Experimental Evaluation of the High-to-Low Speed Transformation Process for a Highly Loaded Core Compressor Stage

Journal of Engineering for Gas Turbines and Power

et al. 2018

Modern high-pressure turbine blades operate at high-speed conditions. The over-tip-leakage (OTL) flow can be high-subsonic or even transonic. From the consideration of problem simplification and cost reduction, the OTL flow has been studied extensively in low-speed experiments. It has been assumed a redesigned low-speed blade profile with a matched blade loading should be sufficient to scale the high-speed OTL flow down to the low-speed condition. In this paper, the validity of this conventional scaling approach is computationally examined. The computational fluid dynamics (CFD) methodology was first validated by experimental data conducted in both high- and low-speed conditions. Detailed analyses on the OTL flows at high- and low-speed conditions indicate that, only matching the loading distribution with a redesigned blade cannot ensure the match of the aerodynamic performance at the low-speed condition with that at the high-speed condition. Specifically, the discrepancy in the peak tip leakage mass flux can be as high as 22%, and the total pressure loss at the low-speed condition is 6% higher than the high-speed case. An improved scaling method is proposed hereof. As an additional dimension variable, the tip clearance can also be “scaled” down from the high-speed to low-speed case to match the cross-tip pressure gradient between pressure and suction surfaces. The similarity in terms of the overall aerodynamic loss and local leakage flow distribution can be improved by adjusting the tip clearance, either uniformly or locally.

Section: Introductionmentioning

confidence: 99%

On Scaling Method to Investigate High-Speed Over-Tip-Leakage Flow at Low-Speed Condition

Jiang

Journal of Engineering for Gas Turbines and Power

et al. 2018

“…The increasing understanding of its intemal intricate flow and the flow mechanism can largely promote the improvement of compressor design ability. However, in view of its small size, high cost, and expensive highspeed test research to develop new design concepts, more and more low-speed large-scale test rigs were utilized to model the complex flow of the middle and late stages of high pressure compressors, such as the low-speed model testing of E'' carried out by GE [3][4][5], low-speed experimental work focused on the fourth stage of a five-stage high pressure compressor C147 in Cranfield University [6][7][8], high-speed and low-speed test research of sweep and dihedral technology on the multistage axial fiow compressor of the Trent engine carried out by Rolls-Royce and Cambridge University [9,10], and also the low-speed modeling work carried out in Nanjing University of Aeronautics and Astronautics during the past five years [11][12][13]. All of the above-mentioned studies have achieved good results.…”

Section: Introductionmentioning

confidence: 99%

Design Work of a Compressor Stage Through High-To-Low Speed Compressor Transformation

Journal of Engineering for Gas Turbines and Power

Wang

et al. 2014

Low-speed model testing has advantages such as great accuracy and low cost and risk, so it is widely used in the design procedure of the high pressure compressor (HPC) exit stage. The low-speed model testing project is conducted in Nanjing University of Aeronautics and Astronautics (NUAA) to represent aerodynamic load and flow field structure of the seventh stage of a high-performance ten-stage high-pressure compressor. This paper outlines the design work of the low speed four-stage axial compressor, the third stage of which is the testing stage. The first two stages and the last stage provide the compressor with entrance and exit conditions, respectively. The high-to-low speed transformation process involves both geometric and aerodynamic considerations. Accurate similarities demand the same Mach number and Reynolds number, which will not be maintained due to motor power/size and its low-speed feature. Compromises of constraints are obvious. Modeling principles are presented in high-to-low speed transformation. Design work was carried out based on these principles. Four main procedures were conducted successively in the general design, including establishment of low-speed modeling target, global parameter design of modeling stage. throughflow aerodynamic design, and blading design. In global parameter design procedure, rotational speed, shroud diameter, hub-tip ratio, midspan chord, and axial spacing between stages were determined by geometrical modeling principles. During the throughflow design process, radial distributions of aerodynamic parameters such as D-factor, pressure-rise coefficient, loss coefficients, stage reaction, and other parameters were obtained by determined aerodynamic modeling principles. Finally, rotor and stator blade profiles of the low speed research compressor (LSRC) at seven span locations were adjusted to make sure that blade surface pressure coefficients agree well with that of the HPC. Three-dimensional fiow calculations were performed on the low-speed four-stage axial compressor, and the resultant flow field structures agree well with that of the HPC. It is worth noting that a large separation zone appears in both suction surfaces of LSRC and HPC. How to diminish it through 3D blading design in the LSRC test rig is our further work.

“…To improve compressor per formance and expand stability range, many passive/active control methods were proposed by researchers, such as casing treatment [4,5], tip micro-injection [6,7], endwall blade profiling [8,9], and positive effects were gained. Study on the endwall flow evolving into the passage is not only an impor tant but also a difficult issue in the past few years.…”

Section: Introductionmentioning

confidence: 99%

“…

In v e s tig a tio n s on th e E ffects of In flo w C o n d itio n and T ip C le a ra n c e S ize to th e P e rfo rm a n c e of a C o m p re s s o r R o to r

To clearly clarify the effects of different upstream boundary layer thickness and tip clear ance size to the detailed tip flow field and flow mechanism, numerical simulations are performed on a subsonic compressor rotor, which is used for low-speed model testing of a rear stage embedded in a modern high-pressure compressor. To improve compressor per formance and expand stability range, many passive/active control methods were proposed by researchers, such as casing treatment [4,5], tip micro-injection [6,7], endwall blade profiling [8,9], and positive effects were gained. Second, comparisons are made for tip leakage vortex (TLV) structure, the interface of leakage flowlmainflow, endwall loss, isentropic efficiency and pressure-rise among different operating conditions.

…”

mentioning

confidence: 99%

Investigations on the Effects of Inflow Condition and Tip Clearance Size to the Performance of a Compressor Rotor

Journal of Engineering for Gas Turbines and Power

Wang

2014

In v e s tig a tio n s on th e E ffects of In flo w C o n d itio n and T ip C le a ra n c e S ize to th e P e rfo rm a n c e of a C o m p re s s o r R o to rTo clearly clarify the effects of different upstream boundary layer thickness and tip clear ance size to the detailed tip flow field and flow mechanism, numerical simulations are performed on a subsonic compressor rotor, which is used for low-speed model testing of a rear stage embedded in a modern high-pressure compressor. First, available experi mental data are adopted to validate the numerical method. Second, comparisons are made for tip leakage vortex (TLV) structure, the interface of leakage flowlmainflow, endwall loss, isentropic efficiency and pressure-rise among different operating conditions. Then, effects of different clearance sizes and inflow boundary layer thicknesses are inves tigated. Finally, the self-induced unsteadiness at one near-stall (NS) operating condition is studied for different cases. Results show that the increment of tip clearance size has a deleterious effect on rotor efficiency and pressure-rise peiformance over the whole oper ating range, while thickening the inflow boundary layer is almost the same except that its pressure-rise performance will be increased at mass flow rate larger than design operat ing condition. Self-induced unsteadiness occurs at NS operating conditions, and its appearance largely depends on tip clearance size, while the effect of upstream boundary layer thickness is little. In tro d u c tio nAs one of the most important components in an aero-engine, the compressor directly determines its aerodynamic performance and stability margin. In addition, upstream conditions of the com pressor have a great impact on its working condition and perform ance. Due to this, it is urgent and essential to carry out detailed and indepth studies on the effect of inlet boundary condition to compressor performance and its internal flow mechanism. A cer tain clearance exists between casing and blade tip in axial com pressors. Studies of Brandt et al. [1] show that thickening the inflow boundary layer will cause the rollup point of the leakage vortex to move upstream and toward the leading edge, while the vortex trajectory is more inclined in the blade passage and the vor tex length decreases. This leads to higher total pressure loss in the endwall region.Flow in and around compressor tip clearance is very complex and has a profound effect on compressor performance. Study on the endwall flow evolving into the passage is not only an impor tant but also a difficult issue in the past few years. Complex endwall flow consists of tip leakage flow, secondary flow, interactions of leakage flow and endwall boundary layer, and interactions between leakage flow and mainflow. In addition, shock wave establishes in transonic flows, which will incite interactions of shock wave and other flows [2,3]. To improve compressor per formance and expand stability range, many passive/active control methods were proposed by researchers, such as casing t...