Cascade Test Facility Research Articles

A study of angle parametric optimization of the composite honeycomb on turbine blade tip

To reduce the secondary flow loss of the high-pressure turbine caused by the tip leakage flow, the design of the composite honeycomb tip has been experimentally and numerically examined in this study. The arrangement angle of the composite honeycomb is defined and optimized using the radial basis function and genetic algorithm. The tip flow field of the high-pressure turbine blade with a 2.128 mm clearance is simulated by CFX 18.0, and the arrangement angle of the composite honeycomb is optimized to obtain a lower loss. Moreover, the flat tip, basic honeycomb tip, and optimized honeycomb tip are tested in a low-speed cascade test facility. The results show that the vortices and radial velocity induced by the honeycomb structure cavity increase the kinetic energy loss of the leakage flow and reduce the leakage flow rate. Compared with the basic honeycomb, the optimized honeycomb has a supplemental blocking effect on the tip leakage flow by reducing the crosswise pressure gradient in the blade tip. The optimized honeycomb also changes the outlet angle of the leakage flow in the tip clearance, reduces the leakage vortex, and the total pressure loss is further reduced by 2.5%.

Effect of the distribution of the upstream probe support on performance of compressor cascade

Abstract Aiming at the problem of aerodynamic coupling interference between disturbance of the built-in measurement probe and internal flow of high load compressor, the experiment of axial position and structure dimension of cylindrical probe support effect on the performance of the compressor cascade was carried out on the transonic plane cascade test facility. The effect characteristics of the distribution of probe support on aerodynamic performance of downstream cascade under different inlet Mach numbers were analyzed through the experimental data. The experimental results show that the axial position of probe support has little impact on the total pressure loss coefficient of cascade under the low-speed inlet flow condition. However, the variety of the axial position of probe support has a great impact on the total pressure loss coefficient of cascade under the high-speed inlet flow condition. With the decrease of the axial distance between probe support and cascade, the total pressure loss coefficient of cascade increases. Under different inlet flow conditions, the variety of the structure dimension of probe support has a significant impact on the total pressure loss coefficient of cascade. With the increase of the structure dimension of probe support, the total pressure loss coefficient of cascade increases.

환형 섹터 터빈 노즐 캐스케이드에서 팬 형상 막냉각 홀의 복합분사각도 적용에 따른 막냉각 효율 특성

The present study investigated the effect of compound angle injection of fan-shaped film cooling holes on film cooling performance on the nozzle vane surface in an annular sector cascade. Two different configurations were examined to investigate effect of compound angle injection. The baseline has streamwise injection holes on the pressure and suction side surfaces and normal injection holes in leading edge region. The other configuration has compound angle injection holes on the leading edge and the pressure side surface. An annular sector turbine cascade test facility was used and the measurements using an infrared thermography method were conducted at the exit Reynolds number of 2.2x10<SUP>6</SUP> and exit Mach number of 0.8. Total coolant mass flow rate ranges between 5 and 10% of mainstream. The results showed that the baseline configuration shows skewed trajectory of coolant downstream of holes on the pressure side surface, which is not observed in ‘flat-plate’ study and strongly related to mainstream flow direction near the nozzle surface. On the other hands, the coolant follows the mainstream direction and spread more uniformly on the pressure side surface with compound angle injection configuration. However as coolant mass flow rate increases, compound angle injection induces stronger interaction of mainstream and coolant and film cooling effectiveness drops quickly in the downstream region. Also, film cooling effectiveness in the suction side surface shows lower values for compound angle injection configuration since compound angle injection on the showerhead region causes non-uniform distributions of film cooling effectiveness due to stronger secondary flow.

천음속 선형 터빈 캐스케이드에서 막냉각 및 내부냉각이 적용된 2단 터빈 베인의 냉각성능 연구: 전냉각효율 측정

The present study investigated the cooling performance of the second stage turbine nozzle vane using a transonic linear turbine cascade test facility. For experiments, 500 hp or 2,250 hp compressor and electric heaters were used for mainstream supply and 50 hp compressor for secondary flow supply and a single nozzle vane with adjusted sidewalls on each side of the nozzle was installed in the test section. The test nozzle model having complex cooling configuration was made of Inconel 718 using metal 3D printing technique. The overall cooling effectiveness of the cooled turbine nozzle vane was measured using two Infrared cameras. Four different sets of experiments were conducted under two different mainstream conditions, subsonic and transonic conditions and the corresponding Reynolds numbers of the mainstream are 6.5x10SUP5/SUP and 1.3x10SUP6/SUP based on chord length, respectively. The results showed that overall cooling effectiveness on the nozzle vane surface is dominated by internal heat convection in a rib-roughened surface, film cooling flow around the leading edge region and heat conduction around the hole and pin-fins. When comparing the results at different inlet Reynolds numbers of the internal cooling passage, the internal cooling configuration has more dominant effect on overall cooling performance than film cooling. Therefore it is obvious that overall cooling effectiveness as well as film cooling effectiveness should be taken into account to assess exact cooling performance and characteristics on turbine nozzle vane cooling design.

An experimental method for squealer tip flow field considering relative casing motion

Squealer tip is widely used in turbines to reduce tip leakage loss. In typical turbine environment, the squealer tip leakage flow is affected by multiple factors such as the relative casing motion and the wide range of variable incidence angles. The development of experimental methods which can accurately model the real turbine environment and influencing factors is of great significance to study the squealer tip leakage flow mechanism. In the present paper, a low-speed turbine cascade test facility which can model the relative casing motion and wide range of variable incidence angles (−25° to 55°) is built. Based on the similarity criteria, a high-low speed similarity transformation method of the turbine cascade is established by considering the thickness of the turbine blade. A combined testing method of Particle Image Velocimetry (PIV) and local pressure measurement is proposed to obtain the complex flow structures within the tip cavity. The results show that the experimental method can successfully model the relative casing motion and the wide range of variable incidence angles. The low-speed cascade obtained by the similarity transformation can model the high-speed flow accurately. The measurement technique developed can obtain the complex flow field and successfully capture the scraping vortex within the squealer tip.

Experimental Investigation of the Effect of the Probe Support Tail Structure on the Compressor Cascade Flow Field

Abstract Aiming at resolving the problem of measuring probe blockage effect in the performance experiments of high loaded axial flow compressors, an experimental investigation of the probe support disturbance effect on the compressor cascade flow field was conducted on a transonic plane cascade test facility. The influence characteristics of the probe support tail structure on the cascade downstream flow field under different operation conditions were revealed through the detailed analysis of the test data. The results show that the aerodynamic coupling effect between the upstream probe support wake and the downstream cascade flow field is very intense. Some factors, i. e. inlet Mach number, probe support tail structure, circumferential installing position of probe, and axial distance from the probe support trailing edge to the downstream cascade, are found to have the most impact on the probe disturbance intensity. Under high speed inlet flow condition, changing probe support tail structure can’t inhibit probe support disturbance intensity effectively. Whereas under low speed inlet flow condition, compared with the cylindrical probe, the elliptic probe can inhibit probe support wake loss and reduce disturbance effects on the downstream cascade flow field.

An experimental investigation of the effects of upstream probe disturbance on compressor cascade performance

An experimental study was conducted in the present study to investigate the effects of upstream probe disturbance on the compressor cascade performance. The experiments were carried out using a plane cascade test facility where the aerodynamic coupling mechanisms between the inner probe and compressor were particularly evaluated in terms of the complex reversed pressure gradient. Moreover, the influence of probe install position near the cascade leading edge on the downstream flow characteristic was evaluated in detail. Results show that the presence of probe at cascade leading edge can reduce the total pressure recovery coefficient, raise the wake loss at cascade outlet and deteriorate the periodicity of conventional cascade flow field distribution. An optimum circumferential installing position at the upstream of the cascade is found with a minimum disturbance, which is not affected by the varying Mach numbers significantly. The probe installed at the intermediate of flow passage has a smaller disturbance on the downstream cascade performance than that installed in front of cascade leading edge.

Improvement of Turbine Vane Film Cooling Performance by Double Flow-Control Devices

This study deals with the studies of the effect of double flow-control devices (DFCDs) on turbine vane film cooling. Aiming for improving film effectiveness, two semispheroid DFCDs per pitch were attached to the vane surface upstream of the cooling hole. Although the DFCDs were successfully applied to the flat-plate film cooling in the previous study, the applicability to the turbine vane was to be investigated. In order to observe the flow field in detail, Reynolds-averaged Navier–Stokes (RANS) simulation was conducted first. The DFCDs were installed upstream of each cooling hole of the pressure and suction sides of the vane to investigate the effect of the device position. In this paper, the effects of blowing ratio and cooling hole pitch were also investigated. The results obtained by CFD showed that the vortex generated from DFCD suppressed lift-off of the secondary air. As a result, the film effectiveness became significantly higher than that without DFCD condition. Moreover, the improvement in the film effectiveness by DFCD was observed by both of the pressure and suction sides of the turbine vane. Based on the findings through RANS simulation, adiabatic effectiveness and total pressure loss coefficient measurement were performed in a linear cascade test facility. The experiment confirmed that the film effectiveness was improved when DFCDs existed.

Experimental and numerical investigation of outlet guide vane and endwall heat transfer with various inlet flow angles

This paper investigates the heat transfer on the outlet guide vane (OGV) surface and its endwall region. The Reynolds number is fixed at 300,000 and the flow is subsonic. The inlet flow angle is varied from +25° (on-design), to +40° and −25° (off-design). Experiments were conducted in a linear cascade test facility using thermochromic liquid crystal technique. Numerical simulations using RANS were carried out with three turbulence models, i.e., standard k–ω model (k–ω), baseline k–ω model (BSL), and shear stress transport k–ω model (SST). Both the experimental and numerical results are provided and compared. On the OGV surface, boundary layer transition and separation affect the heat transfer significantly and they vary with the inlet flow angle. The abilities of the three models to predict these heat transfer behaviors are revealed. For the on-design case, both BSL and SST models capture the main feature of the heat transfer variations due to transition, but the k–ω model fails. For off-design cases where separation occurs, there are discrepancies found between the calculations and experimental data. On the endwall region, the effects of a horseshoe vortex (HV) on the heat transfer are clearly noticed at the leading edge (LE). The three models perform well to simulate the pitchwise averaged Nusselt number while they always over-predict the strength and size of the HV, which leads to higher heat transfer there compared to the measurements. For off-design conditions, the HV becomes more energetic than that of the on-design condition and the pressure side leg departs from the OGV at the inlet flow angle α=−25°.

J0520204 乱流促進デバイスを用いた航空エンジン用低圧タービン翼の高負荷化・高性能化に関する研究

This paper details experimental and numerical studies on the effectiveness in controlling boundary layer separation with the performance of dimples on a low-pressure turbine airfoil. The profile loss due to boundary layer separation in LPT has a significant contribution to its overall loss, especially in a high stage loading condition. Hence it becomes important to study methods to control the boundary layer separation. A linear cascade test facility is employed to investigate the aerodynamic performance of the airfoil by use of a miniature Pitot probe. LES simulations are also carried out to gain a better understanding of the flow field around the airfoil. The Reynolds number based on the chord length was varied in the range of 40,000 to 170,000 in order to simulate a typical flight condition. It is found that the proposed methods have a capability to reduce the profile loss and result in improvement of cascade performance.

J0530202 乱流促進デバイスを用いた航空エンジン用低圧タービン翼の高負荷化に関する研究([J053-02]流体機械の研究開発におけるEFD/CFD(2),流体工学部門)

In order to improve the efficiency of high-lift low pressure turbines, this paper investigates two types of airfoils which have two-dimensional contouring with a small step or dimples on the suction surface to suppress the separation bubble that causes large aerodynamic loss, especially in the case of high stage loading. A linear cascade test facility is employed to investigate the aerodynamic performance of the two-dimensional contoured airfoil by use of miniature Pitot probes and RANS simulations are carried out to investigate the three types of dimple layouts. The Reynolds number based on the chord length was varied in the range of 40,000 to 170,000 to simulate a typical flight condition. It is found that the proposed methods are capable of reducing the profile loss, resulting in an improvement of cascade performance.

Aerothermal Performance of a Cooled Winglet at Engine Representative Mach and Reynolds Numbers

This paper presents an experimental investigation of the aerothermal performance of a cooled winglet tip under transonic conditions (exit Mach number of 1.0, and an exit Reynolds number of 1.27 × 106, based on axial chord). Spatially resolved heat transfer data and film cooling effectiveness data are obtained using the transient infrared thermography technique in the Oxford High-Speed Linear Cascade test facility. Aerodynamic loss data are obtained by traversing a specially made and calibrated three-hole pressure probe and a single-hole probe one axial chord downstream of the blade. Detailed contours of Nusselt number show that for an increase in tip clearance, with and without film cooling, and for coolant injection, for both tip clearances, the Nusselt number increases. Also the smaller tip clearance observes higher film cooling effectiveness overall. Detailed distributions of kinetic energy losses as well as pitch-wise averaged loss coefficients and loss coefficients at a mixed-out plane indicate that the size of the loss core associated with the over-tip leakage vortex decreases with cooling injection.

Aerothermal Performance of a Winglet at Engine Representative Mach and Reynolds Numbers

This paper presents an experimental and numerical investigation of the aerothermal performance of an uncooled winglet tip, under transonic conditions. Spatially resolved heat transfer data, including winglet tip surface and near-tip side-walls, are obtained using the transient infrared thermography technique within the Oxford high speed linear cascade test facility. Computational fluid dynamics (CFD) predictions are also conducted using the Rolls-Royce HYDRA suite. Most of the spatial heat transfer variations on the tip surface are well-captured by the CFD solver. The transonic flow pattern and its influence on heat transfer are analyzed, which shows that the turbine blade tip heat transfer is greatly influenced by the shock wave structure inside the tip gap. The effect of the casing relative motion is also numerically investigated. The CFD results indicate that the local heat transfer distribution on the tip is affected by the relative casing motion but the tip flow choking and shock wave structure within the tip gap still exist in the aft region of the blade.

Measurements of Secondary Losses in a Turbine Cascade With the Implementation of Nonaxisymmetric Endwall Contouring

An approach to endwall contouring has been developed with the goal of reducing secondary losses in highly loaded axial flow turbines. The present paper describes an experimental assessment of the performance of the contouring approach implemented in a low-speed linear cascade test facility. The study examines the secondary flows of a cascade composed of Pratt &amp; Whitney PAKB airfoils. This airfoil has been used extensively in low-pressure turbine research, and the present work adds intrapassage pressure and velocity measurements to the existing database. The cascade was tested at design incidence and at an inlet Reynolds number of 126,000 based on inlet midspan velocity and axial chord. Quantitative results include seven-hole pneumatic probe pressure measurements downstream of the cascade to assess blade row losses and detailed seven-hole probe measurements within the blade passage to track the progression of flow structures. Qualitative results take the form of oil surface flow visualization on the endwall and blade suction surface. The application of endwall contouring resulted in lower secondary losses and a reduction in secondary kinetic energy associated with pitchwise flow near the endwall and spanwise flow up the suction surface within the blade passage. The mechanism of loss reduction is discussed in regard to the reduction in secondary kinetic energy.

Advances in Unsteady Boundary Layer Transition Research, Part II: Experimental Verification

This two‐part article presents recent advances in boundary layer research into the unsteady boundary layer transition modeling and its validation. This, Part II, deals with the results of an inductive approach based on comprehensive experimental and theoretical studies of unsteady wake flow and unsteady boundary layer flow. The experiments were performed on a curved plate at a zero streamwise pressure gradient under periodic unsteady wake flow, in which the frequency of the periodic unsteady flow was varied. To validate the model, systematic experimental investigations were performed on the suction and pressure surfaces of turbine blades integrated into a high‐subsonic cascade test facility, which was designed for unsteady boundary layer investigations. The analysis of the experiment′s results and comparison with the model′s prediction confirm the validity of the model and its ability to predict accurately the unsteady boundary layer transition.

Advances in Unsteady Boundary Layer Transition Research, Part I: Theory and Modeling

This two‐part article presents recent advances in boundary layer research that deal with the unsteady boundary layer transition modeling and its validation. A new unsteady boundary layer transition model was developed based on a universal unsteady intermittency function. It accounts for the effects of periodic unsteady wake flow on the boundary layer transition. To establish the transition model, an inductive approach was implemented; the approach was based on the results of comprehensive experimental and theoretical studies of unsteady wake flow and unsteady boundary layer flow. The experiments were performed on a curved plate at a zero streamwise pressure gradient under a periodic unsteady wake flow, where the frequency of the periodic unsteady flow was varied. To validate the model, systematic experimental investigations were performed on the suction and pressure surfaces of turbine blades integrated into a high‐subsonic cascade test facility, which was designed for unsteady boundary layer investigations. The analysis of the experiment′s results and comparison with the model′s prediction confirm the validity of the model and its ability to predict accurately the unsteady boundary layer transition.

The Application of Particle Image Velocimetry (PIV) in a Short-Duration Transonic Annular Turbine Cascade

A series of experiments have been performed to demonstrate the application of Particle Image Velocimetry (PIV) to turbomachinery flows. The tests were performed at transonic speeds on a fully annular engine size turbine nozzle guide vane. The vane cascade was installed in a short-duration Isentropic Light Piston Cascade (ILPC) test facility operating with high inlet turbulence levels. The technique has been shown to map the whole flow field with a resolution of 0.5 mm. The quality of the results obtained is not significantly affected by local turbulence rates. The accuracy of the measurements is put at around 4 percent of absolute velocity and is limited by the quality of the image on the film plane. The velocities derived from the PIV images have been compared with predictions from a three-dimensional viscous numerical calculation. It is shown that the experimental and predicted results are in good agreement. It is considered that this technique has considerable potential in application to turbomachinery flow field diagnostics.

The Effect of Incident Wake Conditions on the Mean Heat Transfer of an Airfoil

The flow phenomena of wakes shed by upstream blade rows is a well-known problem in turbomachinery, which influences blade forces, vibrations, losses, and heat transfer. With respect to the heat load to turbine blades, this problem becomes even more complex because of the interaction between wake, potential flow, and the boundary layer along the surface of the airfoil. Experimentally evaluated mean heat transfer coefficients obtained under different unsteady initial conditions are reported. The heat transfer measurements have been carried out in the cascade test facility at the ITS in Karlsruhe, using a rotating bar wake generator placed upstream of the cascade to simulate the wake passing process. The variation of the wake parameters includes different wake passing frequencies, cascade inlet Reynolds numbers, and wake inclination angles. In addition, the relevant parameters of the unsteady wake have been measured by means of a fixed hot-wire anemometer using the ensemble-average technique. The results are compared to those from the literature for the wake of a cylinder in crossflow. They also serve as experimental base for parallel theoretical analyses.

Annular Cascade Study of Low Back-Pressure Supersonic Fan Blade Flutter

Low back-pressure supersonic fan blade flutter in the torsional mode was examined using a controlled-oscillating annular cascade test facility. Precise data of unsteady aerodynamic forces generated by shock wave movement, due to blade oscillation, and the previously measured data of chordwise distributions of unsteady aerodynamic forces acting on an oscillating blade, were joined and, then, the nature of cascade flutter was evaluated. These unsteady aerodynamic forces were measured by direct and indirect pressure measuring methods. Our experiments covered a range of reduced frequencies based on a semichord from 0.0375 to 0.547, six interblade phase angles, and inlet flow velocities from subsonic to supersonic flow. The occurrence of unstalled cascade flutter in relation to reduced frequency, interblade phase angle, and inlet flow velocity was clarified, including the role of unsteady aerodynamic blade surface forces on flutter. Reduced frequency of the flutter boundary increased greatly when the blade suction surface flow became transonic flow. Interblade phase angles that caused flutter were in the range from 40 to 160 deg for flow fields ranging from high subsonic to supersonic. Shock wave movement due to blade oscillation generated markedly large unsteady aerodynamic forces which stimulated blade oscillation.

Unsteady Aerodynamic Damping Measurement of Annular Turbine Cascade With High Deflection in Transonic Flow

Unsteady aerodynamic forces acting on oscillating blades of a transonic annular turbine cascade were investigated in both aerodynamic stable and unstable domains, using a Freon gas annular cascade test facility. In the facility, whole blades composing the cascade were oscillated in the torsional mode by a high-speed mechanical drive system. In the experiment, the reduced frequency K was changed from 0.056 to 0.915 with a range of outlet Mach number M2 from 0.68 to 1.39, and at a constant interblade phase angle. Unsteady aerodynamic moments obtained by two measuring methods agreed well. Through the moment data the phenomenon of unstalled transonic cascade flutter was clarified as well as the significance of K and M2 for the flutter. The variation of flutter occurrence with outlet flow velocity in the experiments showed a very good agreement with theoretical analysis.