Supersonic Compressor Cascade Research Articles

An efficient sparse surrogate model for aerodynamic characteristics of a supersonic compressor cascade with uncertain geometric deformations

The increasing demand for high-load compressors necessitates research and design of supersonic compressor blades. Geometric deformations of the blades resulting from manufacturing errors introduce uncertainties in their aerodynamic characteristics. For uncertainty analysis, Polynomial Chaos Expansion (PCE) has gained popularity among engineers in various disciplines. However, as the number of random variables used to describe uncertainties increases, the efficiency of the PCE method diminishes. To overcome this limitation, this paper firstly proposes an Efficient Sparse Surrogate Model (ESSM). The ESSM combines PCE to approximate global characteristics and Gaussian process modeling to capture local variability, resulting in a highly accurate model. Additionally, by utilizing an iteratively diffeomorphic modulation under observable response preserving homotopy, important PCE basis functions are adaptively selected, leading to a sparse surrogate model that significantly reduces the required training samples. Then using the ESSM, efficient uncertainty quantification (UQ) and global sensitivity analysis were conducted to evaluate the impact of uncertain geometric deformations on the aerodynamic characteristics of a supersonic compressor cascade. Principal Component Analysis (PCA) was employed in conjunction with an Improved Class Shape Transformation (ICST) to characterize geometric deformations caused by manufacturing errors. The proposed ICST method exhibits enhanced accuracy in fitting near the leading and trailing edges compared to the CST method. PCA enables a 50% reduction in the number of random variables needed to describe geometric deformations. The UQ results establish a quantitative correlation between geometric deformations and supersonic aerodynamics, and the sensitivity analysis identifies some specific forms of geometric deformations that significantly impact the aerodynamics. Finally, the study investigated the flow mechanisms associated with typical geometric deformation forms, further providing recommendations for manufacturing and testing of supersonic blades to prevent performance deterioration.

Summary of Research Progress in Supersonic Compressor Cascade

In compressors, supersonic/transonic flow is a hot topic in today's research. Its development directly affects the performance and function of compressors. The main purpose of supersonic cascade research is to increase its tip velocity, so as to improve its power capacity and boost ratio. In this paper, important research achievements since 2017 are summarized in the following aspects: the working principle of ultrasonic elements, unique intake Angle, shock wave structure and boundary layer interference. At the same time, the most widely used design blade shapes are introduced. Finally, the future development is prospected.

Trailing-edge shock loss control with self-sustaining synthetic jet in a supersonic compressor cascade

To effectively reduce the loss of strong shock wave at the trailing edge of the supersonic cascade under high backpressure, a shock wave control method based on self-sustaining synthetic jet was proposed. The self-sustaining synthetic jet was applied on the pressure side of the blade with the blow slot and the bleed slot arranged upstream and downstream of the trailing-edge shock, respectively. The flow control mechanism and effects of parameters were investigated by numerical simulation. The results show that the self-sustaining synthetic jet forms an oblique shock wave in the cascade passage which slows down and pressurizes the airflow, and the expansion wave downstream of the blow slot weakens the shock strength which can effectively change the Mach reflection to regular reflection and thus weaken the shock loss. And the suction effect can reduce loss near blade surface. Compared with the baseline cascade, the self-sustaining jet actuator can reduce flow losses by 6.73% with proper location design and vibration of diaphragm.

Research on the Influence of Surface Roughness on the Flow Field in the Cascade

Fouling, corrosion, and icing are common causes of increased compressor blade roughness due to special operating conditions and natural environment. Roughness increase would influence on flow field, especially in supersonic compressor cascades where incident waves can interact with reflected shock waves. Studies have shown that the boundary layer condition of the blade varies with wall roughness, which affects both the incident shock wave and the reflected shock wave. In order to study the effect of the distribution of roughness, the flow field under different roughness settings is simulated. The results show that flow loss is reduced with increasing roughness at the incidence point of an incident wave and the trailing edge of the suction side.

Supersonic compressor cascade flow control using plasma actuation at low Reynolds number

To control the supersonic compressor cascade flow at low Reynolds numbers, this paper describes the use of nanosecond dielectric barrier discharge (NS-DBD) plasma actuation for flow control, and presents the results of large-eddy simulations conducted to investigate the corresponding flow control effects. NS-DBD plasma actuation on both the blade pressure surface and suction surface induces a distorted flow structure (DFS) within the blade passage. In the case of NS-DBD plasma actuation on the blade pressure surface, the influence of the DFS on the flow is suppressed by a shock wave. Even so, the DFS can still trigger the instability in the shear layer between the separated flow and the mainstream flow. Shock-wave-induced large-scale flow separation on the blade pressure surface is then suppressed, and the overall total pressure loss of the blade passage is reduced by 7.4%, despite the increased shock wave loss from the reduced flow blockage within the blade passage. In the case of NS-DBD plasma actuation on the blade suction surface, the DFS is less effective in suppressing the shock-wave-induced small-scale flow separation on the blade suction surface. However, the DFS on the blade suction surface enhances the shock wave oscillations within the blade passage, and this suppresses the flow separation on the blade pressure surface.

Effect of a blade end slot on supersonic compressor cascade hub-corner separation

The strong adverse pressure gradient and shock wave/boundary layer interaction that occurs in supersonic compressor cascade corner regions usually leads to large three-dimensional corner separations with extensive total pressure losses. This substantially restricts compressor load performance. In this paper, the effect of a blade end slot on corner separation caused by shock wave/boundary layer interaction was investigated. Because of the pressure difference between the pressure and suction sides, the slot enables one to induce jet flow into the endwall/blade suction-side corner region. Four different blade end slot locations are designed and their influences are investigated numerically. The mechanism of blade end slot influence on corner separation varies by position. The optimal location is slightly downstream of the initial separation point, which improves corner separation control and can reduce the cascade total pressure loss by 9.7% using the currently prescribed slot size.

Flow Control Effect of Spanwise Distributed Pulsed Arc Discharge Plasma Actuation on Supersonic Compressor Cascade Flow

To achieve efficient control of supersonic compressor cascade flow, a type of spanwise distributed pulsed arc discharge plasma actuation (PADPA) was designed. To simulate the influences of PADPA on the flow field, a phenomenological model was established. Then, the flow control effects of PADPA on supersonic compressor cascade flow were researched numerically. The results show that under low static pressure ratio condition, the compressive wave induced by PADPA reduced the intensity of the passage shock wave, which eventually reduced shock wave loss. It was also found that PADPA produced an adverse pressure gradient (pre-compression effect) around the actuation location, which reduced the strength of the high adverse pressure gradient induced by the passage shock wave. The airflow on both sides of the actuation location was accelerated by PADPA owing to the spanwise distributed layout. Thus, it improved the ability of the boundary layer to resist the effect of the adverse pressure gradient and reduced the separation zone. Consequently, the total pressure loss was reduced by 6.8%. Under high pressure ratio condition, the effect of PADPA on the suction side controlling the large separation of the boundary layer was insignificant. The total pressure loss also increased slightly.

Flow mechanism and aspiration strategies in an ultra-highly loaded supersonic compressor cascade

In order to improve the performance and eliminate flow separations, boundary layer suction/aspiration, was carried out in an ultra-highly loaded (UHL) supersonic compressor cascade with a super-high diffusion (D) factor of 0.81. Firstly, the numerical simulation method is validated using existing experimental data. Secondly, three different suction schemes of blade suction surface are designed and investigated numerically under two-dimensional (2D) conditions. Thirdly, six different endwall suction schemes are designed and calculated, and an optimal endwall suction scheme is obtained. At last, suction surface coupled with endwall suction is integrated in the UHL supersonic compressor cascade so as to eliminate both trailing edge separation and corner separation.Under 2D conditions, the UHL supersonic compressor cascade experiences a strong normal shock wave with pre-Mach number of 1.6, and significant flow separation exists inside the blade passage under shock wave/boundary layer interaction. Due to the flow separation near the suction surface, the normal shock wave becomes an oblique shock wave at the shock wave foot. By arranging suction slot at 30% to 70% axial chord on suction surface, trailing edge separations can be effectively eliminated. Under three-dimensional (3D) conditions, an optimal endwall suction slot is distributed from the shock foot to the trailing edge near the suction surface. With the optimal endwall suction scheme, corner separation is effectively removed. However, after endwall suction, flow separation near mid-span is increased considerably. With appropriate suction mass flow, the separation along spanwise direction can be eliminated and the flow field inside the cascade becomes a 2D pattern. By adopting the coupled suction surface and endwall suction, both corner separation and trailing edge separation are eliminated effectively, and the diffusion ability of the supersonic compressor cascade is also improved significantly. The suction effect of suction slot on suction surface is similar to that of 2D conditions.

Large Eddy Simulation of Pulsed Blowing in a Supersonic Compressor Cascade

Large Eddy Simulation (LES) of a two dimensional supersonic compressor cascade is performed in the current study. It is found that the Shock Wave Boundary Layer Interaction causes a large scale of total pressure losses and presents strong fluctuation features. Thus the pulsed and steady excitation jets are applied to suppress the flow separations and to reduce the total pressure losses. Several impacting parameters, such as jet axial location, jet hole width, jet angle to the local blade surface and jet mass flowrate are chosen based on the primary analysis by the calculations by the Reynolds Averaged Navier-Stokes equations. In addition, based on the results of frequency spectrum and POD analysis, the excitation jet frequency is chosen for the pulsed excitation jet scheme. It is concluded that the pulsed excitation jet scheme achieves a 9.8% reduction of total pressure loss in comparison to the steady excitation jet scheme under the same time-averaged excitation jet mass flow rate. The excitation jets affect both the flow field near the jet hole on the suction surface and the flow field on the pressure surface via the management of the reflection shock wave. In addition, the excitation frequency dominates not only the time-averaged flow field, but also the second and third modes which stand for the unsteady structures in the flow field under the POD analysis. The first mode contains most energy in the flow field and the energy percentage decreases dramatically with the increase of the mode number. In comparison to the steady excitation jet scheme, the pulsed excitation jet scheme gathers more energy to the low orders of the modes, especially the first four modes. With the mixing effect and high dissipation rate of the high-frequency signals, the high-frequency signals shrink in the wake and the flow field builds up more uniformity.

Supersonic Compressor Cascade Shape Optimization under Multiple Inlet Mach Operating Conditions

Transonic compressors are widely used today in propulsion and industrial applications thanks to their higher specific work compared to subsonic. In this work, the aerodynamic optimization of a two-dimensional Computational Fluid Dynamics (CFD) model of the transonic cascade ARL-SL19 is described. The validated computational model is used for a multi-objective optimization of the cascade at three different inlet Mach numbers using a genetic algorithm and an artificial neural network, with the aim of reducing total pressure loss and increasing maximum pressure ratio. Finally, the optimized shapes on the Pareto fronts are investigated, analyzing mechanisms responsible for loss reduction and enhanced compression. Profiles having the lowest losses have flatter camberlines and reduced acceleration of flow on the suction side, while geometries achieving the highest pressure ratio values have a more cambered shape with a concave suction side.

Unsteady Numerical Simulation in a Supersonic Compressor Cascade with a Strong Shock Wave

Abstract The most important flow behaviour of supersonic compressor cascades is the shock wave boundary layer interaction (SWBLI). Large eddy simulation (LES) and multiple analysing methods are applied in current study to capture more details of the flow field. It is noted that the LES can catch the dual peaks feature near the SWBLI region with respect to the experimental results. Besides, SWBLI is not only the main losses source in the cascade, but also the most important origin of the unsteadiness behaviour. The high frequency signals correspond to the coherent structure in the boundary layer and dissipate downstream in the cascade, while the low frequency signals relate to the motion of the reflection point of the passage oblique shock wave and dominate the frequency spectrum downstream.

A new method for rapid shock loss evaluation and reduction for the optimization design of a supersonic compressor cascade

A one-dimensional analytical shock loss prediction method was proposed to tailor the shock system, i.e. the strength of the first and second passage shock, and reduce the shock loss in a supersonic cascade. To develop the one-dimensional analytical model, the shock system in a supersonic cascade was divided into four processes which can be seen in most supersonic compressor cascade, i.e. the flow upstream the extending-external shock, the flow between the extending-external shock and the first passage shock, the accelerating flow from the first passage shock to the second passage shock, and the flow downstream of the second passage shock. Based on some flow assumptions and experimental empirical correlations, the complex flows, containing the shock system, in the blade passage of a supersonic compressor cascade could be described with one-dimensional relationships, which can be used to predict shock losses along the flow passage rapidly and determine the shock system improving direction for achieving lower shock loss while keeping the same cascade static pressure ratio. In order to validate the one-dimensional analytical method, the shock system of two supersonic cascades ARL-SL19 and DLR-PAV-1.5 are modified based on the analysis of the model. The modified cascades achieved about 29% and 25% reduction of shock loss at redesign point compared with baseline cascades, respectively.

Kriging-assisted design optimization of S-shape supersonic compressor cascades

In this paper, multi-objective design optimization of a supersonic compressor cascade was dealt with. The geometry considered was an S-shaped 2D cascade developed and tested at DLR (DLR-PAV-1.5), operating under unique incidence and inlet Mach number equal to 1.5 so as to develop approximately a static pressure ratio of 2. In such conditions, the cascade shows a strong lambda shock-boundary layer interaction. Methodology included the use of an in-house Kriging-assisted evolutionary algorithm driving an iterative loop which includes a parametric code for geometry representation and the ANSYS Fluent© solver for flow calculation. Cascade optimization was carried out by varying the shapes of both the suction and the pressure sides while maintaining the unique incidence condition using an iterative CFD-based procedure. The objectives were to minimize the cascade total pressure losses and to maximize the static pressure ratio using a Pareto ranking criterion. Results showed that loss coefficient could be reduced by 25% and the static pressure ratio could be improved by 6.5% as a results of a decrease in the pre-shock Mach number by virtue of an extended pre-compression mechanism that involves the entire suction surface upstream of the passage shock.

Shock loss model and blade profile optimization design of a supersonic cascade

With the improvement of single-stage compressor load, the inflow velocity and relative Mach number at blade tip of compressor are further increased. Using supersonic blade profile to design highly loaded compressor is an effective method to satisfy the requirements of the compressor aerodynamic design. In the internal flow of a blade row, detached shock is caused at the leading edge of supersonic cascade, and complex passage shock structure is produced by shock and boundary layer interaction at blade surface. Establishing shock loss calculation method for detached shock and passage shock and optimizing the blade profile according to the shock loss mechanisms are effective means to design high-efficiency compressors. The earlier studies of shock loss calculation model in supersonic compressor cascade have shown that shock loss of the detached shock at leading edge is related to leading edge radius and inlet Mach number. Therefore, several shock loss models for detached shock are established and proved to be effective. In previous studies on passage shock loss, the changes in geometry of passage shock, which is caused by shock and boundary layer interaction, are not fully considered. So, it is necessary to do further study on passage shock loss model. Based on the internal flow characteristics in supersonic cascade, passage shock loss model is proposed. Physics-based passage shock geometry, which is determined by static pressure rise, Mach number before wave fronts and blade profile curvature, is applied for this passage shock loss model. Near the blade surface, passage shock geometry is transformed into lambda-type shock that is composed of two branch oblique shocks, and oblique shock relationships are used to calculate their total pressure losses. Away from the blade surface, passage shock geometry is very close to the normal shock shape, so that normal shock relationships are used for loss calculation.

A Numerical Study on Passive Control of Shock Wave/Turbulent Boundary Layer in a Supersonic Compressor Cascade

A Numerical Study on Passive Control of Shock Wave/Turbulent Boundary Layer in a Supersonic Compressor Cascade

A Shock Loss Model for Supersonic Compressor Cascades

Loss models used in compression system performance prediction codes are often developed from the study of two-dimensional cascades. In this paper, compressible fluid mechanics has been applied to the changes in shock geometry that are known to occur with back pressure for unstarted operation of supersonic compressor cascades. This physics-based engineering shock loss model is applicable to cascades with arbitrary airfoil shapes. Predictions from the present method have been compared to measurements and Navier–Stokes analyses of the LO30-4 and L030-6 cascades, and very good agreement was demonstrated for unstarted operation. A clear improvement has been demonstrated over previously published shock loss models for unstarted operation, both in the accuracy of the predictions and in the range of applicability. The dramatic increase in overall loss with increasing inlet flow angle is shown to be primarily the result of increased shock loss, and much of this increase is caused by the detached bow shock. For a given Mach number, the viscous profile loss is nearly constant over the entire unstarted operating range of the cascade, unless a shock-induced boundary layer separation occurs near stall. Shock loss is much more sensitive to inlet Mach number than is viscous profile loss.

Compressor Cascade Flow with Strong Shock-Wave/Boundary-Layer Interaction

A two-dimensional multiblock Navier ‐Stokes solver has been successfully applied to a supersonic compressor cascade e ow with strong shock-wave/boundary-layer interaction. The cascade model investigated is typical of a fan blade section with precompression design, low e ow turning, and a static pressure ratio around 2.0. Numerical calculations have been performed on a very e ne grid to guarantee sufe cient resolution of the supersonic/transonic e owe eldwithacomplexwavepatternandshock-inducedboundary-layerseparation.Highaccuracyandrobustness of the numerical scheme is achieved by the second-order Roe upwind ‐total variation diminishing scheme and a one-equation turbulence model. Investigations were performed for inlet Mach numbers from 1.28 to 1.53 and a Reynolds number of 2 :6 ££ 106. The results are compared with those of previous experiments in a supersonic cascade wind tunnel that provide proe le Mach number distributions, wake traverse data, and measured suction surface boundary-layer proe les throughout the strong interaction region. Furthermore, calculated discontinuities of Mach number and e ow direction across the shock system at blade passage entrance are compared with laser anemometer data. Finally the ine uence of inlet Mach number and axial stream tube thickness variation on the shock-wave/boundary-layer interaction mechanism and blade performance is presented.

Studies of Secondary Flow at Endwall of a Supersonic Compressor Cascade

A combined experimental and numerical study was conducted to investigate the endwall secondary flow in a linear supersonic compressor cascade. The experiment was carried out in a supersonic wind tunnel, with Mach and Reynolds numbers (based on the blade chord) of 2.36 and 4.8 X 106, respectively. A three-dimensional viscous computational fluid dynamics (CFD) calculation was performed to simulate the endwall boundary layer and the three-dimensional secondary flows. The experimental and computational analyses presented are all at the design incidence angle. Experimental measurements include multiple-color surface oil flow visualization applied to the endwall and blade surfaces, blade-to-blade pitot pressure contours at the 90% chordwise position, and upstream and downstream static pressures at the endwall. All of these measurements compare favorably with the CFD calculation. Results from the CFD analysis show that endwall total-pressure losses in this supersonic compressor cascade are a significant contributor to the overall losses and can account for one-third of the overall flowfield loss.

Studies of secondary flow at endwall of a supersonic compressor cascade

Studies of secondary flow at endwall of a supersonic compressor cascade

Compressor cascade flow with strong shock-wave/boundary-layer interaction

A two-dimensional multiblock Navier-Stokes solver has been successfully applied to a supersonic compressor cascade flow with strong shock-wave/boundary-layer interaction. The cascade model investigated is typical of a fan blade section with precompression design, low flow turning, and a static pressure ratio around 2.0. Numerical calculations have been performed on a very fine grid to guarantee sufficient resolution of the supersonic/transonic flowfield with a complex wave pattern and shock-induced boundary-layer separation. High accuracy and robustness of the numerical scheme is achieved by the second-order Roe upwind-total variation diminishing scheme and a one-equation turbulence model. Investigations were performed for inlet Mach numbers from 1.28 to 1.53 and a Reynolds number of 2.6 x 106. The results are compared with those of previous experiments in a supersonic cascade wind tunnel that provide profile Mach number distributions, wake traverse data, and measured suction surface boundary-layer profiles throughout the strong interaction region. Furthermore, calculated discontinuities of Mach number and flow direction across the shock system at blade passage entrance are compared with laser anemometer data