Highly loaded compressor with boundary layer suction

An experimental and numerical investigation of boundary layer suction (BLS) on a transonic compressor blade has been conducted. The objective of the present work is to identify the possible benefit of boundary layer suction via a slot in the region of the shock/boundary layer interaction (SBLI). The study has two major parts: First, an investigation of the flow on an isolated airfoil without and with BLS on the suction side (SS). This investigation is intended to optimise the suction slot geometry in terms of location and width. Second, a study on a transonic cascade with supersonic inlet flow has been conducted with the goal to show the influence of the BLS on the pressure loss coefficient and the surface Mach number distribution. An initial literature study revealed that only few work has been carried out in the field of BLS in transonic compressors and no data close to the suction slot was available. The main steps of the present study are: The design of an isolated airfoil for testing in a Laval nozzle, representative for the flow at a compressor root section, with enough space inside to evacuate the aspirated air. A numerical study of the flow on the isolated airfoil without and with BLS on the SS . An experimental investigation of the isolated airfoil without and with BLS on the SS using static surface pressure taps, PSP and the Schlieren method for data acquisition. Design and numerical investigation of a suitable transonic compressor cascade with enough space inside to evacuate the suction mass flow. An experimental investigation of the designed transonic cascade with an inlet Mach number of 1.23 with two sets of blades in the non-rotating annular cascade test-rig; one without and one with BLS at 40% chord. The main results are: The numerical study of the flow on the isolated airfoil without and with BLS on the SS shows, that the onset of shock-induced separation could be shifted to an increased inlet Mach number. Steady 2D-NS simulations show that the reference configuration exhibits an attached flow behind the shock up to an inlet Mach number of up to 0.70 for an inlet flow angle of +4°. The configuration with BLS shows attached flow behind the shock for an an inlet Mach number of 0.725. Unsteady 2D-NS simulations show for an inlet flow angle of +4° at an inlet Mach number of 0.725 for the reference configuration a shock, which moves periodically on the SS due to a separated boundary layer created by the shock/boundary layer interaction. The same configuration with BLS shows a stable shock position. The experimental investigation of the isolated airfoil without and with BLS on the SS shows that BLS at 15% chord upstream of the initial shock location for an inlet Mach number of 0.725 at a flow angle of +4° leads to a shock location 10% chord downstream of the initial position. Boundary layer suction suppresses flow separation and stabilises the shock. Due to the suction a stagnation point is created at the downstream edge of the suction slot. Downstream of this stagnation point, the flow is re-accelerated to the isentropic Mach number level of the reference configuration without BLS. The design and 3D-numerical investigation of a suitable transonic compressor cascade shows, that BLS with a suction mass flow of 2% can increase the maximum pressure ratio and diffusion factor of approximately 10%. This is the consequence of the stabilisation of the shock in the blade passage due to the suction. The experimental investigation of the designed transonic cascade with an inlet Mach number of 1.23 with 2% suction mass flow shows, that BLS acts locally and that suction leads to an increased isentropic surface Mach number in front of the suction slot and due to the suction related stagnation point, to a locally decreased velocity immediately downstream of the suction slot. Then the flow re-accelerates to the same level as the non-suction configuration attains. The pressure loss coefficient is decreased by up to 3.6% for the most positive incidences. The results of this investigation confirm that boundary layer suction on the suction side of an airfoil is a suitable way to increase the working range of a compressor as the shock position can be stabilised (shock-trapping) and the increase of the pressure loss coefficient can be shifted to higher incidences.

Numerical and experimental investigation of the shock-vane interaction in a transonic compressor

Stefan Kristukat

A numerical and experimental investigation of the shock-IGV (inlet guide vane) interaction in a transonic compressor is conducted at the Laboratoire de Thermique Appliquée et de Turbomachines (LTT) at the Ecole Polytechnique Fédérale de Lausanne (EPFL). The objective of the presented research work is to determine the impact of the upstream propagating shock pattern of an axial transonic compressor rotor on an aerodynamically loaded IGV and the influence of this interacting process on the compressor flow field for different operating points and aerodynamic setups. For the approach to the subject, a literature review shed light on the different research efforts over the past years. A large numerical study is conducted including the work with two different flow solvers. A numerical 3D transient analysis (commercial code) was conducted in order to provide information about the influence of the axial spacing on the overall mass flow through the stage. A numerical Q-3D transient analysis (industry owned code) on one streamline at a constant rotor inlet Mach number is performed in order to compare the shock-IGV interaction for different vane types similar to what is known in literature. A semi-analytical model was developed in order to evaluate in a fast approach the impact of a rotor bow shock on an upstream situated vane. The results of the numerical investigation can be summarized as following: The overall results of this investigation show that the upstream vane of a transonic compressor rotor transforms the rotor forcing function into an excitation with vane shape depending orientation and force. The time averaged flow field of the IGV does not show a measurable change of the exit flow angle due to the shock-IGV interaction for the studied operating points and axial distances. The 3D transient calculations do not show any influence of the examined axial spacing between vane and blade on the mass flow rate through the stage. The results of the Q3D analysis are compared with the semi-analytic approach and the measurements, both showing with some restrictions good accordance to the model. The results of the numerical investigation give an accurate prediction of the transient aerodynamic load of the upstream vane. An analytic tool was developed which uses three input parameters, the rotor relative inlet Mach number, the stagger angle of the rotor and the axial spacing between vane and rotor in order to provide an estimation of the amplitude of the pressure variation on the upstream situated vane. An extensive modification of the existing test facility of the LTT precedes the experimental study. The main steps of the approach are: A subsonic axial compressor test facility was modified to allow test runs at transonic rotor inlet conditions. A perfect leak tightness of the closed circuit was accomplished to enable its operation in heavy gas. Heavy gases like refrigerants have been used in many studies for compressor testing. While respecting certain restrictions during the use of heavy gas, a reduction of operation and manufacturing costs of the facility is beneficial. A high molecular weight of the gas reduces the sonic speed to the half of that of air thus reducing the dimension and the rotational speed of the rotor significantly and consequently reduces the demand in the material quality of the blading. A 1.5 stage transonic compressor had been designed and manufactured for the heavy gas (R134a) in a first step. Preliminary measurements showed that the influence of the downstream stator makes it difficult to separate the influence of shock-IGV interaction on the compressor flow field from loss generation mechanisms downstream of the rotor. Thus, the stator was removed and measurements were conducted in a one stage transonic compressor. Modifications on the construction of the testing setup are done in a way that the distance between IGV and rotor can be varied by means of spacers. Thus the flow can be determined for up to four different axial distances at varying rotational speed for different operating points. For transonic operation the time-dependent load of the IGV (inlet guide vane) due to the forcing function of the rotor bow shock was recorded with a high speed data acquisition system. The results of the experimental investigation can be summarized as following: No influence of the axial spacing on the efficiency and the mass flow was observed in the investigation. The operation characteristics remained equal between the minimum and the maximum vane-blade spacing. The measured global values of the stage are close to the values the design was made for, although the overall efficiency was generally lower than predicted. The experimental results of the pressure variation on the upstream vane agree well with the predicted values. The study in general showed a good matching between the numerical and the experimental investigation taking into account a rotor blade shape which is not optimized to the extend. The shock-vane interaction is responsible for an excitation of the entire vane blade. With decreasing distance between the rotor leading edge and vane trailing edge the excitation increases. The strength and orientation of the reflected shock/pressure wave part depends upon the vane shape. A more sophisticated design of the rotor blade shape should be in the focus of further investigations using this test facility at transonic rotor inlet conditions. Within this investigation a tool was developed in order to predict the load variation on the vane driven by the transient rotor forcing function. The tool maps the fact that the pressure on the vane surface oriented against rotational direction of the rotor exceeds the pressure of the passage shock due to the amplification effect of the reflection of the shock on a surface. Its possible application not only for compressors but also for turbine related topics should be an objective for similar investigations.

EPFL2008

Aerodynamic study of a Hyperloop pod equipped with compressor to overcome the Kantrowitz limit

Yohei Sato

This paper describes an investigation of the aerodynamic performance of a Hyperloop pod equipped with an axial compressor using CFD simulation. The compressor is expected to reduce the drag if the operational speed of the pod exceeds the Kantrowitz Limit (KL). To evaluate the effectiveness of the compressor, two types of pod have been designed: those with and those without a compressor. A validation study was undertaken for supersonic flow around a cylindrical body in a wind tunnel. Several combinations of pod Mach number and blockage ratios were examined. The axial compressor was modeled with the fan interface model and our original method, which we call the source terms model. Results are similar, but our model has the advantage that it can be applied to higher pressure ratios without inducing numerical instabilities. The power consumption was reduced owing to the compressor by decreasing the pressure accumulated in front of the pod. The effectiveness of the compressor was shown to be greater for conditions well above the KL, with a maximum power reduction of 47.5%. Simulations performed half-scale, and at a different system pressure, showed that the effectiveness of the compressor was not greatly influenced by these parameters.

2021

Numerical and experimental investigation of the shock-vane interaction in a transonic compressor

Stefan Kristukat

EPFL2008