Modified calibration technique of a five-hole probe for high flow angles

A new method is described for calibrating a five-hole probe for extending the useful measurement range up to flow angularities of 85°. The calibration method involves adjustment of the calibration coefficients to allow valid calibration at larger flow angles. The extended range calibration curves for flow angularity, total and static pressures are presented. The present range is valid in pitch only when the yaw ports are nulled.     

Smooth leading edge transition in hypersonic flow

The boundary layer transition along the attachment line of a smooth swept circular cylinder in hypersonic flow is investigated in a blowdown wind tunnel. A wide range of spanwise Mach numbers Me (3.28 to 6.78) is covered with the help of different models at several sweep angles (60°?Λ?80°). The transition is indirectly detected by means of heat flux measurements. The influence of the wall to stagnation temperature ratio is investigated by cooling the model with liquid nitrogen.     

Effect of flow regime change from subsonic to transonic on the air loads of an oscillating airfoil

In this research, the effect of flow regime change from subsonic to transonic on the air loads of a pitching NACA0012 airfoil is investigated. To do this, the effect of change in flow regime on the lift and pitching moment coefficients hysteresis cycles is studied. The harmonic balance approach is utilized for numerical calculation due to its low computational time. Verifications are also made with previous works and good agreements are observed. The assessment of flow regime change on the aforementioned hysteresis cycles is accomplished in the Mach number range of M=0.65–0.755. The reduced frequency and pitch amplitude also vary from k=0.03 to 0.1 and α₀=1–2.51°, respectively. Results show that the effect of increase in Mach number is to increase and decrease the lift coefficient during downstroke and upstroke, respectively, whereas at low reduced frequencies, the effect of increase in Mach number may lead to a reverse manner when airfoil moves toward its extremum angle of attack. Results also reveal that as the pitch amplitude varies, the shape of lift coefficient hysteresis cycle depends more on the pitch amplitude than on the appearance of shock. It is shown that as the Mach number increases, the incidence angles correspond to the extremum pitching moment, and depending on the reduced frequency, lie between zero and extremum angle of attack. These incidence angles shift toward the extremum angle of attack as the reduced frequency decreases. Results also show that the increase in pitch amplitude at low Mach number, in such a way that leads to the formation of shock around the extremum angle of attack, causes the extremum pitching moment to appear around these angles and at high Mach number, depending on the reduced frequency, the extremum pitching moment incidence angles would be between zero and extremum incidence angle.     

Aerodynamic performance of an annular cascade of film cooled nozzle guide vanes under engine representative conditions

This paper addresses two important issues relevant to efficiency measurements in film-cooled annular cascades: the definition of the ideal flow to be used in loss calculation, and the measurements that are necessary for such loss calculation. The paper also addresses the question of the correct parameterisation of coolant density effects, showing that the momentum flux ratio, rather than the blowing rate, is the most appropriate parameter. Experiments examining the effect of extensive aerofoil surface film cooling on the aerodynamic efficiency of an annular cascade of transonic nozzle guide vanes are reported. A dense foreign gas (SF₆/Ar mixture) is used to simulate engine representative coolant-to-mainstream density ratios, momentum ratios and blowing rates under ambient temperature conditions. Experiments are also conducted with air coolant to study the effect of density ratio on efficiency. The flowfield measurements have been obtained using a four-hole pyramid probe in a short duration blowdown facility which correctly models engine Reynolds and Mach numbers. This work compares the measured aerodynamic efficiencies of uncooled vanes with those which employ an extensive amount of cooling. The engine-representative cooling geometry investigated features 14 rows of cylindrical cooling holes, and a second geometry where 8 of these rows are replaced by holes having a fan-shaped exit. The effects of adding trailing edge slot ejection are also presented. By selectively blocking rows of holes, the cumulative effect on the mid-span efficiency of adding rows of cooling holes has also been determined. Experimental results are presented as area traverse maps (total pressure, isentropic Mach number and flow angles), from which the relative changes in efficiency due to film cooling have been calculated. These calculations reveal that the presence of foreign-gas coolant from the cylindrical-hole geometry increases the aerodynamic loss (relative to the uncooled baseline) by 6.7%; and coolant from the fan-shaped holes increases the loss by 15%. The effects of different assumptions for the coolant total pressure are shown to significantly change the measured loss; if the loss measurements are based on the mainstream total pressure, rather than the total pressure in the coolant cavity, the respective increase in loss (relative to the uncooled baseline) of cylindrical and fan-shaped holes is 4.5% and 12.5%. Experimental data is compared with loss predictions using a Hartsel model. Received: 4 December 1998/Accepted: 1 September 1999     

Separation control using synthetic vortex generator jets in axial compressor cascade

An experimental investigation conducted in a high-speed plane cascade wind tunnel demonstrates that unsteady flow control by using synthetic (zero mass flux) vortex generator jets can effectively improve the aerodynamic performances and reduce (or eliminate) flow separation in axial compressor cascade. The Mach number of the incoming flow is up to 0.7 and most tested cases are at Ma = 0.3. The incidence is 10° at which the boundary layer is separated from 70% of the chord length. The roles of excitation frequency, amplitude, location and pitch angle are investigated. Preliminary results show that the excitation amplitude plays a very important role, the optimal excitation location is just upstream of the separation point, and the optimal pitch angle is 35°. The maximum relative reduction of loss coefficient is 22.8%.The project supported by the National Natural Science Foundation of China (10477002 and 50476003) and the Ph.D. Innovative Foundation of Beihang University. The English text was polished by Yunming Chen.     

Three-dimensional effects on regular reflection in steady supersonic flows

The reflection of shock waves between two symmetrical wedges is investigated for the case of three-dimensional flows. Oblique shadowgraphs at various optical angles of yaw and pitch were used to examine the nature of fully three-dimensional flows, with wedge aspect ratios as low as 0.25 being considered. These images were used to construct surface models of the overall flow field for various reflection patterns and aspect ratios, which provides a visual indication of the flow field shape. For a Mach number of 3.1, and suitable wedge angles, the flow field with regular reflection on the tunnel centreline and Mach reflection further out is examined. The point of transition from regular reflection to the peripheral Mach surfaces is identified for various wedge angles and aspect ratios. It is shown that the transition points move outwards from the central plane as the aspect ratio decreases. This shows that three-dimensional flows favor regular reflection, because of the increasing curvature of the incident shock as the wedge becomes narrower, causing a decrease in the local angle of incidence. The height of the Mach stem is shown to be highly dependent on the geometry of the test wedge models. The Mach stem height decreases with aspect ratio due to the three-dimensional relieving effect, where the increase in lateral flow relieves the pressure over the surfaces of the wedges. Experimental evidence of the existence of the strong oblique shock solution in steady flows is presented.Received: 7 July 2003, Revised: 20 October 2003, Accepted: 6 November 2003, Published online: 10 February 2004PACS: 47.40.Nm Correspondence to: B.W. Skews     

An experimental investigation of the ELAC 1 configuration at supersonic speeds

Supersonic flight of aerospace planes is of marked interest since several flow regimes characterized by different local flow structures have to be flown through. This problem was investigated experimentally for the hypersonic research configuration ELAC 1. The aim of the study was to detect the influence of the rounded leading edge, of the thickness distribution prescribed, and of the Reynolds number, especially on the flow on the leeward side of the configuration. The experiments were carried out in the transonic wind tunnel of Aerodynamisches Institut of RWTH Aachen, at a freestream Mach number Ma _∞=2, a unit Reynolds number of Re _∞=13×10⁶, angles of attack between ?3°?α?10°, and in a wind tunnel of the Institute for Theoretical and Applied Mechanics of the Russian Academy of Sciences in Novosibirsk. The freestream Mach numbers covered in these experiments were varied between 2?Ma _∞?4, freestream Reynolds numbers per unit length between 25×10⁶?Re _∞?56×10⁶ and angles of attack between ?3°?α?10°. Flow visualization studies, measurements of surface pressure distributions and of aerodynamic forces were used to analyze the flow. The results, which will also be compared with numerical data, clearly indicate marked differences in the location of the separation and reattachment lines, and the formation of the primary, secondary and tertiary vortices, for the flow regimes investigated.     

Heat transfer coefficient measurements on the pressure surface of a transonic airfoil

This paper presents steady-state recovery temperature and heat transfer coefficient measurements on the pressure surface of a modern, highly cambered transonic airfoil. These measurements were collected with a peak Mach number of 1.5 and a maximum turbulence intensity of 30%. We used a single passage model to simulate the idealized two-dimensional flow path between rotor blades in a modern transonic turbine. This set up offered a simpler construction than a linear cascade, yet produced an equivalent flow condition. We performed validated high accuracy (±0.2°C) surface temperature measurements using wide-band thermochromic liquid crystals allowing separate measurements of the previously listed parameters with the same heat transfer surface. We achieved maximum heat transfer coefficient uncertainties that were equivalent to similar investigations (±10%). Two key observations are the heat transfer coefficient along the aft portion of the airfoil is sensitive to the surface heat flux and is highly insensitive to the level of freestream turbulence. Possible explanations for these observations are discussed.     

Enhanced performance of fast-response 3-hole wedge probes for transonic flows in axial turbomachinery

The paper presents the development and application of a three-sensor wedge probe to measure unsteady aerodynamics in a transonic turbine. CFD has been used to perform a detailed uncertainty analysis related to probe-induced perturbations, in particular the separation zones appearing on the wedge apex. The effects of the Reynolds and Mach numbers are studied using both experimental data together with CFD simulations. The angular range of the probe and linearity of the calibration maps are enhanced with a novel zonal calibration technique, used for the first time in compressible flows. The data reduction methodology is explained and demonstrated with measurements performed in a single-stage high-pressure turbine mounted in the compression tube facility of the von Karman Institute. The turbine was operated at subsonic and transonic pressure ratios (2.4 and 5.1) for a Reynolds number of 10⁶, representative of modern engine conditions. Complete maps of the unsteady flow angle and rotor outlet Mach number are documented. These data allow the study of secondary flows and rotor trailing edge shocks.     

Direct numerical simulation of three-dimensional flow past a yawed circular cylinder of infinite length

Direct numerical simulation of flow past a stationary circular cylinder at yaw angles (α) in the range of 0–60° was conducted at Reynolds number of 1000. The three-dimensional (3-D) Navier–Stokes equations were solved using the Petrov–Galerkin finite element method. The transition of the flow from 2-D to 3-D was studied. The phenomena that were observed in flow visualization, such as the streamwise vortices, the vortex dislocation and the instability of the shear layer, were reproduced numerically. The effects of the yaw angle on wake structures, vortex shedding frequency and hydrodynamic forces of the cylinder were investigated. It was found that the Strouhal number at different yaw angles (α) follows the independence principle. The mean drag coefficient agrees well with the independence principle. It slightly increases with the increase of α and reaches a maximum value at α=60°, which is about 10% larger than that when α=0°. The root-mean-square (r.m.s.) values of the lift coefficient are noticeably dependent on α.     

Investigation of turbulent flow past a yawed wavy cylinder

A large eddy simulation (LES) study was conducted to investigate the three-dimensional characteristics of the turbulent flow past wavy cylinders with yaw angles from 0° to 60° at a subcritical Reynolds number of 3900. The relationships between force coefficients and vortex shedding frequency with yaw angles for both wavy cylinders and circular cylinders were investigated. Experimental measurements were also performed for the validation of the present LES results. Comparing with corresponding yawed circular cylinders at similar Reynolds number, significant differences in wake vortex patterns between wavy cylinder and circular cylinder were observed at small yaw angles. The difference in wake pattern becomes insignificant at large yaw angles. The mean drag coefficient and the Strouhal number obey the independence principle for circular cylinders at yaw angle less than 45°, while the independence principle was found to be unsuitable for yawed wavy cylinders. In general, the mean drag coefficients and the fluctuating lift coefficients of a yawed wavy cylinder are less than those of a corresponding yawed circular cylinder at the same flow condition. However, with the increase of the yaw angle, the advantageous effect of wavy cylinder on force and vibration control becomes insignificant.     

Numerical analysis of the flow pattern and vortex breakdown over a pitching delta wing at supersonic speeds

A supersonic compressible flow over a 60° swept delta wing with a sharp leading edge undergoing pitching oscillations is computationally studied. Numerical simulations are performed by the finite volume method with the use of the k?ω turbulence model for various Mach numbers and angles of attack. Variations of flow patterns in a crossflow plane, hysteresis loops associated with the vortex core location, and vortex breakdown positions during a pitching cycle are investigated. Trends for various Mach numbers, mean angles of attack, pitching amplitudes, and pitching frequencies are illustrated.     

Improved-Delayed-Detached-Eddy Simulation of cavity-induced transition in hypersonic boundary layer

Hypersonic flow transition from laminar to turbulent due to the surface irregularities, like local cavities, can greatly affect the surface heating and skin friction. In this work, the hypersonic flows over a three-dimensional rectangular cavity with length-to-width-to-depth ratio, L:W:D, of 19.9:3.57:1 at two angles of attack (AoA) were numerically studied with Improved-Delayed-Detached-Eddy Simulation (IDDES) method to highlight the mechanism of transition triggered by the cavity. The present approach was firstly applied to the transonic flow over M219 rectangular cavity. The results, including the fluctuating pressure and frequency, agreed with experiment well. In the hypersonic case at Mach number about 9.6 the cavity is seen as “open” at AoA of −10° but “closed” at AoA of −15° unconventional to the two-dimensional cavity case where the flow always exhibits closed cavity feature when the length-to-depth ratio L/D is larger than 14. For the open cavity flow, the shear layer is basically steady and the flow maintains laminar. For the closed cavity case, the external flow goes into the cavity and impinges on the bottom floor. High intensity streamwise vortices, impingement shock and exit shock are observed causing breakdown of these vortices triggering rapid flow transition.     

Numerical Modeling of a Viscous-Gas Flow in a Turbine Cascade with Air Ejection

The gas flow pattern, the local friction coefficients, the profile losses, and the outlet flow angles in a plane turbine cascade are modeled numerically using the 2-D Reynolds equations. For describing the turbulence characteristics, a two-parameter q- turbulence model is used. The flow pattern behind the cascade trailing edge is studied. The calculated values of the local gas pressure and the friction coefficient on the profile contour, the profile losses, and the outlet flow angles are compared with the experimental data for a transonic flow in a nozzle cascade at various reduced gas velocities behind the cascade and relative mass flows of the air ejected.     

横摆角下汽车尾部湍流特征量的PIV试验分析

汽车尾部湍流场是汽车压差阻力的主要来源,在HD-2汽车模型风洞中,首先使用测力天平和测压系统,对横摆角工况下汽车模型的气动六分力和纵对称截面48个测点的表面压力进行了测量,然后利用PIV测量技术对模型在横摆角分别为0°、15°的尾部湍流场进行了测量,获得该模型尾流场的速度场、涡量场和雷诺应力流场信息,通过计算得出尾流场区域空间相关系数和湍流积分尺度。结果表明:在横摆角工况下,汽车模型尾部涡流的结构呈现向上发展的趋势;尾流场拖拽涡的范围和强度的增大导致了模型气动力出现较大的增加;湍流积分尺度的变化表明,尾部涡流区的分离噪声与涡流分离位置有关,在汽车尾部造型设计中,要尽量推迟尾部涡流的分离。     

Experimental study on aerodynamic coefficients of yawed cylinders

This paper presents wind tunnel tests on a stationary cylinder inclined with the flow. The cylinder was positioned at different sets of yaw and vertical angles. The flow regime of the tests remained in the subcritical state. Two load cells were designed and installed to measure the aerodynamic forces, with enough sensitivity to measure vortex shedding frequencies. In this paper, the three aerodynamic force coefficients are normalized using the free stream velocity instead of its normal component. The results show that the drag coefficient and the resultant of the lift and side forces coefficients can be described by an empirical function of the incidence angle. The lift and side force coefficients remain however functions of both the horizontal yaw and vertical angles and cannot be expressed as functions of the incidence angle only. The Independence Principle was observed to become inaccurate for yaw angles larger than 40°. However, the measured Strouhal numbers indicate that the vortex shedding frequencies of a yawed cylinder can be predicted using the Independence Principle.     

Drag force on a free surface-piercing yawed circular cylinder in steady flow

The force distribution on a surface-piercing yawed cylinder surface differs significantly from that on a surface-piercing vertical cylinder. The established numerical model for flow past the surface-piercing yawed cylinder with yaw angles from −45° to 45° was solved by the standard large-eddy simulation (LES) methodology. Six cases at intervals of ±15° relative to the vertical were studied at the Reynolds number of 27 000 and the Froude number of 0.8 based on the cylinder diameter and free-stream velocity, among which the drag forces on four cylinders with yaw angles from −15° to 30° were tested for the validation of the LES approach. The results revealed that the time-averaged total drag coefficient for all cases increases with the increase of yaw angle compared to that of the surface-piercing vertical cylinder, even over 2.5 for the ±45°-yawed cylinders. The sectional drag coefficients for the negatively yawed cylinders are much greater than that for the vertical cylinder, and much less for the positively yawed cylinders. The unbalanced hydrostatic pressures on the inclined section are mainly responsible for those increment and decrement. Once the hydrostatic pressure was removed, the sectional drag coefficient on the mid-span of the positively yawed cylinder increases from the top section to the bottom, and decreases for the negatively yawed cylinder. The corresponding integrated total drag coefficient decreases with the increase of the yaw angle to ±15°, then increases with the further increase of the magnitude of yaw angle.     

Shock tunnel study of spiked aerodynamic bodies flying at hypersonic Mach numbers

A three-component accelerometer balance system is used to study the drag reduction effect of an aerodisc on large angle blunt cones flying at hypersonic Mach numbers. Measurements in a hypersonic shock tunnel at a freestream Mach number of 5.75 indicate more than 50% reduction in the drag coefficient for a 120° apex angle blunt cone with a forward facing aerospike having a flat faced aerodisc at moderate angles of attack. Enhancement of drag has been observed for higher angles of attack due to the impingement of the flow separation shock on the windward side of the cone. The flowfields around the large angle blunt cone with aerospike assembly flying at hypersonic Mach numbers are also simulated numerically using a commercial CFD code. The pressure and density levels on the model surface, which is under the aerodynamic shadow of the flat disc tipped spike, are found very low and a drag reduction of 64.34% has been deduced numerically.     

Experimental study of compressible boundary layer on a cone at angles of attack

Experimental study was conducted for boundarylayers on a sharp 5° half-angle cone of 400mm length at angles of attack. The model was tested in the T-326 hypersonic wind tunnel （ITAM） at freestream Mach number M = 5.95. Mean and fluctuation wall characteristics of the boundary layer are measured at 0°, 2°, 3° and 4° angles of attack for different stagnation pressures. Pulsation measurements are carried out by means of ALTP sensor. Pressure and temperature distributions along the model are obtained, and transition beginning and end locations have been found. Boundary layer stabilization with the increase of angle of attack and the decrease of stagnation pressure is observed. High frequency pulsations inherent to hypersonic boundary layer （second mode） have been detected.