Journal of Turbomachinery

A computational study to define the phenomena that lead to the onset of short length-scale (spike) rotating stall disturbances has been carried out. Based on unsteady simulations, we hypothesize there are two conditions necessary for the formation of spike disturbances, both of which are linked to the tip clearance flow. One is that the interface between the tip clearance and oncoming flows becomes parallel to the leading-edge plane. The second is the initiation of backflow, stemming from the fluid in adjacent passages, at the trailing-edge plane. The two criteria also imply a circumferential length scale for spike disturbances. The hypothesis and scenario developed are consistent with numerical simulations and experimental observations of axial compressor stall inception. A comparison of calculations for multiple blades with those for single passages also allows statements to be made about the utility of single passage computations as a descriptor of compressor stall.

Work on rotating stall and its related disturbances have been in progress since the Second World War. During this period, certain “hot topics” have come to the fore—mostly in response to pressing problems associated with new engine designs. This paper will take a semihistorical look at some of these fields of study (stall, surge, active control, rotating instabilities, etc.) and will examine the ideas which underpin each topic. Good progress can be reported, but the paper will not be an unrestricted celebration of our successes because, after 75 years of research, we are still unable to predict the stalling behavior of a new compressor or to contribute much to the design of a more stall-resistant machine. Looking forward from where we are today, it is clear that future developments will come from CFD in the form of better performance predictions, better flow modeling, and improved interpretation of experimental results. It is also clear that future experimental work will be most effective when focussed on real compressors with real problems—such as stage matching, large tip clearances, eccentricity, and service life degradation. Today’s topics of interest are mostly associated with compressible effects and so further research will require more high-speed testing.

Experimental investigations have shown that the enhancement in heat transfer coefficients for air flow in a channel roughened with angled ribs is on the average higher than that roughened with 90 deg ribs of the same geometry. Secondary flows generated by the angled ribs are believed to be responsible for these higher heat transfer coefficients. These secondary flows also create a spanwise variation in heat transfer coefficient on the roughened wall with high levels of heat transfer coefficient at one end of the rib and low levels at the other end. In an effort basically to double the area of high heat transfer coefficients, the angled rib is broken at the center to form a V-shaped rib, and tests are conducted to investigate the resulting heat transfer coefficients and friction factors. Three different square rib geometries, corresponding to blockage ratios of 0.083, 0.125, and 0.167, with a fixed pitch-to-height ratio of 10, mounted on two opposite walls of a square channel in a staggered configuration, are tested in a stationary channel for 5000 < Re < 30,000. Heat transfer coefficients, friction factors, and thermal performances are compared with those of 90 deg, 45 deg, and discrete angled ribs. The V-shaped ribs are tested for both pointing upstream and downstream of the main flow. Test results show that: (a) 90 deg ribs represent the lowest thermal performance, based on the same pumping power, and is essentially the same for the 2:1 change in blockage ratio, (b) low-blockage-ratio (e/Dh = 0.083) V-shaped ribs pointing downstream produced the highest heat transfer enhancement and friction factors. Among all other geometries with blockage ratios of 0.125 and 0.167, 45 deg ribs showed the highest heat transfer enhancements with friction factors less than those of V-shaped ribs, (c) thermal performance of 45 deg ribs and the lowest blockage discrete ribs are among the highest of the geometries tested in this investigation, and (d) discrete angled ribs, although inferior to 45 deg and V-shaped ribs, produce much higher heat transfer coefficients and lower friction factors compared to 90 deg ribs.

Film cooling effectiveness was studied experimentally in a flat plate test facility with zero pressure gradient using a single row of inclined holes, which injected high-density, cryogenically cooled air. Round holes and holes with a diffusing expanded exit were directed laterally away from the free-stream direction with a compound angle of 60 deg. Comparisons were made with a baseline case of round holes aligned with the free stream. The effects of doubling the hole spacing to six hole diameters for each geometry were also examined. Experiments were performed at a density ratio of 1.6 with a range of blowing ratios from 0.5 to 2.5 and momentum flux ratios from 0.16 to 3.9. Lateral distributions of adiabatic effectiveness results were determined at streamwise distances from 3 D to 15 D downstream of the injection holes. All hole geometries had similar maximum spatially averaged effectiveness at a low momentum flux ratio of I = 0.25, but the round and expanded exit holes with compound angle had significantly greater effectiveness at larger momentum flux ratios. The compound angle holes with expanded exits had a much improved lateral distribution of coolant near the hole for all momentum flux ratios.

Experimental measurements in a linear cascade with tip clearance are complemented by numerical solutions of the three-dimensional Navier–Stokes equations in an investigation of tip leakage flow. Measurements reveal that the clearance flow, which separates near the entry of the tip gap, remains unattached for the majority of the blade chord when the tip clearance is similar to that typical of a machine. The numerical predictions of leakage flow rate agree very well with measurements, and detailed comparisons show that the mechanism of tip leakage is primarily inviscid. It is demonstrated by simple calculation that it is the static pressure field near the end of the blade that controls chordwise distribution of the flow across the tip. Although the presence of a vortex caused by the roll-up of the leakage flow may affect the local pressure field, the overall magnitude of the tip leakage flow remains strongly related to the aerodynamic loading of the blades.

The breakdown of tip leakage vortex has been investigated on a low-speed axial compressor rotor with moderate blade loading. Effects of the breakdown on the rotor aerodynamics are elucidated by Navier–Stokes flow simulations and visualization techniques for identifying the breakdown. The simulations show that the leakage vortex breakdown occurs inside the rotor at a lower flow rate than the peak pressure rise operating condition. The breakdown is characterized by the existence of the stagnation point followed by a bubblelike recirculation region. The onset of breakdown causes significant changes in the nature of the tip leakage vortex: large expansion of the vortex and disappearance of the streamwise vorticity concentrated in the vortex. The expansion has an extremely large blockage effect extending upstream of the leading edge. The disappearance of the concentrated vorticity results in no rolling-up of the vortex downstream of the rotor and the disappearance of the pressure trough on the casing. The leakage flow field downstream of the rotor is dominated by the outward radial flow, resulting from the contraction of the bubblelike structure of the breakdown region. It is found that the leakage vortex breakdown plays a major role in characteristic of rotor performance at near-stall conditions. As the flow rate is decreased from the peak pressure rise operating condition, the breakdown region grows rapidly in the streamwise, spanwise, and pitchwise directions. The growth of the breakdown causes the blockage and the loss to increase drastically. Then, the interaction of the breakdown region with the blade suction surface gives rise to the three-dimensional separation of the suction surface boundary layer, thus leading to a sudden drop in the total pressure rise across the rotor.

This paper presents detailed measurements of the film-cooling effectiveness for three single, scaled-up film-cooling hole geometries. The hole geometries investigated include a cylindrical hole and two holes with a diffuser-shaped exit portion (i.e., a fan-shaped and a laid-back fan-shaped hole). The flow conditions considered are the crossflow Mach number at the hole entrance side (up to 0.6), the crossflow Mach number at the hole exit side (up to 1.2), and the blowing ratio (up to 2). The coolant-to-mainflow temperature ratio is kept constant at 0.54. The measurements are performed by means of an infrared camera system, which provides a two-dimensional distribution of the film-cooling effectiveness in the near field of the cooling hole down to x/D = 10. As compared to the cylindrical hole, both expanded holes show significantly improved thermal protection of the surface downstream of the ejection location, particularly at high blowing ratios. The laidback fan-shaped hole provides a better lateral spreading of the ejected coolant than the fan-shaped hole, which leads to higher laterally averaged film-cooling effectiveness. Coolant passage cross-flow Mach number and orientation strongly affect the flowfield of the jet being ejected from the hole and, therefore, have an important impact on film-cooling performance.

In this paper, we describe the structures that produce a spike-type route to rotating stall and explain the physical mechanism for their formation. The descriptions and explanations are based on numerical simulations, complemented and corroborated by experiments. It is found that spikes are caused by a separation at the leading edge due to high incidence. The separation gives rise to shedding of vorticity from the leading edge and the consequent formation of vortices that span between the suction surface and the casing. As seen in the rotor frame of reference, near the casing the vortex convects toward the pressure surface of the adjacent blade. The approach of the vortex to the adjacent blade triggers a separation on that blade so the structure propagates. The above sequence of events constitutes a spike. The computed structure of the spike is shown to be consistent with rotor leading edge pressure measurements from the casing of several compressors: the centre of the vortex is responsible for a pressure drop and the partially blocked passages associated with leading edge separations produce a pressure rise. The simulations show leading edge separation and shed vortices over a range of tip clearances including zero. The implication, in accord with recent experimental findings, is that they are not part of the tip clearance vortex. Although the computations always show high incidence to be the cause of the spike, the conditions that give rise to this incidence (e.g., blockage from a corner separation or the tip leakage jet from the adjacent blade) do depend on the details of the compressor.

Flow separations in the corner regions of blade passages are common. The separations are three dimensional and have quite different properties from the two-dimensional separations that are considered in elementary courses of fluid mechanics. In particular, the consequences for the flow may be less severe than the two-dimensional separation. This paper describes the nature of three-dimensional (3D) separation and addresses the way in which topological rules, based on a linear treatment of the Navier-Stokes equations, can predict properties of the limiting streamlines, including the singularities which form. The paper shows measurements of the flow field in a linear cascade of compressor blades and compares these to the results of 3D computational fluid dynamics (CFD). For corners without tip clearance, the presence of three-dimensional separation appears to be universal, and the challenge for the designer is to limit the loss and blockage produced. The CFD appears capable of predicting this.

The influence of three-dimensional flow structures within a compressor blade passage has been examined computationally to determine their role in rotating stall inception. The computations displayed a short length-scale (or spike) type of stall inception similar to that seen in experiments; to the authors’ knowledge this is the first time such a feature has been simulated. A central feature observed during the rotating stall inception was the tip clearance vortex moving forward of the blade row leading edge. Vortex kinematic arguments are used to provide a physical explanation of this motion as well as to motivate the conditions for its occurrence. The resulting criterion for this type of stall inception (the movement of the tip clearance vortex forward of the leading edge) depends upon local flow phenomena related to the tip clearance with the implication that for this and possibly other stall mechanisms the flow structure within the blade passages must be addressed to explain the stability of an axial compression system that exhibits such short length-scale disturbances.