Measurements of Secondary Flows Downstream of a Turbine Cascade at Off-Design Incidence

This paper presents experimental results of the secondary flows from two large-scale, low-speed, linear turbine cascades for which the incidence was varied. The aerofoils for the two cascades were designed for the same inlet and outlet conditions and differed mainly in their leading-edge geometries. Detailed flow field measurements were made upstream and downstream of the cascades and static pressure distributions were measured on the blade surfaces for three different values of incidence: 0, +10 and +20 degrees. The results from this experiment indicate that the strength of the passage vortex does not continue to increase with incidence, as would be expected from inviscid flow theory. The streamwise acceleration within the aerofoil passage seems to play an important role in influencing the strength of the vortex. The most recent off-design secondary loss correlation (Moustapha et al. [1]) includes leading-edge diameter as an influential correlating parameter. The correlation predicts that the secondary losses for the aerofoil with the larger leading-edge diameter are lower at off-design incidence; however, the opposite is observed experimentally. The loss results at high positive incidence have also high-lighted some serious shortcomings with the conventional method of loss decomposition. An empirical prediction method for secondary losses has been developed and will be presented in a subsequent paper.

Experimental and Numerical Investigation on the Performance of a Family of Three HP Transonic Turbine Blades

Experimental results concerning the performance of three high-pressure (HP) transonic turbine blades having fore-, aft- and mid-loadings have been presented previously by Corriveau and Sjolander [1]. Results from that study indicated that by shifting the loading towards the rear of the airfoil, improvements in loss performance of the order of 20% could be obtained near the design Mach number. In order to gain a better understanding of the underlying reasons for the improved loss performance of the aft-loaded blade, additional measurements were performed on the three cascades. Furthermore, 2-D numerical simulations of the cascade flow were performed in order to help in the interpretation of the experimental results. Based on the analysis of additional wake traverse data and base pressure measurements combined with the numerical results, it was found that the poorer loss performance of the baseline mid-loaded profile compared to the aft-loaded blade could be traced back to the former’s higher rear suction side curvature. The presence of higher rear suction surface curvature resulted in higher flow velocity in that region. Higher flow velocity at the trailing edge in turn contributed to reducing the base pressure. The lower base pressure at the trailing edge resulted in a stronger trailing edge shock system for the mid-loaded blade. This shock system increased the losses for the mid-loaded baseline profile when compared to the aft-loaded profile.

Investigation on the Influence of a Core Chevron Nozzle on the Performance of a Modern Bypass Engine

In light of intensifying environmental concerns, the noise in aircraft gas turbine engines needs to be reduced significantly. Considerable work has been conducted to reduce jet noise produced by the mixing of high velocity gas streams with ambient air. Various nozzle designs such as lobed nozzles, serrated nozzles or chevron nozzles have been used and proposed to control and modify the velocity pattern of exhaust gas streams. This paper presents investigations on the influence of a core chevron nozzle on the performance of a modern bypass engine. The characteristic discharge, velocity and specific thrust coefficients of the chevron and non-chevron nozzles are determined by numerical calculations and are verified with experimental data. The nozzle coefficients form the basis for an engine performance comparison between the two hot nozzle configurations of the bypass engine. The effect of the nozzle configuration on overall engine performance and component working points has been investigated by applying an engine performance synthesis tool. The thrust loss and the corresponding SFC increase which has been observed by using the chevron nozzle have been related to engine internal rematching and changes in nozzle performance.

An Approach to Integrated Multi-Disciplinary Turbomachinery Design

Aeroengines are designed using fractured processes. Complexity has driven the design of such machines to be subdivided by specialism, customer and function. While this approach has worked well in the past, with component efficiencies, current material performance and the possibilities presented by scaling existing designs for future needs becoming progressively exhausted it is necessary to reverse this process of disintegration. Our research addresses this aim. The strategy we use has two symbiotic arms. The first is an open data architecture from which existing disparate design codes all derive their input and to which all send their output. The second is a dynamic design process management system known as “SignPosting”. Both the design codes and parameters are arranged into complementary multiple level hierarchies: fundamental to the successful implementation of our strategy is the robustness of the mechanisms we have developed to ensure consistency in this environment as the design develops over time. One of the key benefits of adopting a hierarchical structure is that it confers not only the ability to use mean-line, throughflow and fully 3D CFD techniques in the same environment but also to cross specialism boundaries and insert mechanical, material, thermal, electrical and structural codes, enabling exploration of the design space for multi-disciplinary non-linear responses to design changes and their exploitation. We present results from trials of an early version of the system applied to the redesign of a generic civil aeroengine core compressor. SignPosting has allowed us to examine the hardness of design constraints across disciplines which has shown that it is far more profitable not to strive for even higher aerodynamic performance, but rather improve the commercial performance by maintaining design and part speed pressure ratios stability and efficiency while increasing rotor blade creep life by up to 70%.

Design of a Transonic High-Pressure Turbine Stage 2D Section With Reduced Rotor/Stator Interaction

In transonic turbine stages, the exit static pressure field of the vane is highly non-uniform in the pitchwise direction. The rotor traverses periodically this non-uniform field and large static pressure fluctuations are observed around the rotor section. As a consequence the rotor blade is submitted to significant variations of its aerodynamic force. This contributes to the high cycle fatigue and may result in unexpected blade failure. In this paper an existing transonic turbine stage section is redesigned in the view of reducing the rotor stator interaction, and in particular the unsteady rotor blade forcing. The first step is the redesign of the stator blade profile to reduce the stator exit pitchwise static pressure gradient. For this purpose, a procedure using a genetic algorithm and an artificial neural network is used. Next, two new rotor profiles are designed and analysed with a quasi 3D Euler unsteady solver in order to investigate their receptivity to the shock interaction. One of the new profiles allows reducing the blade force variation by 50%.

Flow With Shock Waves in a Transonic Centrifugal Compressor With a Low-Solidity Cascade Diffuser Using PIV

A single-stage transonic centrifugal compressor with a pressure ratio greater than six was tested in a closed loop with HFC134a gas. Flow at the inducer of a rotating impeller as well as flow in a stationary low-solidity cascade diffuser was measured using a double-pulse and double-frame particle image velocimetry (PIV). Shock waves in both flows were clearly observed. The effect of flow rate on a 3D configuration of shock wave at the inducer and a so-called rotor-stator interaction between a rotating impeller and a stationary cascade were discussed based on a phase-averaged measurement technique. Furthermore, the unsteadiness of inducer shock wave and the flow in a cascade diffuser during surge were discussed based on instantaneous velocity vector maps.

Investigation on a Type of Flow Control to Weaken Unsteady Separated Flows by Unsteady Excitation in Axial Flow Compressors

By solving unsteady Reynolds-averaged 2-D N-S equations discretized by a high-order scheme, the results showed that the disordered unsteady separated flow could be effectively controlled by periodic suction and blowing in a wide range of incidence, resulting in enhancement of time-averaged aerodynamic performances. The effects of unsteady excitation frequency, amplitude and excitation location were investigated in detail. The effective excitation frequency spans a wide spectrum and there is an optimal excitation frequency that is nearly equal to the Characteristic frequency of vortex shedding. Excitation amplitude exhibits a threshold value (nearly 10% in term of the ratio of maximum velocity of periodic suction and blowing to the velocity of free flow) and an optimal value (nearly 35%). The optimal excitation location is just upstream of the separation point. We also explored feasible unsteady actuators by utilizing upstream wake for constraining unsteady separation in axial flow compressors.

Noise Reduction of Blade-Passing Frequency Components in a Centrifugal Blower

A hybrid-type noise control method is applied to fundamental and higher-order blade-passing frequency components, abbreviated to BPF components, radiated from a centrifugal blower. An active cancellation of the BPF noise source is conducted based on a detailed investigation of the noise source distribution by using correlation analysis. The sound pressure level of 2nd- and/or 3rd-order BPF can be reduced by more than 15 decibels and discrete tones almost eliminate from the power spectra of blower-radiated noise. On the other hand, the sound pressure level of the fundamental BPF is difficult to reduce effectively by the active cancellation method because of the large amplitude of the noise source fluctuation. However, the fundamental BPF is largely influenced by the frequency-response characteristics of the noise transmission passage, and is passively reduced by appropriate adjusting of the inlet duct length. Simultaneous reduction of BPF noise, therefore, can be easily made possible by applying passive and active control methods on the fundamental and higher-order BPF noise, respectively. We also discuss the distribution pattern of BPF noise sources by numerical simulation of flow fields around the scroll cutoff.

Noise Computation of a Turbo-Engine Jet Exhaust Based on LES and Lighthill’s Acoustic Analogy

Increasing noise regulations at urban airports force jet engine manufactures to develop and build more quiet engines. Over recent years, a significant reduction in fan and mechanical noise has been achieved. However, the jet exhaust is the principal source of noise. The acoustical field that is generated by a turbo-engine jet exhaust running near the ground level is considered. The full equations of motion for compressible and unsteady flows describe both flow field and sound generation. The flow variables are decomposed into semi-compressible components and inviscid, irrotational acoustical components. The turbulent flow and mixing are computed using Large Eddy Simulation (LES). The radiated acoustical field is computed using the Lighthill’s acoustic analogy with acoustic sources provided by instantaneous LES data.

Axial Turbine Tip Desensitization by Injection From a Tip Trench: Part 2 — Leakage Flow Sensitivity to Injection Location

In Part 1 of this paper it was shown that discrete jets issuing from a tip platform trench were successful in reducing the total pressure deficit due to tip leakage flow. The specific tip cooling system used in Part 1 had all four injection locations active. This paper examines the effect of the individual location of the injection hole on the tip leakage flow. The investigation was carried out in a large-scale rotating rig. Total pressure downstream of the rotor exit was measured using a Kulite sensor. The measurements were phase-locked and ensemble averaged over 200 rotor revolutions. The injection holes are located at 61%, 71%, 81%, and 91% blade axial chord, in the tip trench of a single blade with a clearance of 1.40% blade height. Individual injection at 61% and 71% chord reduced the leakage vortex size. Coolant injection at 81% chord was the most successful in reducing the total pressure deficit in the leakage vortex. Injection from 91% chord had no effect on the leakage vortex. Injection from combinations of holes had greater effect in reducing the leakage vortex size and the total pressure deficit associated with the vortex. It can be concluded that the individual jets most likely turn the leakage flow towards the trailing edge. Most of the leakage flow that is responsible for the greatest total pressure deficit occurs around 80% chord.