Stereo-Lithography (SL) technique. Internal metal frame structure was designed to be of regular configurations that can be conveniently fabricated by conventionally mechanical manufacturing methods. Outer resin components were designed to meet configuration fidelity and surface quality, which were fabricated by SL facilities. Combination of aerodynamics and structure was utilized to accomplish structural design, strength and stiffness calibration and vibration analysis. It is validated that the manufacture method is suitable to construct the high-speed wind-tunnel models by some typical wind-tunnel models, especially for test models with complex structure.

The Effects of Different Support Forms on the Aerodynamic Characteristics of the High-Speed Train Model

By the method of computational fluid dynamics (CFD), this paper calculates the aerodynamic parameters of one complex high-speed train model which adopts different support forms when the reference velocity is 70m/s under different operating conditions. It also analyses the support interference mechanism from the point of flow field structure. The results show that the distributed cylinder support form causes least interference on the model, while the single big cylinder support forms change the flow field structures much which leads to much change of the aerodynamic parameters of the model. The distributed cylinder support form can be applied as a good support form for the high-speed train wind tunnel tests.