Aerodynamic Investigations of a Variable Inlet Guide Vane With Symmetric Profile

In order to determine the aerodynamic behavior of a Variable Inlet Guide Vane as used in multishaft compressors, extensive experimental investigations with a 2D linear cascade have been conducted. All the experiments were performed at the High-Speed Cascade Wind Tunnel at the Institute of Jet Propulsion. They covered a wide range of Reynolds numbers and stagger angles as they occur in realistic turbomachines. Within this work at first the observed basic flow phenomena (loss development, overturning) will be explained. For the present special case of a symmetric profile and a constant decreasing chord length along the vane height, statements about different spanwise position can be made by investigating different Reynolds numbers. The focus of this paper is on the outflow of the VIGV along the vane height. Results for an open flow separation on the suction side are presented, too. Stall condition can be delayed by boundary layer control. This is done using a wire to trigger an early boundary layer transition. The outcomes of the trip wire measurement are finally discussed. The objective of this work is to evaluate the influence of the stagger angle and Reynolds number on the total pressure losses and the deviation angle. The results of the work presented here, gives a better insight of the efficient use of a VIGV.

Aerodynamic and Aeroacoustic Optimization of a Transonic Centrifugal Compressor

The performance of transonic compressors can be characterized aerodynamically and aeroacoustically. In this paper, the DLR SRV2 compressor without vaned diffusers and its redesigned version are studied. The redesign strategy (Zangeneh et al. 2011 [1]) utilized the 3D inverse design and CFD analysis. Both compressors were analyzed in ANSYS CFX 11, and the computational results show that the predicted pressure-ratio and efficiency of the original compressor have good agreement with experimental results. The simulations have also revealed that the redesigned one is superior at both design and off-design points at different rotating speeds. This work applies a convective FW-H method to further investigate the noise radiation from these two compressors. As the blade tip speed is supersonic, the permeable integral surface scheme must be adopted. The flow quantities needed as the inputs to the FW-H solver were extracted from the CFD solutions. The numerical predictions of the noise SPLs at blade passing frequency and its harmonics match the experimental measurements reasonably well. It is found that the original compressor has significant variations of SPLs as the operating mass flow rate changes whereas the redesigned one has much slighter variations. At peak efficiency the redesigned compressor has a lower noise level. This study provides insights for the optimal design of a transonic compressor when good aerodynamic and aeroacoustic performance are both required.

On the Recessed Blade Tip With and Without a Porous Material in an Axial Compressor

This paper addresses the effect of the recessed blade tip with and without a porous material on the performance of a transonic axial compressor. A commercial flow solver was employed to analyze the performance and the internal flow of the axial compressor with three different tip configurations: reference tip, recessed tip and recessed tip filled with a porous material. It was confirmed that the recessed blade tip is an effective method to increase the stall margin in an axial compressor. It was also found in the present study that the strong vortex formed in the recess cavity on the tip pushed the tip leakage flow backward and weakened the tip leakage flow itself, consequently increasing the stall margin without any penalty of the efficiency in comparison to the reference tip. The recessed blade tip filled with a porous material was suggested with hope to obtain the larger stall margin and the higher efficiency. However, it was found that a porous material in the recess cavity is unfavorable to the performance in both the stall margin and the efficiency. An attempt has been made to explain the effect of the recess cavity with and without a porous material on the flow in an axial compressor.

Robust Compressor Blades for Desensitizing Operational Tip Clearance Variations

Robust design is a multi-objective optimization framework for obtaining designs that perform favorably under uncertainty. In this paper robust design is used to redesign a highly loaded, transonic rotor blade with a desensitized tip clearance. The tip gap is initially assumed to be uncertain from 0.5 to 0.85% span, and characterized by a beta distribution. This uncertainty is then fed to a multi-objective optimizer and iterated upon. For each iteration of the optimizer, 3D-RANS computations for two different tip gaps are carried out. Once the simulations are complete, stochastic collocation is used to generate mean and variance in efficiency values, which form the two optimization objectives. Two such robust design studies are carried out: one using 3D blade engineering design parameters (axial sweep, tangential lean, re-cambering and skew) and the other utilizing suction and pressure side surface perturbations (with bumps). A design is selected from each Pareto front. These designs are robust: they exhibit a greater mean efficiency and lower variance in efficiency compared to the datum blade. Both robust designs were also observed to have significantly higher aft and reduced fore tip loading. This resulted in a weaker clearance vortex, wall jet and double leakage flow, all of which lead to reduced mixed-out losses. Interestingly, the robust designs did not show an increase in total pressure at the tip. It is believed that this is due to a trade-off between fore-loading the tip and obtaining a favorable total pressure rise and higher mixed-out losses, or aft-loading the tip, obtaining a lower pressure rise and lower mixed-out losses.

Numerical Investigation on the Effect of Vortex Generator on Axial Compressor Performance

The performance of the compressor blade is considerably influenced by secondary flow effects, like the cross flow on the end wall as well as corner flow separation between the wall and the blade. The present work is focused on the studying the effects of Vortex Generator (VG) on NASA Rotor 37 test case using Computational Fluid Dynamics (CFD). VG helps in controlling the inception of the stall by generating vortices and energizes the low momentum boundary layer flow which enhances the rotor performance. Three design configuration namely, Counter-rotating, Co-rotating and Plow configuration VG are selected based on the improved aerodynamic performance discussed in reference [1]. These VG are located at 90% span and 42% chord on suction side surface of the blade. Among the three configurations, the first configuration has greater impact on the end wall cross flow and flow deflection which resulted in enhanced numerical stall margin of 5.4% from baseline. The reasons for this numerical stall margin improvement are discussed in detail.

Numerical Investigation of the Solidity Effect on Linear Compressor Cascades

Solidity in compressors is defined as the ratio of the aerodynamic chord over the peripheral distance between two adjacent blades, the pitch. This parameter is simply the inverse of the pitch-to-chord ratio generally used in turbines. Solidity must be selected at the earliest design phase, i.e. at the level of the meridional design and represents a crucial step in the whole design process. Most of the existing studies on this topic rely on low-speed compressor cascade correlations from Carter or Lieblein. The aim of this work is to update those correlations for state-of-the-art controlled diffusion blades, and extend their application to high Mach number flow regimes more typical of modern compressors. Another objective is also to improve the physical understanding of the solidity effect on compressor performance and stability. A numerical investigation has been performed using the commercial software FINE/Turbo. Two different blade profiles were selected and investigated in the compressible flow regime as an extension to the low-speed data on which the correlations are based. The first cascade uses a standard double circular arc profile, extensively referenced in the literature, while the second configuration uses a state-of-the-art CDB, representative of low pressure compressor stator mid-span profile. Both profiles have been designed with the same inlet and outlet metal angles and the same maximum thickness but the camber and thickness distributions, the stagger angle and the leading edge geometry of the CDB have been optimized. The determination of minimum loss, optimum incidence and deviation is addressed and compared with existing correlations for both configurations and various Mach numbers that have been selected in order to match typical booster stall and choke operating conditions. The emphasis is set on the minimum loss performance at mid-span. The impact of the solidity on the operating range and the stability of the cascade are also studied.

Numerical Investigation on the Effects of Circumferential Coverage of Injection in a Transonic Compressor With Discrete Tip Injection

Discrete tip injection upstream of the rotor tip is an effective technique to extend stability margin for a compressor system in an aeroengine. The current study investigates the effects of injectors’ circumferential coverage on compressor performance and stability using time-accurate three-dimensional numerical simulations for multi passages in a transonic compressor. The percentage of circumferential coverage for all the six injectors ranges from 6% to 87% for the five investigated configurations. Results indicate that circumferential coverage of tip injection can greatly affect compressor stability and total pressure ratio, but has little influence on adiabatic efficiency. The improvement of compressor total pressure ratio is linearly related with the increasing circumferential coverage. The unsteady flow fields show that there exists a non-ignorable time lag of the injection effects between the passage inlet and outlet, and blade tip loading will not decline until the injected flow reaches the passage outlet. Stability improves sharply with the increasing circumferential coverage when the coverage is less than 27%, but increases flatly for the rest. It is proven that the injection efficiency which is a measurement of averaged blockage decrement in the injected region is an effective guideline to predict the stability improvement.

Calculation of Flow Instability Inception in High Speed Axial Compressors

This paper applies a theoretical model developed recently to calculate the flow instability inception point in axial high speed compressors system. After the mean flow field is computed by steady CFD simulation, a body force approach, which is a function of flow field data and comprises of one inviscid part and the other viscid part, is taken to duplicate the physical sources of flow turning and loss. Further by applying appropriate boundary conditions and spectral collocation method, a group of homogeneous equations will yield from which the stability equation can be derived. The singular value decomposition method is adopted over a series of fine grid points in frequency domain, and the onset point of flow instability can be judged by the imaginary part of the resultant eigenvalue. The first assessment is to check the applicability of the present model on calculating the stall margin of one single stage transonic compressors at 85% rotational speed. The reasonable prediction accuracy validates that this model can provide an unambiguous judgment on stall inception without numerous requirements of empirical relations of loss and deviation angle. It could possibly be employed to check over-computed stall margin during the design phase of new high speed fan/compressors. The following validation case is conducted to study the nontrivial role of tip clearance in rotating stall, and a parameter study is performed to investigate the effects of end wall body force coefficient on stall onset point calculation. It is verified that the present model could qualitatively predict the reduced stall margin by assuming a simplified body force model which represents the response of a large tip clearance on the unsteady flow field.

Experimental Investigation on the Effect of Porosity of Bend Skewed Casing Treatment on a Single Stage Transonic Axial Flow Compressor

A bend skewed casing treatment was designed, to study the influence of one of its geometrical parameter porosity on the stable performance of single stage transonic axial flow compressor. The compressor was designed for the stage total-to-total pressure ratio of 1.35, corrected mass flow rate of 22 kg/s at corrected design speed of 12930 RPM. Bend skewed casing treatment has an axial inlet segment till 50% of the total length and rear segment that is skewed by 45° in the direction of the rotor tip section stagger. Both the sections were oriented at a skew angle of 45° to the radial plane such that the flow exiting the slot is in counter-clockwise direction to that of the rotor direction. The casing treatment slot width was equal to the maximum thickness of the rotor blades. Three casing treatment configurations were identified for the current experimental investigation. All the treatment geometries considered for the experimental research have lower porosities than reported in the open literatures. The effect of the porosity parameter on the performance of transonic compressor stage was evaluated at two axial coverages of 20% and 40% relative to the rotor tip axial chord. Performance maps were obtained for the solid casing and casing treatment with three different porosities. Comparative studies were carried out and experimental results showed a maximum of 65% improvement in the stable operating range of the compressor for one of the treatment configurations. It was also observed that the stable operating range of the compressor increases with an increase in the casing treatment porosity. All the casing treatment configurations showed that the compressor stall occurs at lower mass flows as compared to the solid casing. Compressor stage peak efficiency shows significant degradations with increase in the porosity as compared to solid casing. Detailed blade element performances were also obtained using calibrated multi-hole aerodynamic probe. Comparative variations of flow parameters like absolute flow angle, Mach number were studied at full flow and near stall conditions for the solid casing and casing treatment configurations. Hot wire measurements show very high fluctuation in the inlet axial velocity in the presence of solid casing as compared to casing treatments. Experimental investigation revealed that the porosity of the casing treatments has strong influence on the transonic compressor stage performance.

Endwall Contouring and Fillet Design for Reducing Losses and Homogenizing the Outflow of a Compressor Cascade

The following paper deals with the development of an optimized fillet and an endwall contour for reducing the total pressure loss and for homogenizing the outflow of a highly loaded cascade with a low aspect ratio. The NACA-65 K48 cascade profile without a fillet and without endwall contouring is used as a basis. Optimizations are performed using the DLR in-house tool AutoOpti and the RANS-solver TRACE. Three operating points at an inflow Mach number of 0.67 with different inflow angles are used to secure a wide operating range of the optimized design. At first only a fillet is optimized. The optimized fillet is small at the leading edge and rather high, wide and thick towards the trailing edge. It reduces the total pressure loss and homogenizes the outflow up to a blade height of 20 %. Following this a combined optimization of the endwall and the fillet is performed. The optimized contour leads to the development of a vortex, which changes the secondary flow in such a way, that the corner separation is reduced, which in turn significantly reduces the total pressure loss up to 16 % in the design operating point. The contour in the outflow region leads to a significant homogenization of the outflow in the near wall region.