A 2D Validation Experiment for Dynamic High-Lift System Aerodynamics

Pressure tap cavity for unsteady aerodynamic pressure measurements

An experimental investigation of unsteady pressures on military aircraft wind tunnel models has been performed using Polyvinylidenfluorid (PVDF) sensors. The unsteady pressures from the wind tunnel measurements are a prerequisite for the prediction of dynamic loads and vibrations during the development and design of new fighter aircraft structures. In the past the wind tunnel unsteady pressure measurements had been performed using wind tunnel models equipped by Kulite pressure sensors. This technique resulted in very high cost wind tunnel models. In contrast the application of the PVDF foil Smart Sensor & Signal Processing Technology would lead to improvements through the application of an affordable test technique which could be also applied in flight with the benefit of more accurate and locally detailed pressure information. The present investigation includes the measurement of unsteady pressures with PVDF foils on a wind tunnel model of a trainer aircraft configuration in the high speed transonic wind tunnel of the NLR Amsterdam. The investigated Mach numbers were M0.5, M0.7 and M0.9 and the incidence range was 0 to 45 degrees. The results of the experiments are discussed using the analysed time histories and power spectral densities of the stochastic unsteady pressure and comparisons to results using direct unsteady pressure measurements. Main aspect is the evaluation and validation with respect to technical industrial applicability of PVDF foils.

The Ohio State 6” × 22” Unsteady Transonic Wind Tunnel was used to study different methods of experimentally capturing the time-varying location of a shock wave on a NACA 0012 airfoil. The unsteady wind tunnel enables a freestream Mach number oscillation of 0.51 ± 0.1 for a Reynolds number range of 17 – 43 million per meter. Fast-response pressuresensitive paint (PSP) and particle image velocimetry (PIV) were evaluated for their ability to capture time-accurate location of the oscillating shock wave. Unsteady surface pressure and temperature measurements were made using a fast-response bi-luminophore pressuresensitive paint with a single-shot intensity-based technique. This method allows for high spatial resolution surface pressure and temperature maps. Phase-locked planar particle image velocimetry data were also acquired in this experiment, demonstrating an ability to resolve a moving shock. These advanced diagnostic tools enable detailed understanding of the impact of time-varying compressibility both on and off the airfoil surface. The information obtained with each of these techniques is compared in terms of spatial resolution and accuracy with data acquired via surface pressure taps. A NACA 0012 airfoil was tested at 9, 10 and 11 degrees angle of attack and at low reduced frequencies.

Experimental Fluid Dynamics (EFD) and Computational Fluid Dynamics (CFD) have been instrumental in Fluid Mechanics to help solve scientific and engineering problems. This research attempts to use both techniques to perform a parametric study of turbulence flow around airfoil ClarkY-14 at various velocity and angle of attack (AoA). Clark Y-14 airfoil was designed in the 1920’s. It demonstrated good overall performance at low and moderate Reynolds numbers. With the progress in the aviation field, its performance was sub-optimal for newer aircraft designs. However, with the advent of RC airplanes and model aircrafts, there is a renewed interest in this airfoil. Various research projects have been conducted using this airfoil, but there hasn’t been a combined EFD and CFD study of the performance characteristics of the airfoil itself, which still finds real world applications today. One important aspect of this research included the investigation of the effects of a Force Measurement Device/Sensor, which is typically used in scaled/full-size wind tunnels to mount the test model as well as measure the forces/moments acting on it during the testing. The presence of such a device could affect the quality of the data obtained from the wind tunnel testing when compared to a real world application scenario where the aforementioned device may not be present. To the best of the author’s knowledge, no detailed study has been published on the effects of such devices. In this study, the results with and without the measuring device were generated by using CFD simulations. The results were then compared to see to what extent the inclusion of these devices will affect the results. The methodology used for this research was experimental as well as computational. In the present research, a commercially available CFD software STAR-CCM+ was employed to simulate the flows around airfoil Clark Y-14. The experimental data was obtained from wind tunnel tests using AEROLAB Educational Wind Tunnel (EWT) and compared with the simulation data from the CFD. The two data sets were in good agreement. Both experimental and simulation results were used to understand the effects of the measurement device/sensor used in the scaled wind tunnel on the lift and drag coefficients of the airfoil. Two separate CFD simulation setups were designed to model the presence and absence of the measurement device/sensor. These setups replicated the wind tunnel setup. The airfoil was tested and simulated at different speeds as well as different AoA. The comparative study gave a useful insight on the accuracy of the CFD simulations in relation to the actual testing. The analysis of results concluded that the force measurement device/sensor had insignificant effects on the accuracy and quality of data collected through wind tunnel testing.

It is often imperative to test and validate the development of active flutter suppression technology in wind tunnels or flight near and beyond the flutter speed. This is done not only to fine-tune mathematical models but also to examine the active control laws that were synthesized from such models. However, in-flight flutter tests are often dangerous, costly, and sometimes infeasible. Wind tunnel tests are usually more affordable and less risky leading to higher productivity in evaluating alternative configurations and control law strategies. The University of Washington’s MARGE-I aeroservoelastic wind tunnel model is a flexible half-wing-body-tail model with active ailerons and elevator. The work here studies SISO mixed-sensitivity H∞ control for active flutter suppression, which is designed subject to requirements and hardware constraints. Such controller is validated by wind tunnel tests, where the closed-loop margins in both simulation and experiments are quantified and compared. A previously published safe wind tunnel and flight test technique is used to generate flutter mechanisms. The set of control surfaces is partitioned into destabilizing and stabilizing ones, where a destabilizing controller reshapes the aeroservoelastic system to yield the desired unstable dynamic characteristics, and the stabilizing controller is used to study active flutter control strategies.

Comparison of wind tunnel and full-scale pressure measurement tests on the Texas Texh Building

Comparison of wind tunnel and full-scale pressure measurement tests on the Texas Texh Building

This paper presents detailed study of force and pressure coefficients of an unsymmetrical ‘E’ plan shaped tall building under wind excitation. The building is unsymmetrical about both direction (along and across wind direction) in plan. Experimental and numerical study is carried out by wind tunnel test in boundary layer wind tunnel and Computational Fluid Dynamics (CFD) technique using ANSYS CFX software respectively to study the above model under wind excitation. Two different types of models were used for force and pressure measurements in wind tunnel. Rigid model was used at a model scale of 1:300 for both the cases. Force measurement model was tested under three different wind speeds of 6, 8, 10 m/s whether pressure measurement model was tested under wind speed of 10 m/s in wind tunnel. Wind flow pattern around the building model is studied by CFD technique for both wind flow direction which predicts about the nature of pressure distribution on different surfaces. Some dynamic nature of wind flow is observed at the rear side in case of across wind effects. Maximum positive force and pressure coefficients noticed in case of along wind effects as the surface area is maximum and maximum wind energy dissipates on frontal surface. Same mean pressure occurred on similar faces in case of along wind effects. Pressure contours are symmetrical about vertical axis on frontal faces for both directions of flow and pressure fluctuation noticed at the rear side face due to dynamic effects of flow pattern in case of across wind effect. A comparative study is carried out between experimental and numerical study and results found from these two methods have good agreement to each other.

The combined trends toward internal carriage of weapons in modern aircraft and the design of new weapons to be neutrally stable for in creased performance have increased the probability that unsteady aerodynamics will effect store separation trajectories. Reports from time accurate computational fluid dynamics trajectory simulations have shown trajectory sensitivity to time of weapon release. Typical wind tunnel data collected to support store separation analysis cannot detect thi s effect since the data collected consists of time averaged store loads. Flight testing to inves tigate effects of bay unsteadiness has not been done since it is expensive and typically there are insufficient funds to conduct repeated store drops at identical flight conditions. Carefu lly designed and conducted small scale drop testing in wind tunnels provides a means to identif y the effect of unsteady weapons bay aerodynamics on store separation trajectories. Time accurate computational fluid dynamics trajectory analysis results on the Small Smart Bomb separating from an F-111 aircraft and a GBU-12 separation from the B-52 aircraft are reviewed. Small scale drop test data from a generic 10% scale rectangular cavity in a flatplate and from a 6% scale B-1B are introduced to provide experimental evidence of the effect of u nsteady flow on the store separation trajectory.

The objective of this paper is to analyze the 3D buffet phenomenon which appears on a swept wing at a high Mach number and/or high angle of attack. This aerodynamic instability induces strong wall pressure fluctuations and as such limits aircraft envelope. Consequently, it is interesting to understand this phenomenon in order to not only improve aircraft performance but also to provide more flexibility during the design phase. Results from two wind tunnel tests on a 3D half wing-body configuration are presented for several freestream Mach numbers (0.78–0.86) and Reynolds numbers (2.83 × 106–8.49 × 106, based on the aerodynamic mean chord). The buffet phenomenon is characterized using steady and unsteady wall pressure measurements. By opposition to the 2D buffet which exhibits rather a well marked peak in the pressure spectra, the 3D buffet is characterized by a broadband bump at a much higher Strouhal number (between 4 and 7 times higher). It is also observed that two different instabilities coexist on the suction side of the wing: the 3D buffet phenomenon (with Strouhal numbers ranging between 0.2 and 0.6) and the Kelvin-Helmholtz instability (with Strouhal numbers ranging between 1 and 4). Each phenomenon has a different Strouhal number but also different convection velocities and propagation directions.

This paper presents results of wind tunnel and acoustic tests to investigate buffet loads on Shuttle Thermal Protection System (TPS) tiles. It also describes the application of these results to the prediction of tile buffet loads for the first Shuttle flight into orbit (STS-1). The tests were conducted in wind tunnels at transonic and supersonic Mach numbers and in a progressive wave tube. The wind tunnel tests included single tiles of various thicknesses and arrays of tiles mounted on both rigid and elastic substructures. The corresponding acoustic tests were of the single tiles only. In most cases the “test” tiles were of Orbiter qualified material and they were bonded to the panels and proof-loaded in accordance with flight-vehicle specifications. The fixtures for the panel installations in the wind tunnels were designed to simulate flow regions on the Orbiter where shock waves and related high pressure gradients and high dynamic loads would occur. The results of the wind tunnel tests indicate that significant tile buffet loads occurred due to separated flows and oscillating shocks. Separated flow excitation caused higher buffet loads than oscillating shocks for a given level of excitation, however, the highest combined steady and dynamic loads occurred when shock waves with ΔP > 1.5 psi were on tiles.

Wind energy, being an indirect form of solar energy would initially seem a very promising form of energy. Unfortunately, it suffers from the problem of dilution. Wind turbine designers naturally try to compensate for this by increasing the size of the rotor to capture more of the kinetic energy of the wind. A major constraint in conventional wind turbine design is the reduction in rotational speed as the size of the rotor is increased. This means expensive gear boxes are unavoidable. The rotor also becomes considerably more complicated in design and heavier as the size increases, to mitigate working stresses. Flow concentrators have been investigated in an attempt to alleviate wind turbine design problems, but flow concentrators usually incur the expense of high structural weight and size. The Helical Vortex Wind Concentrator (HVWC) is a recent addition to the list of flow concentrator types and its economic potential is, as yet unknown. The principle of the HVWC has been demonstrated in a series of wind tunnel tests. The wind tunnel tests involved a direct comparison between the performance of a wind turbine with and without an HVWC attached.During these tests a definite increase in power out was observed when the concentrator was attached to the wind turbine. Previous to these successful wind tunnel tests, other wind tunnel and field tests had been conducted on less successful designs. These other tests were important in the development of the current theory and design or the HVWC. Future research will need to investigate both physical and economic limitations of this type of wind concentrator.

Acoustic wind tunnel test is an important way to study the noise mechanism of helicopter rotors. The broadband dipole noise distribution of helicopter rotors can be obtained from the acoustic wind tunnel measurement, which can be extremely helpful in analyzing the pulsation noise mechanism of rotor blade surfaces. However, during acoustic imaging of broadband dipole sources, there is often a problem of unwanted noise interference from other experimental setups of the wind tunnel, such as the rotor shaft. To accurately visualize the broadband dipole sound sources, in this paper, a moving dipole-type sources beamforming method suitable for the wind tunnel environment is proposed. The reduced-rank cyclic Wiener filter is firstly developed to extract the rotor broadband noise signal in wind tunnel tests; a beamforming method for moving dipole sound sources is then proposed according to the characteristics of the generation and propagation of the dipole sound sources. The BO105 helicopter rotor model is tested in the FL-17 aeroacoustic wind tunnel of the China Aerodynamics Research and Development Center (CADRC). The wind tunnel test results demonstrate that the method can effectively suppress noise interference in the wind tunnel and accurately extract and visualize broadband dipole sources of rotor blade surfaces.

A complete set of unsteady blade surface pressure measurements is presented for a single turbine blade oscillating in a three dimensional bending mode. Results are provided for five spanwise sections at 10%, 30%, 50%, 70% and 90% of span. Steady blade pressure measurements and five-hole probe traverses at the inlet and exit planes of the test section, are also included. The test facility operates at low speed and the working section consists of a single turbine blade mounted in a profiled duct. A rigid blade with constant section was used, and a three dimensional bending mode realised by hinging the blade at root and driving the tip section. The low speed and scale of the test facility allowed low oscillation frequencies (5 to 20 Hz) to be employed, in order to match realistic reduced frequencies. This enabled the unsteady blade surface pressure response to be recorded with externally mounted pressure transducers. The validity of this technique is examined. Results from the test facility demonstrate a noticeable three dimensional behaviour of the unsteady flow.

PurposeSome recent effort showed that usage of Krueger flaps helps to maintain laminar flow in cruise flight. Such flaps are positioned higher relative to the chord to shield the leading edge from the insect contamination during take-off. The flap passes several through critical intermediate position during the deployment to its design position. The purpose of this paper is to analyse the aerodynamics.Design/methodology/approachTo better understand such flow phenomena, the combined approach of computational fluid dynamics and experimental methods were used. Flow simulation was performed with in-house finite volume Navier–Stokes solver in fully turbulent unsteady RANS regime. The experimental data were obtained by means of force and pressure measurements and some areas of the flow field were examined with 2 C particle image velocimetry.FindingsThe airfoil with flap in critical position has a very limited maximum lift coefficient. The maximum achievable lift coefficient during the deployment is significantly affected by the vertical position of the trailing edge of the flap. The most unfavourable position during the deployment is not the flap perpendicular to the chord, but the flap inclined closer to it is the retracted position.Research limitations/implicationsThe flap movement was not simulated either in the simulation or in the experiment. Only intermediate static positions were examined.Practical implicationsA better understanding of aerodynamic phenomena connected with the deployment of a Krueger flap can contribute to the simpler and lighter of kinematics and also to decrease time-to-market.Originality/valueLimited experimental and computational results of Krueger flap in critical positions during the deployment are published in the literature.