Trailing Edge Pressure Research Articles

Aerodynamic analysis of airfoil in steady and unsteady motion

In this paper, several basic methods for solving the wing force are introduced, namely: arranging constant vortices on the surface, linear vortices, arranging sources and sinks on the surface, and arranging vortex systems on the camber. The focus is on the last method. For thinner airfoils, the block form has a greater impact on the results; while for thicker symmetrical airfoils, the block form has little effect on the results. The results can be more accurate if the number of blocks is appropriately increased. The distribution vortices with higher order should be used on the camber surface, such as the plate vortex system and linear vortex system. Moreover, the force of the unsteady motion wing is also discussed, and the method of arranging the source and sink on the surface of the wing and the vortex system on the camber is used. For the wing with unsteady motion, the Kutta condition and the wake vortex system should be dealt with. Since the square term of the velocity term is included, the Kutta condition with equal trailing edge pressure cannot be directly adopted, and it needs to be improved. Regarding the treatment of the wake vortex, the calculation results show that the first free vortex adopts the vortex model, and the result is better. For the same airfoil, the calculation results are not very different under different time steps. Therefore, in order to improve the calculation efficiency, it is appropriate to consider increasing the time step.

Numerical analysis of pitch-break and all moving wingtip aileron of lambda wing configuration

Lambda wing, as a commonly used configuration for modern combat aircraft, tend to reduce stability margin known as Pitch-Break before the wing completely stalls. This phenomenon greatly affected the safety of aircraft at high angles of attack. In this paper, the characteristics of Pitch-Break are clarified and it is found that the initial airflow separation position of the outer wing is the principal reason for generating nonlinear moment of lambda wing. The airflow separation position is generally affected by leading edge vortex and the trailing edge spanwise airflow. Furthermore, the following conclusions are obtained: 1. The distribution of the absolute value of the leading edge radius is a decisive parameter of the initial position of leading edge vortex at the beginning of Pitch-Break zone; 2. The difference of trailing edge pressure gradient brings difference of trailing edge spanwise airflow at the end of Pitch-Break zone. To solve the aerodynamic problem of Pitch-Break, this paper adopts an all moving wingtip (AMT) as auxiliary aileron which has minor fluctuations during the Pitch-Break zone compared with conventional aileron. The result shows the defect of the conventional aileron pitch control at high angle of attack and presents the engineering feasibility of the all moving wingtip auxiliary aileron.

Control strategies for tip leakage vortex using inclined squealer rim in axial turbines

In a typical gas turbine, due to its complicated blade geometry, complex vortex structures appear and cause significant aerodynamic loss. Vortex systems dominated by a tip leakage vortex near the tip region are the primary source of this loss. In this paper, to improve the aerodynamic performance of the turbine, two novel control strategies for tip leakage vortex and the tip leakage flow of the cavity tip are proposed, and their coupling control effects are numerically investigated. The first control strategy is intended to control the loss caused by the breakdown of tip leakage vortex. By inclining the external wall of the suction side rim toward the passage, the emergence of a trailing edge pressure spike is delayed. This significantly reduces the adverse pressure gradient, suppressing the breakdown of tip leakage vortex and reducing tip leakage loss. The second control strategy controls tip leakage flow using the inclined inner wall of the suction side rim, which enhances the separation bubble on the top of the rim of the suction side and reduces the leakage rate by 7.7%. In this way, the formation and development of tip leakage vortex are indirectly manipulated, inhibiting the tip leakage loss. The coupling of the two strategies reduces the blocking effect on tip leakage flow slightly compared to the second strategy. However, the stage efficiency of the turbine is still improved by 0.24% because of the effective suppression of tip leakage vortex breakdown.

Wind peak pressures on a square-section cylinder: Flow mechanism and standard/conditional POD analyses

This study emphasizes the flow mechanism of negative peak pressures on a square-section cylinder at Reynolds number of 22,000, particularly of those immediately upstream of the trailing edges (trailing-edge peak pressure). The minimum trailing-edge pressure coefficients are observed as low as -5 and have a close correlation in a streamwise length of approximately 0.1D (D is the width of the cylinder). The occurrence of small trailing-edge vortex, which has highly concentrated vorticity, is responsible for the trailing-edge peak pressures. The trailing-edge peak pressure occurs immediately before the shear layer moves far from the cylinder and recovers to the symmetric state. Standard proper orthogonal decomposition (POD) is applied to discover the inherent spatial modes with sharp peak pressures upstream of the trailing edges. Instead of the 1st mode (corresponding to vortex shedding), the higher-order modes contribute significantly to the trailing-edge peak pressures, whose mode coefficients have harmonics of vortex-shedding frequency. More than nine POD modes are necessary to reconstruct the trailing-edge peak pressures. As a variant of POD proposed recently, conditional space–time POD is valid in studying the average evolution of wind peak pressures. It further provides the relative phase and frequency ratio between the trailing-edge peak pressure and vortex shedding.

Experimental investigation of shock-buffet criteria on a pitching airfoil

An experimental investigation of the shock-buffet phenomenon subject to unsteady pitching supercritical airfoil around its quarter chord has been conducted in a transonic wind tunnel. The model was equipped with pressure taps connected to the fast response pressure-transducers. Measurements were conducted at different free-stream Mach number from 0.61 to 0.76. The principle goal of this investigation was to experimentally discuss the shock-buffet criterion over a SC(2)-0410 supercritical pitching related to the hysteresis loops of total drag and trailing edge pressure, the behaviour of the shock wave foot location, the pressure distribution over the upper surface, and by implementing the wavelet analysis of the normal force. To ensure capturing the buffet phenomenon by utilizing these criteria, a pressure port has been drilled exactly at the trailing edge of the airfoil where its output was used to detect the buffet phenomenon for different conditions. Visual representation of the flow using the shadow graph flow visualization technique for different test cases is further used to illustrate the unsteady shock wave motion. A comparative analysis of experimental measurements shows that the conducted criteria confirm each other when the buffet phenomenon occurs at the position of the oscillating cycle.

Multi-axial large-scale testing of a 34 m wind turbine blade section to evaluate out-of-plane deformations of double-curved trailing edge sandwich panels within the transition zone

Transverse cracks in the double curved trailing edge panels within the transition zone are among one of the increasingly encountered in-field damages found on wind turbine blades today. Believed to be root cause of these transverse cracks, are the out-of-plane deformation of the double curved trailing edge pressure side panels. These deformations are evaluated on the inner 15 m section of a 34 m wind turbine blade – referred to here as the root section. Through a parametrical study the free end of the root section is loaded in the quasi-static regime comprising edgewise loading (Fy) and torsional moment (Mz) around the longitudinal axis of the blade. The root section is through a multi-scale numerical analysis found to exhibit representative structural behavior in terms of out-of-plane deformations within the area of interest. A combination between Fy and Mz are found to generate the highest peak-to-peak out-of-plane deformation of 15.9 mm.

Finite Element Vortex Method for Hydrodynamic Analysis of Vertical Axis Cycloidal Tidal Turbine

Jiang, J.; Ju, Q., and Yang, Y., 2019. Finite element vortex method for hydrodynamic analysis of vertical axis cycloidal tidal turbine. In: Guido-Aldana, P.A. and Mulahasan, S. (eds.), Advances in Water Resources and Exploration. Journal of Coastal Research, Special Issue No. 93, pp. 988–997. Coconut Creek (Florida), ISSN 0749-0208.A finite vortex method was developed based on the Kutta condition of pressure difference at the trailing edge of an airfoil aiming at the analysis and prediction of hydrodynamic load and wake field of cycloidal vertical axis tidal turbine. The empirical value of the trailing edge pressure difference parameter was obtained by comparing with the experimental value of static lift coefficient of NACA0015 airfoil. The instantaneous and average loads of fixed pitch turbine and cycloidal turbine with different eccentricity are compared. The results show that the positive torque of the turbine is mainly concentrated in the upstream area of the actuator disk. Cycloidal turbine can not only improve the power coefficient, but also reduce the force's amplitude of the blade. By comparing the structure and evolution process of turbine wake, the sources of the trailing vortex in the upper and lower regions are found. They are obtained from the starting vortex and the vortex flipped off by the blades, respectively. The wake of a turbine with high power coefficient develops more rapidly downstream.

Experimental and numerical investigations of trailing edge injection in a transonic turbine cascade

Trailing edge mixing flows associated with coolant injection are complex, in particular at transonic flows, and result in significant aerodynamics losses. A series of tests and calculations on a high-pressure turbine vane cascade with trailing edge injection were conducted to investigate the influence of trailing edge injection on the loss characteristics of the transonic profile. Wake traverses with a five-hole probe and measurements of the pressure distribution on the profile were taken for injection mass flow ratios of 0 and 4.1% under a test Mach number of 0.7. In the meantime, two-dimensional numerical predictions were carried out for exit isentropic Mach numbers of 0.7, 0.85 and 1.1 and injection mass flow ratios of 0, 4.1%, 4.7% and 5.2%. Numerical predictions show a reasonable agreement with the experimental data, and wake total pressure losses and flow angles as well as pressure distributions on the profile were compared to calculations without trailing edge injection, showing a significant effect of trailing edge injection on the wake development and its blockage effect on the flow in the vane cascade passage. The outlet total pressure loss and flow angle are increased with increasing exit Mach number. The trailing edge injection of 4.1% increases the static pressure in the rear part of vane suction side, and then reduces the vane loading, thus causing decreased wake losses at different exit Mach numbers. At transonic flows, the injection also reduces the strength of the shock wave at the trailing edge pressure side. In addition, the injection decreases the outlet pressure circumferential-unevenness at an exit isentropic Mach number of 0.7, while it increases the pressure circumferential-unevenness at higher exit Mach numbers.

CFD Simulations of the New MEXICO Rotor Experiment under Yawed Flow

This paper describes the conduction of CFD simulations with DLR’s flow solver TAU for comparison with the experimental data that has been used within the third phase of the MexNext-III project. More precisely computations for three different inflow velocities under yawed flow condition have been conducted and compared to the experimental results. Important aspects of the simulations such as a sufficiently long simulation time or a sufficiently small time-step are discussed. At first local flow quantities such as the flow velocity along axial and radial traverses, pressure distributions at four radial stations and normal and tangential forces at these sections are compared. Subsequently the integral rotor quantities i.e. rotor thrust and rotor torque are revised. The integration of the CFD results is performed using the complete surface dataset as well as two subsets. The comparison proves that the agreement can substantially be improved if the subset resembles the experimental dataset. Furthermore the integration of the experimental dataset is reviewed and a weak spot in the current reconstruction procedure of the trailing edge pressure is identified. The proposed alternative provides evidence that further improvements are possible. The final comparison shows an acceptable agreement for the medium and high inflow speed case yielding 0% and 10% difference for integral thrust and 9% and 14% for integral torque. The results for the low inflow speed case are divers: while the agreement for the integral thrust is excellent, it is less satisfactory for the integral torque which amounts to 41%.

Experimental and numerical investigations of hole injection on the suction side throat of transonic turbine vanes in a cascade with trailing edge injection

Trailing-edge mixing flows associated with coolant injection are complex, in particular at transonic flows, and result in significant aerodynamics losses. The objective of this paper is to evaluate the impacts of hole injection near the suction side throat on shock wave control and aerodynamic losses. A series of tests and calculations on effects of hole injection on the suction-side throat of a high-pressure turbine vane cascade with and without trailing-edge injection were conducted. Wake traverses with a five-hole probe and tests of pressure distributions on the turbine profile were taken for total injection mass flow ratios of 0% and 1.2% under test Mach numbers of 0.7, 0.78, and 0.87. Meantime, numerical predictions are carried out for exit isentropic Mach numbers of 0.7, 0.78, 0.87, and 1.1 and hole-injection mass flow ratios of 0%, 0.17%, 0.3%, and 0.89%. Numerical predictions show a reasonable agreement with the experimental data, and wake total pressure losses and flow angles as well as pressure distributions on the turbine profile were compared to calculations without hole injection, indicating a significant effect of hole injection on the profile wake development and its blockage effect on the shock-wave flow in the vane cascade passage. At subsonic flows, the hole injection on the suction side throat thickens the suction-side boundary layer, and increases the flow mixing, thus causing increased wake losses and flow angles. At transonic flows, while the trailing-edge injection reduces the strength of the shock wave at the trailing-edge pressure side, the hole injection on the suction side throat alters the local pressure fields, and then tends to enhance the shock-wave at the trailing-edge pressure-side; however, it seems to reduce the strength of the shock-wave at the trailing-edge suction side.

Burst-Mode Frequency Effects of Dielectric Barrier Discharge Plasma Actuator for Separation Control

The various separation control mechanisms of burst-mode actuation with a dielectric barrier discharge plasma actuator were experimentally investigated in this study. The control of the separated flow around a NACA 0015 airfoil at a Reynolds number of was investigated using a plasma actuator mounted at a distance from the leading edge of 5% of the chord length. A parametric study on the nondimensionalized burst frequency was conducted at three poststall angles of attack and various input voltages using time-averaged pressure measurements and time-resolved particle imaging velocimetry (PIV) results. The measurement results of the trailing edge pressure, which was selected as the index of separation control, indicate that the optimal burst frequency varies with the angle of attack. Several flow fields are discussed in detail in this paper, and two flow control mechanisms were observed: the use of a large-scale vortex and the promotion of turbulent transition. With regard to the first mechanism, the phase-locked PIV results indicate that a vortex structure, the size of which increases with decreasing burst frequency in the experimental range, is shed from the shear layer for each burst actuation. With regard to the second mechanism, time-averaged pressure and PIV measurements reveal that a burst frequency of is able to promote turbulent transition. Among these two mechanisms, at higher angles of attack, the use of a large-scale vortex structure provides better separation control, whereas near the stall angle, the promotion of the turbulent transition provides better separation control.

Film Cooling Effectiveness Downstream of Trailing Edge Slots Including Cutback Surface Protuberances

The trailing edge of the high pressure turbine blade and vane presents significant challenges to the turbine cooling engineer. The current research has focused specifically on the effect of cutback surface protuberance, or “land”, shapes on film cooling effectiveness. A set of six different land geometries has been investigated in a large scale model of the trailing edge pressure side ejection slot exit. Slot height and width and lip height was maintained. Pressure sensitive paint was used to measure adiabatic film cooling effectiveness at five blowing ratios ranging from 0.6 to 1.4 in increments of 0.2. High-resolution full surface distributions of film cooling effectiveness both on the cutback surface and the top of the lands were recorded. It was found that tapering the lands did not significantly increase effectiveness on the lands and slightly reduced effectiveness near the lands. Using a diffuser shape improved average effectiveness greatly and gave the best overall performance up to the end of the lands except at the lowest blowing ratio of 0.6, where having no lands was slightly better.

Prediction of the unsteady turbine trailing edge wake flow characteristics and comparison with experimental data

The trailing edge base pressure is known to play a significant role in the profile losses of turbine blades in particular at transonic outlet Mach numbers. Recent tests have shown that the assumption of an isobaric near wake region is wrong because of the highly unsteady character of the near wake flow which affects directly the trailing edge base pressure. New experimental data show the evolution of the trailing edge base pressure in function of the downstream Mach number from a uniform base pressure at moderate subsonic Mach numbers to an increasingly strong non-uniform distribution up into the transonic range followed by a sudden return to an isobaric base trailing edge pressure. To get more insight into the unsteady flow features, numerical simulations are carried out. A LES approach is validated and employed in this work to analyze three different operating conditions for the profile, representing three configurations of the flow and the trailing edge shock system.

An experimental study of buffet detection on supercritical airfoils in transonic regime

Conventional buffet onset methods for a 2D supercritical airfoil, SC0410, in transonic regime for various Mach number and various angles of attack have been surveyed. The existing methods give good results for high subsonic and transonic regimes, but demand a computational procedure to detect the buffet onset. One of these methods, trailing edge pressure divergence, that have been recognized inappropriate in other studies for supercritical airfoils, shows acceptable result at least for the present supercritical airfoil. A new method has been proposed by the authors for transonic regime that is based on the physical definition of the buffet onset from the surface pressure distribution diagram. This method does not require any special calculations. The data scrutiny shows good agreement by this method in comparison with the conventional schemes.

Reynolds Number Effects Investigation of Supercritical Airfoil Based on EFD and CFD

This study aimed to investigate the transonic Reynolds number effects of supercritical airfoil by EFD and CFD method. An experiment was conducted in NF-6 wind tunnel, to obtain the pressure distribution and aerodynamic coefficients of a typical supercritical airfoil through pressure measuring, with Reynolds numbers varied from 3.5×10 6 to 1.0×10 7 per airfoil chord, Mach numbers from 0.6 to 0.8, angles of attack from 0° to 8°. Also, flows over the supercritical airfoil were numerically studied; the two-dimensional Navier-Stokes equations were solved with structure grids by utilizing the Spalart-Allmaras (S-A) turbulence model, with Reynolds numbers varied from 2.0×10 6 to 50×10 6 per airfoil chord and Mach numbers from 0.6 to 0.8. Computational results compared well with experimental results. It is shown that the upper surface pressure distribution of supercritical airfoil including the location and intensity of shock wave and trailing-edge pressure coefficient, changed apparently with variable Reynolds numbers, when shock-induced trailing-edge separation existed. It is also noticed that the lift coefficient increased, drag and pitching moment coefficient decreased as Reynolds number increasing. Results implied that Reynolds number effects should be considered during the early designing stage and optimization of large aircrafts applied supercritical airfoil.

Engineering Extrapolation to Flight Reynolds Number for Supercritical Airfoil Pressure Distribution Based on CFD Results

Pressure distribution of supercritical airfoil at flight Reynolds number could not be fully simulated except in cryogenic wind tunnel such as NTF (National Transonic Facility) and ETW (European Transonic Wind tunnel), which is costly and time resuming. This paper aimed to explore an engineering extrapolation to flight Reynolds number from low Reynolds number wind tunnel data for supercritical airfoil pressure distribution. However, the extrapolation method requiring plenty of data was investigated based on the CFD results for the reason of low cost and short period. Flows over a typical supercritical airfoil were numerically simulated by solving the two dimensional Navier-Stokes equations, with applications of ROE scheme spatial discretization and LU-SGS time march. Influence of computational grids convergence and turbulent models were investigated during the process of simulation. The supercritical airfoil pressure distribution were obtained with Reynolds numbers varied from 3.0×106 to 30×106 per airfoil chord, angles of attack from 0 degree to 6 degree and Mach numbers from 0.74 to 0.8. Simulated results indicated that weak shock existed on the upper surface of supercritical airfoil at cruise condition, that the shock location, shock strength and trailing edge pressure were dependent of Reynolds number, attack angles and Mach numbers. A similar parameter describing the Reynolds number effects factors was obtained by analyzing the relationship of shock wave location, shock front pressure and trailing edge pressure. Based on the similar parameter, airfoil pressure distribution at Reynolds number 30×106 was obtained by extrapolation. It was shown that extrapolated result compared well with simulated result at Reynolds number 30×106, implying that the engineering method was at least promising applying to the extrapolation of low Reynolds number wind tunnel data.

Engineering Extrapolation to Flight Reynolds Number for Supercritical Airfoil Pressure Distribution Based on CFD Results

Pressure distribution of supercritical airfoil at flight Reynolds number could not be fully simulated except in cryogenic wind tunnel such as NTF (National Transonic Facility) and ETW (European Transonic Wind tunnel), which is costly and time resuming. This paper aimed to explore an engineering extrapolation to flight Reynolds number from low Reynolds number wind tunnel data for supercritical airfoil pressure distribution. However, the extrapolation method requiring plenty of data was investigated based on the CFD results for the reason of low cost and short period. Flows over a typical supercritical airfoil were numerically simulated by solving the two dimensional Navier-Stokes equations, with applications of ROE scheme spatial discretization and LU-SGS time march. Influence of computational grids convergence and turbulent models were investigated during the process of simulation. The supercritical airfoil pressure distribution were obtained with Reynolds numbers varied from 3.0×106to 30×106per airfoil chord, angles of attack from 0 degree to 6 degree and Mach numbers from 0.74 to 0.8. Simulated results indicated that weak shock existed on the upper surface of supercritical airfoil at cruise condition, that the shock location, shock strength and trailing edge pressure were dependent of Reynolds number, attack angles and Mach numbers. A similar parameter describing the Reynolds number effects factors was obtained by analyzing the relationship of shock wave location, shock front pressure and trailing edge pressure. Based on the similar parameter, airfoil pressure distribution at Reynolds number 30×106was obtained by extrapolation. It was shown that extrapolated result compared well with simulated result at Reynolds number 30×106, implying that the engineering method was at least promising applying to the extrapolation of low Reynolds number wind tunnel data.

Variable Reynolds Number Experimental Study on Aerodynamic Characteristic of Supercritical Airfoil RAE2822

An experimental study of pressure distributions over RAE2822 airfoil in the two-dimensional test section 0.8×0.4 meter of a transonic wind tunnel which is the first pressruized continuous wind tunnel in China is presented. This paper in order to further study the influence of the dynamic of continuous changes Reynolds number at Mach number is 0.66 and 0.80, and the attack angle is from-2 degree to 10 degree, and especially the Reynolds number range from3.0×106to 12×106. The study is focalized on the subsonic range of flow conditions with separation and shock wave in the boundary layer. The influence of pressure distribution and pressure coefficient and moment coefficient caused by Reynolds number increasing are analyzed and discussed. The conclusions showed that the pressure distribution of the lower surface of the airfoil get the influence of the Reynolds number is negligible. The Reynolds number impact on the pressure distribution is faintness at Ma=0.66. Reynolds number increases affect the airfoil central and trailing edge pressure. As the Reynolds number increases, the CL curve move and the gradient increasing. The moment coefficient decreased as the Reynolds number increasing. The CL curve with Cd curve moves left as Reynolds number increasing.

A Maximum Curvature Method for Predicting Transonic Buffet Onset Based on Steady Aerodynamic Parameters

It is difficult to predict the buffet onset by kink analysis in steady aerodynamic parameter curves such as the lift coefficient curves, pitching moment coefficient curves and trailing edge pressure coefficient curves at a certain flow conditions. Thus, a maximum curvature method is proposed which is based on the aerodynamic parameter data, to find the buffet onset point at which the curve curvature achieves an extreme. The method is to fit a fourth order polynomial using the aerodynamic parameter data calculated by steady N-S equation. Subsequently, curvature analysis is carried on with the fitted curve and the stationary point in a particular interval of the curve is obtained. And this point is regarded as the buffet onset point. This method can be applied to predict the transonic buffet onset for airfoils based on steady aerodynamic parameters and good agreement is achieved with the comparison of predicted results and wind tunnel test data.

Numerical Investigation on the Reynolds Number Effects of Supercritical Airfoil

Supercritical airfoil has been widely applied to large aircrafts for sake of high aerodynamic efficiency. The aerodynamic characteristics of supercritical airfoil are rather sensitive to Reynolds number. Therefore, Reynolds number has great effects on the economy, comfort and even safety of large aircrafts. Flows over a typical supercritical airfoil XY are numerically investigated for different Reynolds numbers in this paper; the two-dimensional Navier-Stokes equations are solved with structure grids by utilizing the Spalart-Allmaras (S-A) turbulence model. Computational results of RAE2822 airfoil compare well with wind tunnel results. The computation Reynolds numbers of XY airfoil vary from 2.0×106 to 50×106 per airfoil chord while mach numbers equal 0.74 and 0.8. It is shown that the upper surface pressure distribution including the location and intensity of shock wave and trailing-edge pressure coefficient, changed apparently with variable Reynolds numbers, when shock-induced trailing-edge separation exists. Results implied that Reynolds number effects should be considered while designing and optimizing large aircrafts applied supercritical airfoil.