Single-stage Transonic Compressor Research Articles

Study on the Influence of Different Slot Sizes on the Flow Field of Transonic Compressor Rotors

Blade slotting technology is an effective measure to improve the flow structure on the suction surface of a blade and enhance the performance of turbomachinery. To investigate the impact of various slot sizes on the flow field of a single-stage transonic compressor rotor, seven kinds of slot schemes were designed and calculated by numerical simulations. The results show that the above slotting schemes significantly enhance the stability margin of the compressor. In particular, the slotting scheme H9W3 increases the surge margin by 60.9% and slightly reduces peak efficiency by 0.3%, with an almost identical maximum pressure ratio. Slotting promotes high-energy fluid to generate jets from the slot located at the exit of the suction side, effectively controlling blade surface flow separation and reducing channel blockage. Square slots are more effective than elongated slots for controlling separation when using differently shaped slots with equal areas. Increasing slot area gradually decreases outlet total pressure at a constant aspect ratio. A slight increase in the overall blade load causes a backward shift in the front portion load.

Development of Automated Processes for Three-Dimensional Numerical Simulation of Compressor Performance Characteristics

Compressor characteristic evaluation is a critical step in design and optimization. Corrected characteristic curves are typically derived via experimental testing or CFD computation which is typically executed through manual manipulation. For compressors necessitating extensive characteristic computation across multiple speeds and operational conditions, the involved process is inherently complex. This paper introduces an automation approach, employing dichotomy and optimization algorithms aligned with a 3D numerical solver, to streamline the derivation of compressor characteristic curves. Initially, the paper delineated the procedural frameworks for two methodologies and validated them using a single-stage transonic compressor. Both methods achieved the automated resolution of the characteristic lines. The result indicates that the volume of the iterative computations hinges on the granularity of the space searching step and the precision in identifying the stall boundary point. Pertaining to the subject investigated herein, the dichotomy method entailed fewer iterations relative to the optimization algorithm. Subsequent adjustments to the optimization algorithmic process revealed that direct optimization mitigated computational demands substantially by reducing a cycle, thereby attesting to its superior efficacy. In summary, when simulating compressor characteristic curves, the dichotomy method proves more efficient when the interval value of the stall boundary range can be approximated. Conversely, direct optimization is preferable when the estimation of the compressor’s stall boundary is imprecise. The proposed methodologies contribute to compressor research and expedite the progression of compressor development and performance advancement.

Multi-Row Turbomachinery Aerodynamic Design Optimization by an Efficient and Accurate Discrete Adjoint Solver

This paper proposes an approach that combines manual differentiation (MD) and automatic differentiation (AD) to develop an efficient and accurate multi-row discrete adjoint solver. In this approach, the structures of adjoint codes generated using an AD tool are first analyzed. Then, the AD-generated codes are manually adjusted to reduce memory and CPU time consumption. This manual adjustment is performed by replacing the automatically generated low-efficient differentiated codes with manually developed ones. To demonstrate the effectiveness of the proposed approach, the single-stage transonic compressor–NASA Stage 35 and the 1.5-stage Aachen turbine–are used. The solution information exchange at a rotor-stator/stator-rotor interface is achieved by a conservative, non-reflective, and robust discrete adjoint mixing plane method. The results show that the discrete adjoint solver developed by hybrid automatic and manual differentiation is more economical in computational cost than that developed purely by an AD tool and has higher sensitivity accuracy than the adjoint solver with the constant eddy viscosity (CEV) assumption. Moreover, the multi-row turbomachinery design optimizations can be efficiently performed by the discrete adjoint solver developed by the hybrid automatic and manual differentiation.

A Combined Artificial-Intelligence Aerodynamic Design Method for a Transonic Compressor Rotor Based on Reinforcement Learning and Genetic Algorithm

An aircraft engine’s performance depends largely on the compressors’ aerodynamic design, which aims to achieve higher stage pressure, efficiency, and an acceptable stall margin. Existing design methods require substantial prior knowledge and different optimization algorithms to determine the 2D and 3D features of the blades, in which the design policy needs to be more readily systematized. With the development of artificial intelligence (AI), deep reinforcement learning (RL) has been successfully applied to complex design problems in different domains and provides a feasible method for compressor design. In addition, the applications of AI methods in compressor research have progressively developed. This paper described a combined artificial-intelligence aerodynamic design method based on a modified deep deterministic policy gradient algorithm and a genetic algorithm (GA) and integrated the GA into the RL framework. The trained agent learned the design policy and used it to improve the GA optimization result of a single-stage transonic compressor rotor. Consequently, the rotor exhibited a higher pressure ratio and efficiency owing to the sweep feature, lean feature, and 2D airfoil angle changes. The separation near the tip and the secondary flow decreased after the GA process, and at the same time, the shockwave was weakened, providing improved efficiency. Most of these beneficial flow field features remained after agent modification to improve the pressure ratio, showing that the policy learned by the agent was generally universal. The combination of RL and other design optimization methods is expected to benefit the future development of compressor designs by merging the advantages of different methods.

Impact of Boundary Layer Suction on Clearance Leakage Flow in a Cantilever Stator of Transonic Compressor

To control secondary flow effects and enhance the aerodynamic performance of the compressor, the flow control effects of the flow suction at the endwall with different circumferential positions and at the blade tip were numerically investigated in the cantilever stator of an axial single-stage transonic compressor. The main purpose was to gain a better understanding of the application of boundary layer suction and the associated control mechanisms in the cantilever stator. The studies show that the optimal position of the endwall suction slot should be located up the stator blade, in terms of the leakage flow structures and the blade tip unloading effect. In addition, the flow control effects of the suction at the blade tip on leakage flow upstream is better than that of the endwall flow suction with the same structure. Further, the studies of compressor aerodynamic performance curves illustrate that the efficiency and pressure ratio is increased by 0.34% and 1.09% at the peak efficiency point, and are increased by 0.39% and 0.14% at the near stall point, respectively.

Improvements of performance and stability of a single-stage transonic axial compressor using a combined flow control approach

The performance and the stability of axial compressors are seriously concerned as it's widely used in aero-engines and power machines. A new combined flow control approach was developed previously using blade slot and vortex generator, which was used to reduce flow loss and improve flow stability. The approach was experimentally proved to be very effective in linear compressor cascades. To further evaluate its availability in actual compressors, the new approach is implemented in a single-stage transonic compressor, which introduces a part-span slot into the rotor while a full-span slot and a vortex generator into the stator. CFD simulations are performed to evaluate the gains of compressor performance, and flow structures are analyzed to explain the corresponding mechanisms. Results show great improvements of performance and stability of the compressor. The averaged total pressure ratio and isentropic efficiency improve by 1.82% and 0.88% respectively due to the great reduction of separations in both the rotor and the stator. Because the compressor instability is initiated by the stator corner stall, the delay of the corner separations leads to a great increase of the stable mass-flow range by 26.1%. The compressor flow structures are also well improved because the detrimental backflows and large-scale vortices are mostly removed by the combined approach.

Investigation of an innovative non-free vortex aerodynamic procedure to design a single-stage transonic compressor

This paper presents an innovative design method for a transonic compressor based on the radial equilibrium theory by means of increasing blade loading. Firstly, the rotor blade of a transonic compressor is redesigned based on the constant spanwise de-Haller number and diffusion. The design method leads to an unconventional increased axial velocity distribution in tip section, which originates from non-uniform enthalpy distribution assumption. A code is applied to extract the compressor meridional plane and blade-to-blade geometry containing rotor and stator in order to design the blade three-dimensional view. A structured grid is generated for the numerical domain of fluid. Finer grids are used for the regions near walls to capture the boundary layer effects and behavior. Reynolds-averaged Navier–Stokes equations are solved by finite volume method for rotating zones (rotor) and stationary zones (stator). The experimental data, available for the performance map of NASA Rotor67, is used to validate the results of the current simulations. Then, the capability of the design method is validated by computational fluid dynamics that is capable of predicting the performance map. The numerical results of the new geometry by representing 11% improvement in efficiency and 19% in total pressure ratio verify the new method advantages. The computational fluid dynamics results also show that the newly designed rotor blades due to a higher velocity in the tip section have a special capacity to increase the loading without any separation. The mass flow reduction is observed in the new geometry, which could be easily improved by changing stagger angle.

One-Dimensional Performance Improvement of a Single-Stage Transonic Compressor

One-Dimensional Performance Improvement of a Single-Stage Transonic Compressor

Further Investigation on Transonic Compressor Stall Margin Enhancement With Stall Precursor-Suppressed Casing Treatment

A kind of casing treatment, named as stall precursor-suppressed (SPS), has been developed recently, which was proved to be able to effectively improve stall margin (SM) without significant efficiency loss in low-speed axial flow compressors and a transonic compressor rotor. In this paper, the effectiveness of the SPS casing treatment is investigated in a single-stage transonic compressor. Based on an extended stall inception model, the quantitative evaluation of the SM enhancement by the SPS casing treatment is presented for the transonic compressor stage. The model predicts that a 2.5–6.8% of stall margin improvement (SMI), which is defined in terms of mass flow rate at stall inception, can be achieved at the design rotational speed. The experimental results show that the SPS casing treatment can achieve 3.5–9.3% of the SMI at 95% design rotational speed. Due to the fact that the distributions of the total pressure ratio along the spanwise direction are kept the same as those of the solid wall casing at the same mass flow rate, the SPS casing treatments with a small open area ratio and large backchamber enhance the SM without a recognizable efficiency loss and a migration of the pressure-rise characteristics. Furthermore, the mechanism of SMI with the SPS casing treatment is investigated in the experiments. In comparison with the solid wall casing, the emergence and the evolution of the stall inception waves are suppressed and the nonlinear development of the stall process is delayed with the SPS casing treatment.

Analysis and Prediction of Shock-Induced Vortex Circulation in Transonic Compressors

Multiple high-fidelity time-accurate computational fluid dynamics simulations were performed to investigate the effects of upstream stator loading and rotor shock strength on vortex shedding characteristics in a single-stage transonic compressor. Three loadings on the upstream stator row of decreased, nominal, and increased loading in conjunction with three axial spacings of close, mid, and far were studied for this analysis. The time-accurate urans code turbo was used to generate periodic, quarter annulus simulations of the blade row interaction (BRI) compressor rig. It was observed that vortex shedding was synchronized to the passing of a rotor bow shock. Results show that vortex strength increases linearly with stator loading and rotor bow shock strength. “Normal” and “large” shock-induced vortices formed on the stator trailing edge (TE) immediately after the shock passing, but the large vortices were strengthened at the TE due to a low-velocity region on the suction surface. This low-velocity region was generated upstream on the suction surface from a shock-induced thickening of the boundary layer or separation bubble. The circulation of the large vortices was greater than the normal vortices by a factor of 1.7, 1.83, and 2.04 for decreased, nominal, and increased deswirler loadings. At decreased loading, only 24% of the measured vortices were considered large, while at nominal loading 58% were large. A model was developed to predict shock-induced vortex circulation from a known rotor bow shock strength, stator diffusion factor, and near-wake parameters. The model predicts the average vortex circulation very well, with 5% difference between predicted and measured values. An understanding of the unsteady interactions associated with blade loading and rotor shock strength in transonic stages will help compressor designers account for unsteady flow physics at design and off-design operating conditions.

Experimental Investigation of Tip Clearance Flow in a Transonic Compressor With and Without Plasma Actuators

Blade tip losses represent a major performance penalty in low aspect ratio transonic compressors. This paper reports on the experimental evaluation of the impact of tip clearance with and without plasma actuator flow control on performance of an U.S. Air Force-designed low aspect ratio, high radius ratio single-stage transonic compressor rig. The detailed stage performance measurements without flow control at three clearance levels, classified as small, medium, and large, are presented. At design-speed, increasing the clearance from small to medium resulted in a stage peak efficiency drop of almost six points with another four point drop in efficiency with the large clearance (LC). Comparison of the speed lines at high-speed show significantly lower pressure rise with increasing tip clearance, the compressor losing 8% stall margin (SM) with medium clearance (MC) and an additional 1% with the LC. Comparison of the stage exit radial profiles of total pressure and adiabatic efficiency at both part-speed and design-speed and with throttling are presented. Tip clearance flow-control was investigated using dielectric barrier discharge (DBD) type plasma actuators. The plasma actuators were placed on the casing wall upstream of the rotor leading edge and the compressor mapped from part-speed to high-speed at three clearances with both axial and skewed configurations at six different frequency levels. The plasma actuators did not impact steady state performance. A maximum SM improvement of 4% was recorded in this test series. The LC configuration benefited the most with the plasma actuators. Increased voltage provided more SM improvement. Plasma actuator power requirements were almost halved going from continuous operation to pulsed plasma. Most of the improvement with the plasma actuators is attributed to the reduction in unsteadiness of the tip clearance vortex near-stall resulting in additional reduction in flow prior to stall.

수치해석을 활용한 1단 천음속 압축기 내부 유동장 분석

Numerical simulations on a single stage transonic compressor which is developed by Korea Aerospace Research Institute are carried out and their results are compared with experimental data for cross validations. Comparisons between experimental data and numerical simulation results show good agreements on a performance curve, static pressure and total pressure distributions. CFD results show that there is a clear interaction between tip leakage flow and normal shock in the rotor passage. Tip leakage flows are almost dissipated after the strong normal shock and it forms a strong recirculation near the blade tip. Also a large separation region grows on the suction surface just after the normal shock. As the pressure ratio and blade loading increase, the normal shock line moves upstream and it starts to deviate from the blade leading edge. Then the tip leakage flow does not overcome the strong adverse pressure gradient and flow blockage originated from the tip recirculation region. As a result, the tip leakage flow heads for the neighboring blade leading edge, which results in a compressor stall.

Constrained Adjoint-Based Aerodynamic Shape Optimization of a Single-Stage Transonic Compressor

This paper develops a discrete adjoint formulation for the constrained aerodynamic shape optimization in a multistage turbomachinery environment. The adjoint approach for viscous internal flow problems and the corresponding adjoint boundary conditions are discussed. To allow for a concurrent rotor/stator optimization, a nonreflective adjoint mixing-plane formulation is proposed. A sequential-quadratic programming algorithm is utilized to determine an improved airfoil shape based on the objective function gradient provided by the adjoint solution. The functionality of the proposed optimization method is demonstrated by the redesign of a midspan section of a single-stage transonic compressor. The objective is to maximize the isentropic efficiency while constraining the mass flow rate and the total pressure ratio.

Numerical Investigation of the Unsteady Flow in a Transonic Compressor with Curved Rotors

The unsteady 3D flow fields in a single-stage transonic compressor under designed conditions are simulated numerically to investigate the effects of the curved rotors on the stage performance and the aerodynamic interaction between the blade rows. The results show that, compared to the compressor with un-curved rotors, the compressor under scrutiny acquires remarkable increases in efficiency with significantly reduced amplitudes of the time-dependent fluctuation. The amplitude of the pressure fluctuation around the stator leading edge decreases at both endwalls, but increases at the mid-span in the curved rotors. The pressure fluctuation near the stator leading edge, therefore, becomes more uniform in the radial direction of this compressor. Except for the leading edge area, the pressure fluctuation amplitude declines remarkably in the tip region of stator surface downstream of the curved rotor, but hardly changes in the middle and at the hub.

Effect of forward sweep in a transonic compressor rotor

This paper presents the design and testing of a transonic compressor rotor with forward sweep. The rotor was used to investigate the influence of forward sweep on the performance and stability of a single-stage transonic compressor compared with a baseline design with radially stacked blade sections. The comparison was done numerically with the three-dimensional Navier—Stokes code TRACE-S and experimentally in the Darmstadt transonic compressor test rig. It was found that the new rotor with forward sweep has an increased efficiency and also a much better stall margin (much more in the rig test than predicted by the three-dimensional Navier—Stokes calculation). Particularly close to stall, the forward sweep diverts the flow towards the blade tip region which helps to stabilise this region. For that reason it is possible to throttle the forward-swept rotor much further than the radially stacked rotor, although the forward-swept rotor does already suffer from separated flow in the hub.

Darmstadt Rotor No. 2, II: Design of Leaning Rotor Blades

For Darmstadt University of Technology′s axial singlestage transonic compressor rig, a new three‐dimensional aft‐swept rotor was designed and manufactured at MTU Aero Engines in Munich, Germany. The application of carbon fiber–reinforced plastic made it possible to overcome structural constraints and therefore to further increase the amount of lean and sweep of the blade. The aim of the design was to improve the mechanical stability at operation that is close to stall.To avoid the hazard of rubbing at the blade tip, which is found especially at off‐design operating conditions close to the stability limit of the compression system, aft‐sweep was introduced together with excessive backward lean.This article reports an investigation of the impact of various amounts of lean on the aerodynamic behavior of the compressor stage on the basis of steady‐state Navier‐Stokes simulations. The results indicate that high backward lean promotes an undesirable redistribution of mass flow and gives rise to a basic change in the shock pattern, whereas a forward‐leaning geometry results in the development of a highly back‐swept shock front. However, the disadvantage is a decrease in shock strength and efficiency.

Application of pressure sensitive paint measurement to a low-solidity cascade diffuser of a transonic centrifugal compressor

A pressure sensitive paint (PSP) measurement has been known as a pressure field measurement technique based on the oxygen quenching phenomenon of luminescence of specific luminophores. A PSP measurement was applied for pressure field measurement in a low-solidity circular cascade diffuser of a single-stage transonic centrifugal compressor with 5 in pressure ratio for HFC134a gas. The oxygen concentration was about 500 ppm. Ru (bath-phen) was adsorbed on a silica-gel thin-layer chromatography sheet, and the sheet was pasted onto the side-wall between the cascade vanes. A drastic change in luminescent intensity was recognized during a surge condition. Also the pressure variations based on luminescent intensity agreed well with the pressure fluctuations measured using a semiconductor pressure sensor with high-frequency-response. It was shown that a PSP measurement worked well to investigate the unsteady pressure fields in a circular cascade diffuser of a transonic centrifugal compressor. Moreover, the time response of PSP becomes clear as a problem to be overcome for the present.

Numerical Simulation of the Shock-Tip Leakage Vortex Interaction in a HPC Front Stage

For a single-stage transonic compressor rig at the TU Darmstadt, three-dimensional viscous simulations are compared to L2F measurements and data from the EGV leading edge instrumentation to demonstrate the predictive capability of the Navier–Stokes code TRACE_S. In a second step the separated regions at the blade tip are investigated in detail to gain insight into the mechanisms of tip leakage vortex-shock interaction at operating points close to stall, peak efficiency, and choke. At the casing the simulations reveal a region with axially reversed flow, leading to a rotationally asymmetric displacement of the outermost stream surface and a localized additional pitch-averaged blockage of approximately 2 percent. Loss mechanisms and streamline patterns deduced from the simulation are also discussed. Although the flow is essentially three-dimensional, a simple model for local blockage from tip leakage is demonstrated to significantly improve two-dimensional simulations on S1-surfaces.