High resolution PIV measurements around a model turbine blade trailing edge film-cooling breakout

Film cooling downstream of a model turbine blade trailing edge has been studied experimentally. High resolution particle image velocimetry was used to obtain spatially resolved mean velocity and turbulence measurements in the immediate vicinity of the trailing edge breakout. The mean velocity measurements imply the presence of a pair of counter-rotating longitudinal vortices shed from the sides of the breakout lands. The turbulent shear stress measurements above the breakout are significantly intensified as blowing ratio is increased. These results suggest that there is a strong mixing between the film cooling slot jets and the mainstream flow which degrades the film cooling effectiveness.     

An inviscid model for the base pressure of transonic turbine cascade

An inviscid base pressure model for transonic turbine blade has been presented. It has been shown that for a given back pressure the base pressure at the trailing edge, and the profile loss of a turbine blade are fixed according to the model and the base pressure can be calculated with the help of an inviscid numerical scheme. A parameteric study on the model shows that a blade profile with positive curvature downstream of the throat is advantageous for generating less loss, whilst the worst situation is when the exit flow reaches the sonic condition.     

DIRECT NUMERICAL SIMULATION OF TURBINE CASCADE FLOW WITH HEAT TRANSFER EFFECTS

低雷诺数流动对高空动力装置, 特别是涡轮部件的性能产生重要的影响. 本文采用具有7阶精度的差分格式, 通过直接求解二维瞬态可压缩Navier-Stokes方程组, 对雷诺数为241 800 (基于叶片弦长)时的叶片表面带有热传导效应的平面涡轮叶栅流动进行了二维直接数值模拟, 对低雷诺数平面涡轮叶栅流动的非定常流动现象作了初步的探索.数值结果表明:在叶栅通道入口处, 流场的非定常性很弱;在叶栅尾缘处, 具有正负涡量的尾涡交替地从压力面和吸力面上脱落;周期性的涡脱落使得叶栅通道内和尾迹区的总压发生(准)周期的变化, 并且, 尾迹区总压变化主频率是通道内总压变化主频率的2倍;在时均流场中, 叶片表面压力的分布与实验值吻合良好, 表征热传导效应的斯坦顿数除湍流区外与实验值基本吻合;尾迹区速度脉动的2阶统计量与圆柱绕流尾迹区速度脉动2阶统计量具有基本相似的分布特征.     

Loss Reduction by Means of Two-Dimensional Roughness Elements on the Suction Surface of a Linear Turbine Rotor Cascade

This paper reports the results of experimental investigations carried out to reduce pressure losses by means of two-dimensional roughness elements (in the form of stainless steel tubes of different diameters). The roughness elements are fixed at various axial stations on the suction surface of 120° turning, 175 mm chord, impulse turbine rotor blades. Flow measurements are carried out at the exit of the cascade at five axial stations, using a five-hole probe operating in non-nulling mode. In addition, the blade surface static pressure distribution is measured. The data from the five-hole probe measurements are used to calculate pressure, velocity and flow angle distributions at the cascade exit and these results are used to calculate mass averaged values and integral parameters such as wake half-width, loss coefficient, etc. The static pressure distribution is altered very little except near the roughness element. The lift coefficient remains almost constant for all configurations and the drag coefficient is reduced for some configurations. The non-dimensional total pressure defects in the wake for all configurations followed Gaussian distribution. A two-dimensional roughness element of 0.6 mm diameter placed at 0.65 chord on the suction surface showed an appreciable reduction in pressure losses.     

Laser velocimetry measurements in a laboratory vaned diffuser

Laser velocimetry measurements were made within a laboratory radial vaned diffuser with three different blade configurations. Measurements were made through passages with four, six and eight blades installed at off design conditions. Also, in the eight blade diffuser measurements were made between the blade passage exit and diffuser exit so that the complete secondary flow could be defined. The flow was found to separate from the blades and form large separation zones. The separation zones consisted primarily of two vortices rotating in opposite directions. At the passage exit the separation region encompassed 23% of the circumferential area for the four blade diffuser, 45% for the six blade and 40% in the eight blade diffuser. Separation occurred at 23%, 27% and 50% from the leading edge of the blades for the 4, 6 and 8 bladed diffusers, indicating that more blades better controlled the separation. Turbulence intensities ranged from approximately 5% to 15% in the primary flow and reached a few hundred percent in the secondary flow within the separation regions.     

Seven hole probe measurement of leading edge vortex flows

This paper discusses the use of a seven-hole probe on measurements of leading edge vortices of highly sweep delta wing planforms. Intrusive probe data taken with the pressure probe were compared with non-intrusive measurements made with laser Doppler anemometry system. In addition to probe size, the natural position of breakdown and the sweep angle of the wing are also factors in determining sensitivity of the flow to probe interference. At low angles of attack vortex breakdown does not occur in the vincinity of the model and the seven hole probe was found to yield reasonably accurate measurements. When the angle of attack of the model was increased so that vortex breakdown was near the trailing edge, introducing the probe over the wing would cause the breakdown position to move ahead of the probe. However, when breakdown naturally occurred ahead of the mid-chord of the wing the vortices were found to be less sensitive to a probe placed behind the breakdown point. Vortex breakdown on a lower swept wing is found to be more sensitive to interference. Near the breakdown region, seven hole probe measurement is less accurate due to a combination of probe interference and flow reversal.     

Experimental investigation of the interaction of two flows on the axial fan hollow blades by flow visualization and hot-wire anemometry

Experimental investigation of the interaction of internal flow with external flow around hollow airfoil NACA series in a low-speed wind tunnel was conducted and is presented in the paper. The region near the trailing edge of the hollow airfoil was studied in detail and measurements of velocity and turbulence intensities were performed with hot-wire anemometry. Determination of flow structure on the hollow airfoil was performed with computer-aided visualization. It can be concluded from the measurement analysis that higher values of velocities, lower turbulence intensities and a significant decrease of circulation effects on the suction side of the hollow blade were achieved, due to the introduction of internal flow. The results obtained on the hollow airfoil were applied on the rotating axial fan. Influence of the internal flow of the hollow blade on the flow field of the axial fan was studied. With the introduction of the internal flow a reduction of circulation effects on the fan hollow blade was achieved. Aerodynamic characteristic of the axial fan reached higher degree of total pressure difference and normalized efficiency through the entire fan working conditions.     

Flow field analysis inside a gas turbine trailing edge cooling channel under static and rotating conditions

The flow field inside a modern internal cooling channel specifically designed for the trailing edge of gas turbine blades has been experimentally investigated under static and rotating conditions. The passage is characterized by a trapezoidal cross-section of high aspect-ratio and coolant discharge at the blade tip and along the wedge-shaped trailing edge, where seven elongated pedestals are also installed. The tests were performed under engine similar conditions with respect to both Reynolds (Re = 20,000) and Rotation (Ro = 0, 0.23) numbers, while particular care was put in the implementation of proper pressure conditions at the channel exits to allow the comparison between data under static and rotating conditions. The flow velocity was measured by means of 2D and Stereo-PIV techniques applied in the absolute frame of reference. The relative velocity fields were obtained through a pre-processing procedure of the PIV images developed on purpose.Time averaged flow fields inside the stationary and rotating channels are analyzed and compared.A substantial modification of the whole flow behavior due to rotational effects is commented, nevertheless no trace of rotation induced secondary Coriolis vortices has been found because of the progressive flow discharge along the trailing edge. For Ro = 0.23, at the channel inlet the high aspect-ratio of the cross section enhances inviscid flow effects which determine a mass flow redistribution towards the leading edge side. At the trailing edge exits, the distortion of the flow path observed in the channel central portion causes a strong reduction in the dimensions of the 3D separation structures that surround the pedestals.     

Tip gap height effects on the aerodynamic performance of a cavity squealer tip in a turbine cascade in comparison with plane tip results: part 1—tip gap flow structure

Tip gap height effects on the flow structure over a cavity squealer tip have been investigated in a linear turbine cascade for power generation, in comparison with the corresponding plane tip results. Oil film flow visualizations are conducted on the tip surface and casing wall for tip gap height-to-chord ratios of h/c = 1.0, 2.0, and 3.0%. The squealer tip has a recessed cavity enclosed by a full length squealer with its rim height-to-chord ratio of 5.51%. The results show that most of in-coming fluid entering the tip gap inlet for the cavity squealer tip is entrapped by the suction-side squealer rim, and the cavity fluid is discharged into the blade flow passage over the suction-side squealer rim in the region from the mid-chord to the trailing edge. Regardless of h/c, the cavity squealer tip makes the leakage flow zone narrower than the plane tip, and is superior to the plane tip in reducing the tip leakage mass flow rate. A qualitative flow model describing full flow features over the cavity squealer tip is suggested. In this flow model, the tip gap exit area is classified into four different regions, and the tip gap height effects on the discharge characteristics in each region are discussed in detail.     

Turbulence measurements in a transonic two-passage turbine cascade

This paper presents detailed turbulence measurements in a two-dimensional, transonic, double passage turbine cascade. Particle image velocimetry was used to obtain mean velocity and turbulence measurements all around a single turbine blade within about 2 mm of the blade and wall surfaces. The passage walls were designed using an optimization procedure so that the blade surface pressure distribution matches that of the blade in an infinite cascade. The resulting experimental model captures much of the complexity of a real turbine stage (including high streamline curvature, strong accelerations, and shocks) in a passage with a continuous wall shape, allowing for high measurement resolution and well controlled boundary conditions for comparison to CFD. The measurements show that in the inviscid regions of the passage the absolute level of the turbulent fluctuations does not change significantly as the flow accelerates, while the local turbulence intensity drops rapidly as the flow accelerates. These results provide a benchmark data set that can be used to improve turbulence models.     

Mean and turbulence measurements in the boundary layer and wake of a symmetric aerofoil

Detailed measurements of two-dimensional profiles of static pressure, mean velocity, turbulence intensity and Reynolds shear stress were carried out with conventional pressure probes and hot wire probes at preselected streamwise stations in the boundary layer and wake of a 12.5% thick, 600 mm chord two-dimensional symmetric aerofoil mounted at zero incidence in a low speed wind tunnel. The chord Reynolds number was one million and the wake measurements extended up to three chord lengths (or nearly 660 trailing edge momentum thicknesses) downstream of the trailing edge. The data indicate rapid interaction of the wall layers immediately behind the trailing edge, leading to significant changes in the flow parameters close to the trailing edge. The relaxation of the wake is preceded by initial ‘overshoot’ in the streamwise profiles of mean-flow parameters and peak values of turbulence components. Further growth of the wake towards similarity/equilibrium is discussed.     

Effects of camberwise varying tip injection on loss and wake characteristics of a low pressure turbine blade

This paper presents the results of an experimental study that investigates the effects of camberwise varying tip injection on the total pressure loss and wake flow characteristics downstream of a row of Low Pressure Turbine (LPT) blades. This injection technique involves spanwise jets at the tip that are issued from a series of holes distributed along the camber line. The injection from each hole is individually and separately controlled using a computer driven solenoid valve and therefore the flow injection geometrical pattern at the tip can be adjusted to any desired variation. Three different injection cases are investigated including triangular, reversed triangular and uniform injection patterns. Here, triangular and reversed triangular cases refer to discrete blowing from the blade tip in which the blowing velocity increases (triangular) or decreases (reverse triangular) linearly from the leading to trailing edge along the camber. For uniform injection, the injection velocity is kept constant for all injection holes. The total mass injection from the tip is kept the same for all injection cases. The experiments are conducted in a continuous-flow wind tunnel with a linear cascade test section and measurements involve Kiel probe traverses 0.5 axial chords downstream of the blades covering a region between 85% and 100% span as well as two-dimensional Particle Image Velocimetry (PIV) measurements on 50%, 85% and 95% spanwise planes. For all injection cases, results show that tip injection reduces the total pressure loss levels in general. Highest measured overall loss reduction occurs in the case of reversed-triangular injection. The least effective waveform is observed to be triangular injection. There is significant reduction in the extent of the low momentum zone of the leakage vortex with injection. This effect is much less pronounced for the passage vortex. On the other hand, complex flow patterns are observed within the passage vortex, especially in the case of reversed-triangular injection, such as a possible embedded vortical structure along the passage vortex core, which creates double peaks in the velocity and turbulent kinetic energy fields.     

Investigation of aerodynamic effects of coolant ejection at the trailing edge of a turbine blade model by PIV and pressure measurements

In order to simulate the thick trailing edges of turbine blades a slotted plate profile together with a newly designed nozzle was installed into the high-speed wind tunnel of the DLR Göttingen. At different supersonic Mach numbers and at four coolant flow rates in the range of 0–2.5% pressure distribution measurements and probe measurements were performed. The flow field was visualized by schlieren photos and the instantaneous velocity field was quantitatively investigated by Particle Image Velocimetry (PIV). The measurements of the velocity field gave an insight into stationary effects, for example the change of shock strength with coolant flow rate, and instationary effects such as the existence of a vortex street in the wake. The PIV technique offers special advantages for the investigation of transonic flow fields, but also yields to special experimental difficulties, which are also described in this article. Measured losses display a maximum at the downstream Mach number 1. This is strongly related to the behaviour of the base pressure. A loss minimum is achieved at moderate coolant flow rates, showing that an optimum coolant flow rate exists. The loss was analysed and separated into the loss contributions from the profile upstream of the trailing edge and the mixing loss due to the coolant flow.     

Isothermal velocity and turbulence measurements downstream of a model multilobed turbofan mixer

An isothermal experimental investigation of the three dimensional flow field downstream of a model multilobed turbofan forced mixer is presented. LDA measurements of the three mean velocities and corresponding turbulence intensities were obtained in the downstream duct where the turbine (primary) and fan (secondary) streams emerging from the lobes mix together. The flow development in the near field was quantified by measuring the cross plane velocities. These were found to consist of large radial flows, of order 15% of the mean axial velocities at the lobe inlet, with changing sign depending on location. The cross-plane flow is consistent with a large scale axial vortex pair (per lobe) which persists throughout the downstream duct and enhances mixing in this region. Turbulence generation and anisotropy of the turbulence structure were evident mainly in the shear layers formed as the fan and turbine streams emanated from the lobe trailing edge. Spatial uniformity in the mean and turbulent fields was measured as little as five heights downstream of the mixer exit, indicating the rapid mixing achievable in these systems.     

Particle image velocimetry （PIV） measurements of tip vortex wake structure of wind turbine

Large-view flow field measurements using the particle image velocimetry (PIV) technique with high resolution CCD cameras on a rotating 1/8 scale blade model of the NREL UAE phase VI wind turbine are conducted in the engineering-oriented Φ3.2 m wind tunnel.The motivation is to establish the database of the initiation and development of the tip vortex to study the flow structure and mechanism of the wind turbine.The results show that the tip vortex first moves inward for a very short period and then moves out...     

DNS of secondary flows over oscillating low-pressure turbine using spectral/hp element method

This paper investigates the secondary vortex flows over an oscillating low-pressure turbine blade using a direct numerical simulation (DNS) method. The unsteady flow governing equations over the oscillating blade are discretized and solved using a spectral/hp element method. The method employs high-degree piecewise polynomial basis functions which results in a very high-order finite element approach. The results show that the blade oscillation can significantly influence the transitional flow structure and the wake profile. It was observed that the separation point over vibrating T106A blades was delayed 4.71% compared to the stationary one at Re = 51,800. Moreover, in the oscillating case, the separated shear layers roll up, break down and shed from the trailing edge. However, the blade vibration imposes additional flow disturbances on the suction surface of the blade before leaving from the trailing edge. Momentum thickness calculations revealed that after flow separation point, the momentum thickness grows rapidly which is due to the inverse flow gradients which generate vortex flows in this area. It was concluded that the additional vortex generations due to the blade vibrations cause higher momentum thickness increment compared to the conventional stationary LPT blade.     

Instantaneous and time-averaged flow fields of multiple vortices in the tip region of a ducted propulsor

The instantaneous and time-averaged flow fields in the tip region of a ducted marine propulsor are examined. In this flow, a primary tip-leakage vortex interacts with a secondary, co-rotating trailing edge vortex and other co- and counter-rotating vorticity found in the blade wake. Planar particle imaging velocimetry (PIV) is used to examine the flow in a plane approximately perpendicular to the mean axis of the primary vortex. An identification procedure is used to characterize multiple regions of compact vorticity in the flow fields as series of Gaussian vortices. Significant differences are found between the vortex properties from the time-averaged flow fields and the average vortex properties identified in the instantaneous flow fields. Variability in the vortical flow field results from spatial wandering of the vortices, correlated fluctuations of the vortex strength and core size, and both correlated and uncorrelated fluctuations in the relative positions of the vortices. This variability leads to pseudo-turbulent velocity fluctuations. Corrections for some of this variability are performed on the instantaneous flow fields. The resulting processed flow fields reveal a significant increase in flow variability in a region relatively far downstream of the blade trailing edge, a phenomenon that is masked through the process of simple averaging. This increased flow variability is also accompanied by the inception of discrete vortex cavitation bubbles, which is an unexpected result, since the mean flow pressures in the region of inception are much higher than the vapor pressure of the liquid. This suggests that unresolved fine-scale vortex interactions and stretching may be occurring in the region of increased flow variability.     

Tip leakage aerodynamics over stepped squealer tips in a turbine cascade

Tip gap flow physics and aerodynamic loss generations for two stepped squealer tips of a “Higher Pressure-side rim and Lower Suction-side rim” (HPLS) tip and a “Lower Pressure-side rim and Higher Suction-side rim” (LPHS) tip have been investigated in a turbine cascade. For a fixed tip gap height-to-chord ratio of h/c = 2.0%, oil film flow visualizations are performed on the casing wall as well as on the cavity floor, and three-dimensional flow fields downstream of the cascade are measured with a five-hole probe. For the HPLS tip, the leakage inflow over the pressure-side rim cannot reach the suction-side rim in the upstream region due to the presence of an inlet flow intrusion, and there exists a strong near-wall flow heading toward the trailing edge all over the cavity floor. On the other hand, the LPHS tip has a mid-chord leakage flow penetration into the blade flow passage, and also provides a downstream leakage flow penetration deeper than that for the HPLS tip. Its cavity floor can be divided into a backward flow region and a wide separation bubble. Aerodynamic loss for the HPLS tip, which is nearly identical to that for the cavity squealer tip, is lower than those for the LPHS and plane tips in a considerable degree.     

AN APPLICATION OF TOPOLOGICAL ANALYSIS TO STUDYING THE THREE-DIMENSIONAL FLOW IN CASCADES; PART I—TOPOLOGICAL RULES FOR SKIN-FRICTION LINES AND SECTION STREAMLINES

Based on the working of Lighthill and Hunt et al., in the present paper the author has established the topological rules adapting to analysing the skin-friction lines and the section streamlines in cascades. These rules are (1) for a rotor cascade without shroud band, the total number of nodal points equals that the saddle points on the skin-friction line vector fields in each pitch range; (2) for an annular or straight cascade with no clearances at blade ends, the total number of saddle points is two more than that of nodal points on the skin-friction line fields in a pitch; (3) the total number of saddles in the secondary flow fields on cross-sections in cascade is one less than that of nodes; (4) in the section streamline vector fields on a meridian surface penetrating a flow passage, and on leading and trailing edge sections, the total number of nodes is equal to that of saddles; (5) on the streamline vector fields of a blade-to-blade surface, the total number of nodes is one less than that of saddles.     

Calibration of a four-hole pyramid probe and area traverse measurements in a short-duration transonic turbine cascade tunnel

A four-hole pyramid probe has been calibrated for use in a short-duration transonic turbine cascade tunnel. The probe is used to create area traverse maps of total and static pressure, and pitch and yaw angles of the flow downstream of a transonic annular cascade. This data is unusual in that it was acquired in a short-duration (5 s of run time) annular cascade blowdown tunnel. A four-hole pyramid probe was used which has a 2.5 mm section head, and has the side faces inclined at 60° to the flow to improve transonic performance. The probe was calibrated in an ejector driven, perforated wall transonic tunnel over the Mach number range 0.5–1.2, with pitch angles from -20° to + 20° and yaw angles from-23° to +23°. A computer driven automatic traversing mechanism and data collection system was used to acquire a large probe calibration matrix (～ 10,000 readings) of non dimensional pitch, yaw, Mach number, and total pressure calibration coefficients. A novel method was used to transform the probe calibration matrix of the raw coefficients into a probe application matrix of the physical flow variables (pitch, yaw, Mach number etc.). The probe application matrix is then used as a fast look-up table to process probe results. With negligible loss of accuracy, this method is faster by two orders of magnitude than the alternative of global interpolation on the raw probe calibration matrix. The blowdown tunnel (mean nozzle guide vane blade ring diameter 1.1 m) creates engine representative Reynolds numbers, transonic Mach numbers and high levels (≈ 13%) of inlet turbulence intensity. Contours of experimental measurements at three different engine relevant conditions and two axial positions have been obtained. An analysis of the data is presented which includes a necessary correction for the finite velocity of the probe. Such a correction is non trivial for the case of fast moving probes in compressible flow.