Recent Advances in Simulating Unsteady Flow Phenomena Brought About by Passage of Shock Waves in a Linear Turbine Cascade

1993 ◽  
Vol 115 (4) ◽  
pp. 687-698 ◽  
Author(s):  
J. C. Collie ◽  
H. L. Moses ◽  
J. A. Schetz ◽  
B. A. Gregory
Keyword(s):  
Shock Waves ◽  
Mechanical Performance ◽  
Pressure Ratio ◽  
High Pressure Ratio ◽  
Large Fluctuations ◽  
Blade Row ◽  
Flow Phenomena ◽  
Response Pressure ◽  
Unsteady Effects ◽  
Transonic Cascade

High-pressure-ratio turbines have flows dominated by shock structures that pass downstream into the next blade row in an unsteady fashion. Recent numerical results have indicated that these unsteady shocks may significantly affect the aerodynamic and mechanical performance of turbine blading. High cost and limited accessibility of turbine rotating equipment severely restrict the quantitative evaluation of the unsteady flowfield in that environment. Recently published results of the Virginia Tech transonic cascade facility indicate high integrity in simulation of the steady-state flow phenomena. The facility has recently been modified to study the unsteady effects of passing shock waves. Shock waves are generated by a shotgun blast upstream of the blade row. Shadowgraph photos and high-response pressure data are compared to previously published experimental and numerically predicted results. Plots are included that indicate large fluctuations in estimated blade lift and cascade loss.

scholarly journals Recent Advances in Simulating Unsteady Flow Phenomena Brought About by Passage of Shock Waves in a Linear Turbine Cascade

1992 ◽  
Author(s):  
J. C. Collie ◽  
H. L. Moses ◽  
J. A. Schetz ◽  
B. A. Gregory
Keyword(s):  
Shock Waves ◽  
Mechanical Performance ◽  
Pressure Ratio ◽  
High Pressure Ratio ◽  
Large Fluctuations ◽  
Blade Row ◽  
Flow Phenomena ◽  
Response Pressure ◽  
Unsteady Effects ◽  
Transonic Cascade

High-pressure ratio turbines have flows dominated by shock structures that pass downstream into the next blade row in an unsteady fashion. Recent numerical results have indicated that these unsteady shocks may significantly affect the aerodynamic and mechanical performance of turbine blading. High cost and limited accessibility of turbine rotating equipment severely restrict the quantitative evaluation of the unsteady flowfield in that environment. Recently published results of the Virginia Tech transonic cascade facility indicate high integrity in simulation of the steady state flow phenomena. The facility has recently been modified to study the unsteady effects of passing shock waves. Shock waves are genarated by a shotgun blast upstream of the blade row. Shadowgraph photos and high-response pressure data are compared to previously published experimental and numerically predicted results. Plots are included which indicate large fluctuations in estimated blade lift and cascade loss.

The Present Challenge of Transonic Compressor Blade Design

2018 ◽  
Author(s):  
A. Hergt ◽  
J. Klinner ◽  
J. Wellner ◽  
C. Willert ◽  
S. Grund ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Shock Waves ◽  
Unsteady Flow ◽  
Flow Behavior ◽  
Flow Behaviour ◽  
Compressor Cascade ◽  
Transonic Compressor ◽  
Flow Phenomena ◽  
Transonic Cascade

The flow through a transonic compressor cascade shows a very complex structure due to the occuring shock waves. In addition, the interaction of these shock waves with the blade boundary layer inherently leads to a very unsteady flow behaviour. The aim of the current investigation is to quantify this behaviour and its influence on the cascade performance as well as to describe the occuring transonic flow phenomena in detail. Therefore, an extensive experimental investigation of the flow in a transonic compressor cascade has been conducted within the transonic cascade wind tunnel of DLR at Cologne. In this process, the flow phenomena were thoroughly examined for an inflow Mach number of 1.21. The experiments investigate both, the laminar as well as the turbulent shock wave boundary layer interaction within the blade passage and the resulting unsteady behaviour. The experiments show a fluctuation range of the passage shock wave of about 10 percent chord for both cases, which is directly linked with a change of the inflow angle and of the operating point of the cascade. Thereafter, RANS simulations have been performed aiming at the verification of the reproducibility of the experimentally examined flow behavior. Here it is observed that the dominant flow effects are not reproduced by a steady numerical simulation. Therefore, a further unsteady simulation has been carried out in order to capture the unsteady flow behaviour. The results from this simulation show that the fluctuation of the passage shock wave can be reproduced but not in the correct magnitude. This leads to a remaining weak point within the design process of transonic compressor blades, because the working range will be overpredicted. The resulting conclusion of the study is that the use of scale resolving methods such as LES or the application of DNS is necessary to correctly predict unsteadiness of the transonic cascade flow and its impact on the cascade performance.

Design, Test and Analysis of a High-Pressure-Ratio Transonic Fan

2003 ◽  
Author(s):  
W. John Calvert ◽  
Paul R. Emmerson ◽  
Jon M. Moore
Keyword(s):  
High Performance ◽  
Mechanical Design ◽  
Pressure Ratio ◽  
Operating Conditions ◽  
Pressure Transducers ◽  
Technology Acquisition ◽  
Laser Anemometry ◽  
High Pressure Ratio ◽  
Response Pressure ◽  
Stage Matching

Aircraft gas turbine engines require compression systems with high performance and low weight and cost. There is therefore a continuing drive to increase compressor stage pressure ratios, particularly for military fans. To meet this need, a technology acquisition programme has been carried out by QinetiQ and Rolls-Royce. Firstly, the stage matching issues for an advanced two-stage military fan were investigated, including the effects of employing variable inlet guide vanes. From this, the requirements for the first stage together with key operating conditions for the blading were defined. The blade profiles were then designed to satisfy the range of aerodynamic conditions using a quasi-3D calculation system. A satisfactory compromise between the aerodynamic and mechanical design requirements was reached in which a blisk construction was employed for the rotor, machined from a single piece of titanium. The new stage was manufactured and tested successfully, and it achieved its target flow, pressure ratio and efficiency on the first build. Detailed measurements of the internal flows using laser anemometry and high response pressure transducers were taken. Finally, these data have been analysed and used to calibrate current 3D multi-row CFD methods.

The Present Challenge of Transonic Compressor Blade Design

2019 ◽  
Vol 141 (9) ◽  
Author(s):  
Alexander Hergt ◽  
J. Klinner ◽  
J. Wellner ◽  
C. Willert ◽  
S. Grund ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Shock Waves ◽  
Unsteady Flow ◽  
Flow Behavior ◽  
Complex Structure ◽  
Compressor Cascade ◽  
Transonic Compressor ◽  
Flow Phenomena ◽  
Transonic Cascade

The flow through a transonic compressor cascade shows a very complex structure due to the occurring shock waves. In addition, the interaction of these shock waves with the blade boundary layer inherently leads to a very unsteady flow behavior. The aim of the current investigation is to quantify this behavior and its influence on the cascade performance as well as to describe the occurring transonic flow phenomena in detail. Therefore, an extensive experimental investigation of the flow in a transonic compressor cascade has been conducted within the transonic cascade wind tunnel of DLR Institute of Propulsion Technology at Cologne. In this process, the flow phenomena were thoroughly examined for an inflow Mach number of 1.21. The experiments investigate both the laminar and the turbulent shock wave boundary layer interaction within the blade passage and the resulting unsteady behavior. The experiments show a fluctuation range of the passage shock wave of about 10% chord for both cases, which is directly linked with a change of the inflow angle and of the operating point of the cascade. Thereafter, Reynolds-averaged Navier–Stokes (RANS) simulations have been performed aiming at the verification of the reproducibility of the experimentally examined flow behavior. Here, it is observed that the dominant flow effects are not reproduced by a steady numerical simulation. Therefore, a further unsteady simulation has been carried out to capture the unsteady flow behavior. The results from this simulation show that the fluctuation of the passage shock wave can be reproduced but not in the correct magnitude. This leads to a remaining weak point within the design process of transonic compressor blades because the working range will be overpredicted. The resulting conclusion of this study is that the use of scale-resolving methods such as LES or the application of DNS is necessary to correctly predict unsteadiness of the transonic cascade flow and its impact on the cascade performance.

On the Interpretation of Measured Profile Losses in Unsteady Wake–Turbine Blade Interaction Studies

1998 ◽  
Vol 120 (2) ◽  
pp. 276-284 ◽  
Author(s):  
H. P. Hodson ◽  
W. N. Dawes
Keyword(s):  
Unsteady Flow ◽  
Stagnation Pressure ◽  
Energy Separation ◽  
Unsteady Wake ◽  
Profile Losses ◽  
Large Fluctuations ◽  
Blade Row ◽  
Measured Profile ◽  
Unsteady Effects ◽  
Stagnation Enthalpy

The interaction of wakes shed by a moving blade row with a downstream blade row causes unsteady flow. The meaning of the free-stream stagnation pressure and stagnation enthalpy in these circumstances has been examined using simple analyses, measurements, and CFD. The unsteady flow in question arises from the behavior of the wakes as so-called negative jets. The interactions of the negative jets with the downstream blades lead to fluctuations in static pressure, which in turn generate fluctuations in the stagnation pressure and stagnation enthalpy. It is shown that the fluctuations of the stagnation quantities created by unsteady effects within the blade row are far greater than those within the incoming wake. The time-mean exit profiles of the stagnation pressure and stagnation enthalpy are affected by these large fluctuations. This phenomenon of energy separation is much more significant than the distortion of the time-mean exit profiles that is caused directly by the cross-passage transport associated with the negative jet, as described by Kerrebrock and Mikolajczak. Finally, it is shown that if only time-averaged values of loss are required across a blade row, it is nevertheless sufficient to determine the time-mean exit stagnation pressure.

Shock Induced Vortices in Transonic Compressors: Aerodynamic Effects and Design Correlations

2012 ◽  
Author(s):  
Henry Knobbe ◽  
Eberhard Nicke
Keyword(s):  
Pressure Ratio ◽  
Computational Domain ◽  
Tip Leakage ◽  
Count Ratio ◽  
Flow Phenomena ◽  
Shock Mach Number ◽  
Compressor Performance ◽  
Stage Matching ◽  
Unsteady Effects ◽  
Highly Loaded

Gas turbine total pressure ratio, efficiency and accurate stage matching prediction are of increasing importance in multi-stage compressor design. These highly challenging objectives can only be met if all components are highly loaded and optimally designed. Stage matching and efficiency improvements still depend on the designer’s experience and on empirical correlations. The upper blade part in highly loaded transonic compressors is especially difficult to design because of complex flow phenomena like compression shocks. On the one hand this region is of major interest, because of the high pressure ratios. On the other hand it is the most difficult area for the designer because of blade row interaction effects, tip leakage flows and high gradients in general. Recent publications investigated shock induced vortices (SIV), caused by the rotor bow shock. The shock interacts with the trailing edge of the upstream stator/IGV blade row. These vortices convect downstream through the rotor passage. But the unsteady flow phenomena inside transonic compressors are still worthwhile endeavor because of insufficient understanding regarding the unsteady effects onto the overall compressor performance. The vortex trajectory is predictable inside the rotor-passage. However, correlations for vorticity magnitude, vortex-frequency (number of vortices) and vortex-trajectory on the overall compressor performance were never described by equations. Furthermore it has not yet been clarified whether a small or a wide axial spacing is beneficial in highly loaded axial compressors. Therefore a transonic front stage (DLR Test Rig 250, 4 front stages of a state of the art gas turbine compressor) was chosen. The Q3D (quasi 3D) planes were extracted from 3D-Simulations for IGV/Rotor1 and Stator1/Rotor2. The approach is to separate the blade to blade effects from 3D effects, like tip leakage flow. This was achieved by different Q3D streamtubes in the upper blade part. These streamtubes allow a variation of the axial spacing without changing the steady flow solution. A wide range of the axial spacings have been simulated to get an overview about the resulting performance change. Furthermore a change of blade count ratio, inlet condition, outlet conditions and computational domain should lead to a better understanding. Physical relations between the shedded vortices and compressor overall performance should be derived. The results show a wide spreading of the compressor performance speedlines. This spreading indicates the unsteady effects caused by interaction effects. The spreading becomes wider towards the surge margin. The reduced number of IGVs result into a smaller spreading. The higher inlet temperature result into a neglectable change in data spreading. The changed computational domain (stator/rotor) result into a very small data spreading, compared to the front stage data distribution. The change of performance data is periodic to the established B3-factor [1]. This factor predicts the vortex trajectory inside the rotor passage. The analysis of rotor pre shock Mach number and blade count ratio leads to a systematic correlation factor (HK). This correlation takes the pre shock Mach number, the blade count ratio, the B3-factor and some algebraic elements into account to make a prediction of the unsteady effects regarding the total pressure ratio: HK = pt, unsteady−pt, steady. The developed correlations may be useful in 3D to calculate optimal axial spacings at a specific blade to blade plane or to make compressor performance prediction during the design process based on RANS-Simulations (reduced gap between simulation and measurement). Furthermore it was identified that neither a wide nor a short axial spacing is beneficial for a transonic compressor inside a blade to blade plane.

Performance Prediction for High Pressure-Ratio Radial Inflow Turbines

2007 ◽  
Author(s):  
Xuwen Qiu ◽  
Nick Baines
Keyword(s):  
High Pressure ◽  
Early Stage ◽  
Pressure Ratio ◽  
Fluid Motion ◽  
Trailing Edge ◽  
Flow Area ◽  
High Pressure Ratio ◽  
Flow Capacity ◽  
Blade Row ◽  
Map Generation

This paper describes the latest developments in a method for predicting the design and off-design performance of radial inflow turbines, using a one-dimensional analysis. As such, it is suitable for preliminary design purposes and also for turbine map generation as an aid to the modelling of systems including such turbines. Previous development work has resulted in methods of loss correlation allowing the power output and efficiency to be predicted with confidence. The focus of this paper is on extending the calculation method to high-pressure ratios, and the accurate prediction of flow capacity for unchoked and choked conditions. A numerical method provides for the identification of subsonic, transonic, and supersonic flow regimes in the bladed rows of the turbine, and allows their solution in a consistent manner. Numerically stable and validated solutions have been obtained for a cryogenic expander case with the stage pressure ratios as high as 13.6. In this paper, we will report cases with pressure ratio up to 4.0, where the nozzle and rotor are operating at the choked condition. When a blade row is choked, the flow capacity depends on the throat area, and accurate predictions require that this area is known with confidence. Previous meanline methods have typically concentrated on unchoked flow conditions, in which it is not necessary to know the throat area accurately. In turbine design, the method thus enables the necessary throat areas to be established at an early stage in the design process, and this information is required for the subsequent blade design. In analysis, comparison with test data has revealed the importance of throat aerodynamic blockage, which has hitherto largely been overlooked in meanline prediction methods. Estimates of appropriate blockages have been obtained from such comparisons. An unusual feature of radial inflow turbine nozzles is the reduction of annulus area downstream of the blade row. This can lead to situations where it is the flow area at the trailing edge rather than the throat that limits the flow capacity in choking conditions. The method accommodates this by introducing additional deviation at the trailing edge to ensure that the throat remains choked for all blade row pressure ratios greater than the critical pressure ratio, and the flow between the throat and trailing edge develops in a form that is fully consistent with the basic principles of fluid motion.

Effects of Reynolds number on the performance of a high pressure-ratio turbocharger compressor

2013 ◽  
Vol 56 (6) ◽  
pp. 1361-1369 ◽  
Author(s):  
XinQian Zheng ◽  
Yun Lin ◽  
BinLin Gan ◽  
WeiLin Zhuge ◽  
YangJun Zhang
Keyword(s):  
Reynolds Number ◽  
High Pressure ◽  
Pressure Ratio ◽  
High Pressure Ratio

Effect of Recirculation Device With Counter Swirl Vane on Performance of High Pressure Ratio Centrifugal Compressor

2011 ◽  
Author(s):  
Hideaki Tamaki
Keyword(s):  
High Pressure ◽  
Mach Number ◽  
Flow Rate ◽  
Centrifugal Compressor ◽  
Pressure Ratio ◽  
Negative Slope ◽  
Tip Leakage Flow ◽  
Recirculation Flow ◽  
Operating Range ◽  
High Pressure Ratio

Centrifugal compressors used for turbochargers need to achieve a wide operating range. The author has developed a high pressure ratio centrifugal compressor with pressure ratio 5.7 for a marine use turbocharger. In order to enhance operating range, two different types of recirculation devices were applied. One is a conventional recirculation device. The other is a new one. The conventional recirculation device consists of an upstream slot, bleed slot and the annular cavity which connects both slots. The new recirculation device has vanes installed in the cavity. These vanes were designed to provide recirculation flow with negative preswirl at the impeller inlet, a swirl counterwise to the impeller rotational direction. The benefits of the application of both of the recirculation devices were ensured. The new device in particular, shifted surge line to a lower flow rate compared to the conventional device. This paper discusses how the new recirculation device affects the flow field in the above transonic centrifugal compressor by using steady 3-D calculations. Since the conventional recirculation device injects the flow with positive preswirl at the impeller inlet, the major difference between the conventional and new recirculation device is the direction of preswirl that the recirculation flow brings to the impeller inlet. This study focuses on two effects which preswirl of the recirculation flow will generate. (1) Additional work transfer from impeller to fluid. (2) Increase or decrease of relative Mach number. Negative preswirl increases work transfer from the impeller to fluid as the flow rate reduces. It increases negative slope on pressure ratio characteristics. Hence the recirculation flow with negative preswirl will contribute to stability of the compressor. Negative preswirl also increases the relative Mach number at the impeller inlet. It moves shock downstream compared to the conventional recirculation device. It leads to the suppression of the extension of blockage due to the interaction of shock with tip leakage flow.

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