Cold and hot air injection upstream of the first rotor tip of a multistage compressor was tested experimentally. The compressor operating range was extended toward lower mass flow by more than 60% indicating a better throttling capability when air injection was activated. A strong dependency of the stability enhancement on the injected mass flow and injection velocity was found. Both increasing injection mass flow rate and increasing injection velocity led to a considerable extension of the throttling line. Comparable enhancements were achieved when reducing the number of nozzles and hence the injection mass flow. It was also found that injection of hot air, at temperatures comparable to air that bled off at a following stage, had no penalty on the stability enhancement. Investigation of the influence of air injection on radial work distribution showed that only small amounts of injected air were sufficient to lead to a significant radial work redistribution. This in turn changed the operating point of the first stage, leading to axial rematching and thus changed the whole operational behavior of the compressor.

This paper majorly aims to identify and understand the driving flow phenomena when the blading aspect ratio of a 1.5-stage axial compressor is increased so that its overall axial length is reduced. The blading is representative for a state-of-the-art high-pressure compressor (HPC) front-stage design. As part of the investigation steady-state RANS simulations are performed to evaluate the impact on its performance and operability. Moreover, an optimized high aspect ratio (HAR) design is introduced to recover performance penalties. In order to achieve the desired reduction in axial stage length at constant blade row spacing and blade height, numerous possible combinations of increased rotor and stator aspect ratios exist. The impact on compressor efficiency and surge margin will be more or less severe, depending on the chord length reduction in rotor and stator. One intermediate combination of both changes in rotor and corresponding stator aspect ratio is analyzed in detail. The results show that by reducing rotor chord length, the compressor’s stability is predominantly compromised, whereas a shorter stator chord has a bigger impact on efficiency than the rotor. For each HAR configuration, profile loss is increased through a reduced blade chord Reynolds number and a higher profile edge thickness-to-chord ratio. Secondary loss is significantly reduced. However, this effect is extenuated by an increased endwall boundary layer thickness-to-chord ratio. Ultimately, this yields a diminished overall stage efficiency. In general, current HPC blade designs exhibit a lower initial rotor aspect ratio compared to the stator vanes. Consequently, an equivalent stage length reduction has a less crucial impact on Reynolds number — hence profile loss — for rotor blades than for stator vanes. Thus, regarding efficiency, there is an optimum of balancing rotor and stator chord length reduction yielding the least efficiency drop. On the contrary, the stability margin for the compressor stage analyzed is primarily driven by the rotor’s clearance-to-chord ratio. Hence, at constant tip clearance an increase in the rotor’s aspect ratio is proportional to the resulting lack of stability. However, specific compressor design modifications are introduced in order to recover the stability margin without adversely affecting design point efficiency, such that the optimized HAR compressor stage exhibits at least the same performance specifications of the baseline design. This study’s findings also encourage that increasing the blading aspect ratio is a feasible measure for reducing the compressor’s overall axial length aiming a compact design. An optimized HAR compressor allows additional design flexibility, which provides potential for performance improvements.

By the use of decomposition for the three-dimensional flow field in compressors into independent through flow and cross flow as first proposed by Chen et al., the tip gap flow structure for different tip gap heights and boundary conditions is examined. Tracing the roll up of a shear layer in planes normal to the blade camber line the position of the tip leakage vortex is presented in a non-dimensional formulation. Tests for different boundary conditions, i.e. the distinction between a stationary and a rotating end-wall as well as the use of different fluid models are made to quantify their influence on the tip leakage vortex position. By comparing the analytical result to data extracted from three-dimensional RANS computations and to measurement data the validity of the model is presented. Finally, an attempt is made to find a criterion for the occurrence of tip leakage vortex induced stall of tip critical rotor blades. This is done by a correlation based on the previously derived non-dimensional vortex trajectory and the stagger angle. This criterion is again tested on results of a three dimensional RANS computation proving its validity.

Fluid injection at the tip of highly loaded compressor rotors is known to be effective in suppressing the onset of rotating stall and eventually compressor instability. However, using such stability enhancement methods in a multistage compressor might not only stabilize certain stages but has also an impact on radial and axial matching. In order to account for tip injection during the early stages of compressor design, this paper focuses on the development of a method to model the physical effects underlying tip injection within a streamline curvature method. With the help of system identification it could be shown that a rotor subject to the discrete jets of tip injection adapts to the varying flow conditions according to a first order model. This information was used to generate a time-dependent input for the steady equations used with a streamline curvature method and eventually to model the unsteady response of the rotor to tip injection. Comparing the results obtained with the enhanced streamline curvature model to measurement results, good agreement could be shown which raised confidence that the influence of tip injection on axial and radial matching was sufficiently captured.

Fluid injection at the tip of highly loaded compressor rotors is known to be very effective in suppressing the onset of rotating stall and eventually compressor instability. To understand the effects of tip injection, the flow field at the tip region of a transonic compressor rotor with and without fluid injection was investigated in this paper. Using results acquired by phase-locked PIV measurements as well as the static pressure field obtained by fast response pressure transducers, the unsteady interaction between the injection jet and the rotor could be described thoroughly. Both, an influence of the rotor’s flow field on the jet as well of the jet on the rotor was clearly visible. Since unsteady inflow conditions to the front rotor in the relative frame of reference were imposed by the injection jets, the rotor’s unsteady response was investigated by inspection of the position of the tip leakage vortex trajectory. It could be shown that due to a short time for the flow to adapt at the rotor’s leading edge, its position didn’t change distinctly. Because a significantly longer time was needed for the overall passage flow to adapt, it was concluded that this causes the beneficial effect of tip injection.