MTU Aero Engines AG Compressor Competence Center

  Axial Compressor Copyright: © RWTH Aachen | IST

As part of the "Competence Centre for Compressors", the IST maintains a strategic partnership with the industrial partner MTU Aero Engines for joint research and development of technologies for next-generation aircraft engines. Since 2007, numerous research projects have been carried out at the IST in cooperation with MTU to investigate issues relating to the aerodynamics and aeroelasticity & aeroacoustics of compressors.

 

Technical Introduction

Blade Copyright: © RWTH Aachen | IST

The compressor system is considered the key technology of an aircraft engine. More than any other component, the compressor determines the performance of the engine. The compressor system requires more than half of the total length of the engine and the weight of the compressor system contributes to almost half of the total engine weight. The proportionate manufacturing and maintenance costs are about 30-40 %.

Since the thermal efficiency of the engine increases and the specific fuel consumption decreases with increasing total pressure ratio, this has been increased over the past 50 years by increasing the stage pressure ratios from below 10:1 to over 40:1 for civil engines with high bypass ratios. Due to the maximum possible aerodynamic blade loading, the increase in stage pressure ratios could only be realised to a small extent by higher deflection in the blade grid. The larger part of the increase in the stage pressure ratio results from a higher circumferential speed.

A special characteristic of the compressor is the surge line, which limits the stable operating range of the compressor in the compressor map. It is caused by the aerodynamic load limit of the blading (breakaway limit) and by the detuning of the stages at partial load.

In addition to the surge limit, the efficiency of the compressor is of decisive importance, as this essentially determines the fuel consumption of the engine. In the core flow of a blade grid, the blading is flow-optimised and contributes less than half of the total losses in the compressor. Large losses are caused by gap and secondary flow. Secondary flows increase in magnitude with load, as they are caused by pressure gradients in the blading. The decelerated velocity cannot withstand the counteracting pressure gradient in the boundary layer - caused by the curvature of the core flow. As a result, flow processes occur near the wall that run perpendicular to the main flow direction and lead to local detachments, resulting in increased losses.

Modern simulation techniques for turbomachinery flow allow aerodynamicists to mitigate the causes of secondary flow by optimising the design of the blading and thus reduce secondary losses. Examples of three-dimensional blade design features are circumferential pitch (lean), sweep or contouring.

 

Current Research Trends

The following trends dominate current research at IST.

High Aspect Ratio (HAR)

Increasing the aspect ratio of the blades in a multi-stage compressor creates additional axial space and thus increases design flexibility, for example, to fit more stages into the same axial space.

Most of the literature mentions negative effects of an increased aspect ratio on the aerodynamics and mechanics of the compressor blades. In particular, the reduction of the surge margin and the greater sensitivity to forced response and self-excited vibrations (flutter) are highlighted as motivations for the development of low-aspect-ratio blades. However, these negative effects can be greatly mitigated by three-dimensional blade designs and intentional mistuning in integral blisk design. Regarding the effect on compressor efficiency, the literature gives contradictory results. For example, both reduced, unchanged and increased efficiency is reported for increased aspect ratios.

At the IST, in cooperation with MTU, a modern HAR blading for the high-pressure compressor in blisk design of the rotors on the highly instrumented 2.5-stage axial compressor rig is being investigated. In addition to map measurements to determine the performance and operating behaviour of the compressor, the investigations focus on aeroelastic and aeroacoustic phenomena. Detailed measurements of the laminar-turbulent transition on the blades as well as the turbulent state between the grids also contribute to the validation of the numerical analysis tools.

Secondary flows

In modern axial compressors for aircraft engines, further loss reduction can be achieved primarily in the vicinity of the side walls, where a large part of the losses result from secondary flows. Thus, in the corner area between the side wall and the blade suction side, a region of detached flow can form. Both rotors (hub side) and stators (on the side walls without radial gap) are affected by this three-dimensional flow phenomenon. Depending on the size of the separation area, related to the channel height or passage width, one speaks of corner separation (none) or corner stall (large). Corner separation results in total pressure loss on the one hand and blockage of the flow channel on the other, which can lead to flow collapse if the compressor is increasingly throttled back.

In addition, leakage flow occurs across the radial gap due to the pressure gradient between the suction and discharge sides of the blade. The radial gap flow is strongly oriented in the circumferential direction due to the high circumferential pressure gradient. The different flow direction between the radial gap flow and the main flow results in the leakage flow rolling up to the radial gap vortex. The high total pressure losses are primarily caused by the strong shearing as a result of the different velocities during the suction-side re-entry of the leakage flow.

However, the leakage flow can also be used to weaken the corner separation on the hub side of cantilever stators. In this case, the leakage flow is used to flow against the secondary flow in the wall boundary layer and thus suppresses the accumulation of low-energy fluid in the corner area between the blade suction side and the hub.

In cooperation with MTU, the IST is investigating the boundary zone flow at the hub of cantilever stators on a ring grid wind tunnel with rotating hub. The focus is on the influence of the radial gap on the weakening of the corner separation as well as optimised three-dimensional blade geometries to reduce the losses caused by secondary flow. With this project, the IST is supporting the industrial partner MTU in building up expertise in the design optimisation of the rear high-pressure compressor stages.