Blade Row Performance Research Articles

Modeling of a turbine bladerow with stagger angle variation using the multi-fidelity influence superposition method

Manufacturing variabilities can significantly affect a turbine bladerow's performance. With the push for turbine's efficiency and reliability more than ever, the understanding of manufacturing variability impacts is sought after. It has been shown that a multi-passage simulation domain is required to capture the interaction among varied blades. This is a challenge for conventional single-passage methods, largely due to the requirement of prohibitive computational resources. The present work introduces an attempt to solve this challenge by combining the multi-fidelity method and the influence superposition approach. The proposed methodology combines the accuracy of a high-fidelity simulation and the speed of a low-fidelity simulation. Therefore, the multi-fidelity predictions tend to be both accurate and fast. The key enabler is that only a small set of configurations is needed to pre-compute the source term. Two representative geometries of a subsonic low-pressure turbine and a transonic high-pressure turbine have been chosen as the test cases. Each test case is subject to two further stagger angle variation patterns, namely alternating and sinusoidal pattern. The low-fidelity method has been shown to be inadequate for the transonic high-pressure turbine test case, owning to its incapability to capture correctly the shockwave formation and the shockwave/wake interaction. On the other hand, the proposed multi-fidelity method has been successful to match qualitatively and quantitatively compared to the direct high-fidelity solution. More interestingly, the multi-fidelity method has a reduced computational overhead of one order of magnitude compared to the direct high-fidelity simulation. With an ability to accurately and efficiently predict the manufacturing variability effects, this method provides a tool for engineers to explore and optimize their blading designs.

Effects of Stator Platform Geometry Features on Blade Row Performance

This paper details an experimental investigation, using a linear cascade, into the effects of real geometry features on the aerodynamic performance of stator blade rows within axial flow compressors. The specific geometric features investigated include shroud cavities, inter-platform gaps, vane-pack gaps and the effects of misalignment of the platform endwalls due to manufacturing tolerances. A computational investigation into these effects is also included. To ensure that the linear cascade measurements are representative of a multi-stage compressor environment a novel experimental technique was developed to generate a hub endwall boundary layer which had skew. The boundary layer skew generation method involves injecting flow along the cascade endwall in such a manner as to control both the displacement thickness and tangential momentum thickness of the resulting boundary layer. Without the presence of the endwall boundary layer skew the linear cascade could not reproduce the flow features typically observed in a multi-stage compressor. The investigation reveals that real geometry features can have a significant impact on the flowfield within a blade passage. For a shrouded stator, increasing the leakage flow rate increases the stagnation pressure loss coefficient. However, high levels of whirl pickup of the leakage flow as it passes through the stator-shroud cavity can offset the natural secondary flow within the stator passage and thus reduce the stagnation pressure loss. All of the steps and gaps that were observed to be present in real compressors were found to increase the stagnation pressure loss relative to that of a smooth endwall. It is also shown that the computational method is capable of capturing the trends observed in the experiments.

Development of a coupled supersonic inlet-fan Navier–Stokes simulation method

A coupled supersonic inlet-fan Navier–Stokes simulation method was developed by using COMSOL-CFD code. The flow turning, pressure rise and loss effects across blade rows of the fan and the inlet-fan interactions were taken into account as source terms of the governing equations without a blade geometry by a body force model. In this model, viscous effects in blade passages can also be calculated directly, which include the exchange of momentum between fluids and detailed viscous flow close to walls. NASA Rotor 37 compressor test rig was used to validate the ability of the body force model to estimate the real performance of blade rows. Calculated pressure ratio characteristics and the distribution of the total pressure, total temperature, and swirl angle in the span direction agreed well with experimental and numerical data. It is shown that the body force model is a promising approach for predicting the flow field of the turbomachinery. Then, coupled axisymmetric mixed compression supersonic inlet-fan simulations were conducted at Mach number 2.8 operating conditions. The analysis includes coupled steady-state performance, and effects of the fan on the inlet. The results indicate that the coupled simulation method is capable of simulating behavior of the supersonic inlet-fan system.

Modeling Particle Deposition Effects in Aircraft Engine Compressors

Airborne particles ingested in aircraft engines deposit on compressor blading and end walls. Aerodynamic surfaces degrade on a microscopic and macroscopic scale. Blade row, compressor, and engine performance deteriorate. Optimization of maintenance scheduling to mitigate these effects requires modeling of the deterioration process. This work provides a deterioration model on blade row level and the experimental validation of this model in a newly designed deposition test rig. When reviewing previously published work, a clear focus on deposition effects in industrial gas turbines becomes evident. The present work focuses on quantifying magnitudes and timescales of deposition effects in aircraft engines and the adaptation of the generalized Kern and Seaton deposition model for application in axial compressor blade rows. The test rig's cascade was designed to be representative of aircraft engine compressor blading. The cascade was exposed to an accelerated deposition process. Reproducible deposition patterns were identified. Results showed an asymptotic progression of blade row performance deterioration. A significant increase in total pressure loss and decrease in static pressure rise were measured. Application of the validated model using existing particle concentration and flight cycle data showed that more than 95% of the performance deterioration due to deposition occurs within the first 1000 flight cycles.

Two-Dimensional Moving Blade Row Interactions in a Stratospheric Airship Contra-Rotating Open Propeller Configuration

The numerical simulation of two-dimensional moving blade row interactions is conducted by CFD means to investigate the interactions between the front and rear propeller in a stratospheric airship contra-rotating open propeller configuration caused by different rotational speeds. The rotational speed is a main factor to affect the propeller Reynolds number which impact the aerodynamic performance of blade rows significantly. This effect works until the Reynolds number reaches a high enough value beyond which the coefficients become independent. Additionally, the interference on the blade row has been revealed by the investigation. The front blade row moves in the induced-velocity field generated by the rear blade row and the aerodynamic coefficients are influenced when the rear blade row has fast RPMs. The rear blade row moving behind the front one is affected directly by the wake and eddies generated by the front blade row. The aerodynamic coefficients reduce when the front blade row has slow RPMs while increase when the front blade row moves faster than itself. But overall, the interference on the front blade row due to the rear blade row is slight and the interference on the rear blade row due to the front blade row is much more significant.

Effect of Casing Suction on Stage Matching and Flow Loss in a High Pressure Axial Compressor

A numerical simulation of a 1.5 stages high pressure axial compressor is presented. The effect of casing suction on compressor aerodynamic performance, stage matching as well as flow loss is discussed in detail. Using modern CFD codes Numeca/FINE, the 3D steady flow inside the 1.5 stages axial compressor has been simulated and visualized, respectively at cruise speed and maximum climbing speed. Numerical solutions show that casing suction partly removes the low-momentum fluid and reduces the flow blockage at tip region, thus improves stage matching mildly between rotor and stator. In general, casing suction, which provides main flow for flight air system, has slight impact on compressor blade row performance and loss characteristics.

Averaging Nonuniform Flow for a Purpose

Averaging nonuniform flow is important for the analysis of measurements in turbomachinery and gas turbines; more recently an important need for averaging arises with results of computational fluid dynamics (CFD). In this paper we show that there is a method for averaging which is “correct,” in the sense of preserving the essential features of the nonuniform flow, but that the type of averaging which is appropriate depends on the application considered. The crucial feature is the decision to retain or conserve those quantities which are most important in the case considered. Examples are given to demonstrate the appropriate methods to average nonuniform flows in the accounting for turbomachinery blade row performance, production of thrust in a nozzle, and mass flow capacity in a choked turbine. It is also shown that the numerical differences for different types of averaging are, in many cases, remarkably small.

Prediction of Main/Secondary-Air System Flow Interaction in a High-Pressure Turbine

The results from a steady, viscous flow simulation of the main flowpath of a modern jet engine transonic high-pressure turbine, transition duct, and first vane of the low-pressure turbine, coupled with the secondary-air system under the high-pressure turbine, is presented. The secondary-air system configuration includes the lower cavity near the centerline, the mid-cavity with the tangential onboard injection nozzle and pump, and the upper rim cavity. In addition, the flow through the inner tangential onboard injection seal and labyrinth seals that separate the cavities is included. The purpose of this investigation is to determine the feasibility of the use of computational procedures to predict the steady-flow interaction effects between the main- and secondary-air system flow, the ability to predict flow splits in the secondary-air system, and the numerical issues associated with such a computation. A description of the numerical procedure, along with technical details of the initial and boundary conditions and convergence of the numerical procedure are presented. Details of the predicted flow physics in the secondary-air system flowpath, including the rim-cavity and inner tangential onboard injection/labyrinth seals, the main flowpath, and the interaction region at the junction between the two flowpaths are presented. The effect of the secondary-air system flow on the aerodynamic performance and hub endwall temperature distribution of the main flowpath blade rows is also discussed.

Development of Hub Corner Stall and Its Influence on the Performance of Axial Compressor Blade Rows

A detailed investigation has been performed to study hub corner stall phenomena in compressor blade rows. Three-dimensional flows in a subsonic annular compressor stator and in a transonic compressor rotor have been analyzed numerically by solving the Reynolds-averaged Navier–Stokes equations. The numerical results and the existing experimental data are interrogated to understand the mechanism of compressor hub corner stall. Both the measurements and the numerical solutions for the stator indicate that a strong twisterlike vortex is formed near the rear part of the blade suction surface. Low-momentum fluid inside the hub boundary layer is transported toward the suction side of the blade by this vortex. On the blade suction surface near the hub, this vortex forces fluid to move against the main flow direction and a limiting stream surface is formed near the hub. The formation of this vortex is the main mechanism of hub corner stall. When the aerodynamic loading is increased, the vortex initiates further upstream, which results in a larger corner stall region. For the transonic compressor rotor studied in this paper, the numerical solution indicates that a mild hub corner stall exists at 100 percent rotor speed. The hub corner stall, however, disappears at the reduced blade loading, which occurs at 60 percent rotor design speed. The present study demonstrates that hub corner stall is caused by a three-dimensional vortex system and that it does not seem to be correlated with a simple diffusion factor for the blade row.

Wake-Induced Unsteady Flows: Their Impact on Rotor Performance and Wake Rectification

The impact of wake-induced unsteady flows on blade row performance and the wake rectification process is examined by means of numerical simulation. The passage of a stator wake through a downstream rotor is first simulated using a three-dimensional unsteady viscous flow code. The results from this simulation are used to define two steady-state inlet conditions for a three-dimensional viscous flow simulation of a rotor operating in isolation. The results obtained from these numerical simulations are then compared to those obtained from the unsteady simulation both to quantify the impact of the wake-induced unsteady flow field on rotor performance and to identify the flow processes which impact wake rectification. Finally, the results from this comparison study are related to an existing model, which attempts to account for the impact of wake-induced unsteady flows on the performance of multistage turbomachinery.

Cooling-Air Injection Into Secondary Flow and Loss Fields Within a Linear Turbine Cascade

In order to understand overall performance and internal flows of air-cooled turbine blade rows, flows in a model linear cascade were surveyed with secondary air injection from various locations of the blade surfaces. The secondary air interacted with the cascade passage vortices and changed the loss distribution significantly. The cascade overall loss decreased when the air was injected along the mainstream and increased when the air was injected against the mainstream from some locations of the blade leading edge. Effects on overall kinetic energy of the secondary flows and on the cascade outlet flow angle were also discussed in this paper.

Nonaxisymmetric Flow through Annular Actuator Disks: Inlet Distortion Problem

The passage of distorted flow through an annular axial compressor rotor or stator is analyzed in the actuator disk limit. In such a description the flow is steady in absolute coordinates. The resulting analysis yields an overall description of the blade row performance in the presence of inlet flow defects. The present (actuator-disk) analysis is compared successfully with Dunham’s earlier three-dimensional analysis as well as with recent experimental data. In addition, comparison with a recent theory by Greitzer, employing an alternative approach, is made. The analysis shows that at least two important distinct types of vorticity arise, the one being directly analogous to the trailing vorticity (shed circulation) of classical wing theory, and the second being analogous to the vorticity occurring in “secondary flow” theory. The latter arises directly as a result of inlet distortion. The resulting varying flow angles produce spanwise variation in blade loading, and consequent trailing vorticity (Beltrami Flow). The two types of vorticity are therefore interrelated. The static pressure field is also affected by this coupling in agreement with experiment. This problem provided a striking example for which three-dimensional and two-dimensional analyses disagree qualitatively. The vorticity as viewed in coordinates fixed in the rotor is discussed in the Addendum.

Performance characteristics of a model VTOL lift fan in crossflow.

This paper presents a summary of principal results obtained from crossflow tests of a model 15-in.-diam lift fan installed in a wing in the NASA Lewis Research Center, 9 by 15 ft V/STOL Propulsion Wind Tunnel. Tests were run with and without exit louvers over a range of tunnel air speeds, fan speeds, and wing angle of attack. Fan thrust in crossflow was influenced by two principal factors: the effects of inflow distortion on blade-row performance, and changes in fan stage operating point brought about by changes in back pressure ratio. In this particular fan, flow separation on the inlet bellmouth did not appear to be a serious problem for crossflow operation.

Closure to “Discussion of ‘Performance of Compressor Blade Rows in a Sloping Flowpath’” (1972, ASME J. Eng. Power, 94, p. 68)

Closure to “Discussion of ‘Performance of Compressor Blade Rows in a Sloping Flowpath’” (1972, ASME J. Eng. Power, 94, p. 68)

Performance of Compressor Blade Rows in a Sloping Flowpath

In an effort to reduce engine length and weight it is desirable in some applications, such as high bypass ratio turbofan engines, to introduce compressor stages into a downward sloping flowpath. In this “mixed flow” compressor the airfoil orientation with respect to the flowpath becomes important to ensure good performance. The performance of cascades and compressor stages with sloping walls is presented. Three methods of introducing blading into the flowpath were evaluated. It is shown that by canting the blading to be normal to the flow direction, no penalty in performance is incurred compared to the performance of a typical axial compressor with cylindrical flowpath. Test results are compared with predictions based on intrablade analysis and a wall boundary layer calculation. Qualitatively good agreement is obtained except for the case when the blading is swept with respect to the flow direction. Need for further investigtion of three dimensional internal viscous flows is indicated.

Stage Performance and Radial Matching of Axial Compressor Blade Rows

To obtain the ultimate in performance from an axial compressor stage it is necessary that the operation of the blade rows comprising the stage be thoroughly understood. Appreciable advances have been accomplished in recent years, but much still remains to be done. This paper aims to present practical hypotheses concerning the nature of the flow between the blades and deals with design techniques in relation to this flow. In addition, some elements of two-dimensional cascade data and a discourse on practical secondary flows are presented.