Supersonic Compressor Rotor Research Articles

Controlling the flow loss of a supersonic compressor rotor using the blade slotting method

The flow loss in the blade passage of a supersonic compressor rotor mainly comes from the boundary layers on the blade surface and end wall, the shock wave, the shock wave/boundary layer interaction, and tip leakage flow; the instability is mainly caused by the shock wave near the rotor blade tip exiting the blade passage. This paper adopts an internal slot in the blade, with the inlet of the slot located at the leading edge of the blade and the outlet located on the suction surface of the blade, by using the momentum of the incoming flow to form a high-velocity jet to control the flow loss and improve the stall margin of the supersonic rotor. The mechanism of reducing flow loss by a slotting jet was studied, and a genetic algorithm optimization platform was further used for the coupled optimization design of the slot and blade. The numerical calculation results showed that the slotting jet can effectively suppress the development of the boundary layer on the suction surface while reducing the intensity of the shock wave, thereby reducing the loss of the boundary layer and shock wave, significantly improving the peak efficiency of the rotor, and increasing the mass flow rate at the peak efficiency point. The slotting jet can cause the shock wave in the passage to move downstream, thereby improving the stall margin of the rotor. Due to the strong shock wave in the blade passage near the blade tip, the slot outlet should be near and upstream of the shock wave; the shock wave in the middle and root regions of the blade is weaker, and the slot outlet should be located downstream of the shock wave.

Hub aspiration strategies and flow loss analysis of a shock-in type supersonic compressor rotor with negative inlet pre-whirl

Large-scale corner separation is one of the key factors that limit the performance improvement of high-load compressors. In this paper, the hub aspiration method was applied to a supersonic shock-in type compressor rotor with negative inlet pre-whirl to suppress the hub corner separation. Six suction schemes composed of aspiration slots with different locations and orientations have been designed and numerically tested. The results revealed that the aspiration location is better to be placed starting from the post-shock point and extending to the deep separation zone. The streamwise slot satisfying such conditions could effectively suppress corner separation but narrows the rotor stable mass flow range by increasing the pitchwise secondary flow. The circumferential skew slot could exactly remedy such defects. The streamwise slot combined with a downstream skew slot is the optimal aspiration strategy that effectively improves the rotor performance without reducing mass flow range. At the design and near-stall operating point, the rotor efficiency (EF) and total pressure ratio (TPR) could be improved by 1.3% and 5.0%, 3.3% and 7.2%, respectively. Consequently, the stall margin was increased by 2.0%. According to the flow loss analysis, the combined aspiration scheme mainly alleviated losses from hub to middle span by reducing the shock strength and the magnitude of separation as well as its radial migration. But in the tip region, it strengthened the leakage flow and its interaction with shock waves, thus increasing the flow losses.

Aerodynamic Optimization Design of a Supersonic Compressor Rotor with High Pressure Ratio

Supersonic compressors have a high wheel speed and operational capability, which facilitate a high stage pressure ratio. However, the strong shock waves in the passage of a supersonic rotor and the interference between shock waves and boundary layers can lead to large flow loss and low efficiency. Moreover, the existing design of a high-load supersonic compressor has the problem of small stall margin. In this study, an automatic optimization method including 2D profile optimization and 3D blade optimization is proposed to achieve a high efficiency at the design point of a supersonic compressor rotor under the premise of reaching the desired mass flow rate and total pressure ratio. According to the analysis of flow near the stall point of the supersonic compressor rotor, the mechanism responsible for rotor tip stall is established, that is, the aerodynamic throat appeared inside the flow passage, reducing the ability of the blade tip to withstand back pressure, and the low-speed areas caused by the tip-leakage-vortex breakage and boundary layer separation reduced the flow capacity of the blade tip. Based on the reasons for rotor stall, three methods are proposed to improve the stall margin, which include increasing the exit radius of the upper meridian, forward sweep of the blade tip, and increasing the chord length of the blade tip. The above method is used to design a supersonic rotor with a total pressure ratio of 2.8, which exhibits an efficiency of 0.902 at the design point and a stall margin of 18.11%.

Design and Experimental Validation of a Supersonic Concentric Micro Gas Turbine

This paper presents the design and experimental results of a new micro gas turbine architecture exploiting counterflow within a single supersonic rotor. This new architecture, called the supersonic rim-rotor gas turbine (SRGT), uses a single rotating assembly incorporating a central hub, a supersonic turbine rotor, a supersonic compressor rotor, and a rim-rotor. This SRGT architecture can potentially increase engine power density while significantly reducing manufacturing costs. The paper presents the preliminary design of a 5 kW SRGT prototype having an external diameter of 72.5 mm and rotational speed of 125,000 rpm. The proposed aerodynamic design comprises a single stage supersonic axial compressor, with a normal shock in the stator, and a supersonic impulse turbine. A pressure ratio of 2.75 with a mass flow rate of 130 g/s is predicted using a 1D aerodynamic model in steady state. The proposed combustion chamber uses an annular reverse-flow configuration, using hydrogen as fuel. The analytical design of the combustion chamber is based on a 0D model with three zones (primary, secondary, and dilution), and computational fluid dynamics (CFD) simulations are used to validate the analytical model. The proposed structural design incorporates a unidirectional carbon-fiber-reinforced polymer rim-rotor, and titanium alloy is used for the other rotating components. An analytical structural model and numerical validation predict structural integrity of the engine at steady-state operation up to 1000 K for the turbine blades. Experimentation has resulted in the overall engine performance evaluation. Experimentation also demonstrated a stable hydrogen flame in the combustion chamber and structural integrity of the engine for at least 30 s of steady-state operation at 1000 K.

A Numerical Study on the Effect of Bleed System on Starting Ability and Flow Performance of Rampressor Inlet

A unique supersonic compressor rotor with high pressure ratio, termed the Rampressor, is presented by Ramgen Power Systems, Inc., (RPS). Based on the models of Rampressor inlet, the inlet flow field with bleed system is numerically studied. Validation of the employed computational fluid dynamics (CFD) scheme is provided through test cases. The effects of boundary layer bleed location and bleed amount on Rampressor rotor inlet start and flow performance are analyzed. The results indicate that the boundary layer bleed has a significant effect for start and flow performance of Rampressor inlet. Boundary layer bleed technique has been applied to eliminate the emerging flow separation zone for enhancing Rampressor rotor inlet performance and enlarging its stable working range. The starting ability and flow performance of Rampressor inlet are efficiently improved by bleeding system, but the improvement effect is different for Rampressor inlet with different bleed location. Along the position of bleeding system moves forward, the range of Rampressor inlet normal work rotation speed is enlarged. The flow performance of Rampressor inlet improves obviously with the increment of bleed flow rate, and exit stability of Rampressor inlet enhances. And in the same back pressure work condition of Rampressor inlet, bleed system has been shown to be effective that exit stability of Rampressor inlet ameliorates, but the loss of compressed air from the bleed system has a negative effect on overall Rampressor inlet efficiency.

A Numerical Study for Flow Excitation and Performance of Rampressor Inlet considering Rotor Motion

A unique supersonic compressor rotor with high pressure ratio, termed the Rampressor, is presented by Ramgen Power Systems, Inc. (RPS). In order to obtain the excitation characteristic and performance of Rampressor inlet flow field under external excitation, compression inlet flow of Rampressor is studied with considering Rampressor rotor whirling. Flow excitation characteristics and performance of Rampressor inlet are analyzed under different frequency and amplitude of Rampressor rotor whirling. The results indicate that the rotor whirling has a significant effect for flow excitation characteristics and performance of Rampressor inlet. The effect of rotor whirling on the different inlet location excitation has a definite phase difference. Inlet excitation becomes more complex along with the inlet flow path. More frequency components appear in the excitation spectrum of Rampressor inlet with considering Rampressor rotor whirling. The main frequency component is the fundamental frequency, which is caused by the rotor whirling. Besides the fundamental frequency, the double frequency components are generated due to the coupling between inlet compression flow of Rampressor rotor and rotor whirling, especially in the subsonic diffuser of Rampressor rotor inlet. With the increment of rotor whirling frequency and whirling amplitude, the complexity of Rampressor inlet excitation increases, and the stability of Rampressor inlet performance deteriorates.

Supersonic Compressor Cascades—An Analysis of the Entrance Region Flow Field Containing Detached Shock Waves

When the incoming air flow to a supersonic compressor rotor or cascade has a subsonic axial velocity component, a system of shock and Mach waves originates in the blade leading-edge region, propagates upstream of successive blades, and alters the incident flow conditions. The continuity constraint establishes a unique relationship between the incident Mach number and air flow angle. The shocks are detached and curved with blunt leading edged blades thereby violating the assumptions inherent in the classical simple wave analysis of this flow induction process. The current work employs detached shock models and a rotational flow analysis by the method of characteristics to investigate the effects of airfoil leading edge radius and cascade solidity on the unique incidence values and entrance region total pressure losses. The calculated results are compared to simple wave theory for the entrance region and it is shown that the simple wave assumptions lead to appreciable errors when detached shocks are present in the flow field.

A Contribution to the Theoretical and Experimental Examination of the Flow Through Plane Supersonic Deceleration Cascades and Supersonic Compressor Rotors

The results of the experimental investigations of plane supersonic deceleration cascades (tandem cascades) conducted in recent years at the Institut fu¨r Strahlantriebe und Turboarbeitsmaschinen der Technischen Hochschule Aachen (Institute for Jet Propulsion and Turbine Machines of the Aachen Technical University) are summarized. The flow through a supersonic compressor rotor is determined by means of a suitably enlarged method of characteristics. A division of the blade channel into two sets of stream surfaces permits an iterative calculation of the three-dimensional flow. To determine the flow values in the subsonic range, 5-hole-probes featuring a semi-spherical head are used in front of the rotor. The supersonic flow is determined by means of 5-hole-probes of the conical type arranged at varying distances behind the rotor. To determine the efficiencies, all probes were equipped with two NTC-elements each, for temperature measurement. These measurements were carried out at different rotational speeds and throttle positions and supplied rotor characteristics as well as information on the course of the flow behind the rotor. With unthrottled flow, a total pressure ratio of 3.9 resulted, with a total isentropic efficiency of 89 percent. The measurement of wall pressures on hub and tip furnishes additional information on the course of flow. The test values measured show a satisfactory agreement with the calculated design values.