Investigation of the Influence of Trailing Edge Shock Waves on Film Cooling Performance of Gas Turbine Airfoils

2007 ◽  
Author(s):  
M. Ochs ◽  
A. Schulz ◽  
H.-J. Bauer
Keyword(s):  
Mach Number ◽  
Shock Waves ◽  
Film Cooling ◽  
Suction Side ◽  
Test Rig ◽  
Local Cooling ◽  
Trailing Edge ◽  
Number Distribution ◽  
Mach Number Distribution ◽  
Turbine Airfoils

Transonic turbine stage flows are strongly influenced by shock waves. The oblique trailing edge shock generated at the pressure side impinges on the suction side of the neighboring airfoil leading to a significant alteration of the Mach number distribution. On film cooled turbine airfoils this shock interacts with the local cooling film. The present study deals with the investigation of this kind of shock wave – film cooling interaction. Experiments are conducted in a high pressure high temperature transonic test rig which allows setting engine realistic Reynolds numbers and Mach numbers, as well as temperature and density ratios. The generic test rig simulates a transonic region of an airfoil passage with the advantage of accessibility for optical measurement techniques. Coolant is ejected from a row of 5 cylindrical and 5 fanshaped holes at different locations relative to the position of shock impingement. Blowing ratios are varied within a range of 0.25<M<1.5. A simulated suction side Mach number distribution is generated with a Mach number Mam = 1.45 upstream and Mam = 1.14 downstream of the shock. Experimental data presented comprise spatially resolved and laterally averaged film cooling effectiveness and heat transfer coefficients within the vicinity of the interaction zone.

Numerical Investigations of Wake and Shock Wave Effects on Film Cooling Performance in a Transonic Turbine Stage: Part 1 — Methodology Development and Qualification Over Stationary Stators and Rotors

2013 ◽  
Author(s):  
Huazhao Xu ◽  
Jianhua Wang ◽  
Ting Wang
Keyword(s):  
Experimental Data ◽  
Mach Number ◽  
Film Cooling ◽  
Suction Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Numerical Investigations ◽  
Blowing Ratio

To understand the unsteady shock wave and wake effects on the film cooling performance over a transonic 3-D rotating stage, a series of numerical investigations have been conducted and are presented in this two-part paper. Part 1 is focused on the development of the computational model and methodology of the system setup and model qualification; Part 2 is to investigate the unsteady effects of shock waves and wakes on film cooling performance in a transonic rotating stage. In Part 1, the film cooling experimental conditions (non-rotating) and test sections of Kopper et. al. and Hunter are selected for model qualification. The numerical computation is carried out by the commercial software Ansys/Fluent using the pressure based compressible flow governing equations. The effects of four turbulence models are carefully compared with the experimental data. The Realizable k-ε turbulence model is found to match the experimental data better than the other models and is thus used for the rest of the study, including Part 2. The results show that 1) the weak shock emanating from the neighboring stator’s trailing edge results in a temperature rise and a reduction of film cooling effectiveness on the suction side near the trailing edge, 2) cooling ejection from the trailing edge reduces the shock strength in the stator passage, 3) an increase in Mach number from 0.84 to 1.50 can reduce the total pressure losses of fluid flow near the end-walls, 4) the film cooling effectiveness increases with increasing blowing ratio and becomes more even on the stator with a higher blowing ratio, and 5) an increase in Mach number from 0.84 to 1.50 gives rise to a higher cooling effectiveness in the region from the cooling holes to 80% of the chord length of the stator on the pressure side, but becomes lower after this up to the trailing edge. However, on the stator’s suction side, higher Mach number results in a lower cooling effectiveness region around the film holes from 30% to 55% of the chord length, but cooling effectiveness increases downstream.

A Numerical Investigation of Stator-Rotor Interaction Effects on Flow Field and Film Cooling Effectiveness in a 3D Transonic Turbine Stage With Highly Twisted Rotors

2013 ◽  
Author(s):  
Huazhao Xu ◽  
Jianhua Wang ◽  
Ting Wang
Keyword(s):  
Shock Waves ◽  
Film Cooling ◽  
Static Pressure ◽  
Pressure Fluctuations ◽  
Suction Side ◽  
Turbine Stage ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Air

To reduce aerodynamic losses and optimize turbine blade cooling designs, a comprehensive understanding of rotor-stator interaction effects on the blade aerodynamics and film cooling performance is essential. This paper focuses on the numerical analysis of the interactions between shock waves and unsteady wakes and their effects on cooling effectiveness of a highly twisted rotor within a transonic turbine stage. The parameters of the turbine stage are from the Pratt & Whitney Energy Efficient Engine (E3) program. The Realizable k-ε turbulence model was selected as the suitable turbulence model by our previous study. The investigation is conducted first by analyzing mean static pressure and the Root Mean Square (RMS) of the static pressure, followed by a detailed study of the flow field in the rotor passage at blowing ratios (Br) of 0.5, 1.0 and 1.5. Effects of the complicated interactions among shock waves, trailing edge wake shedding, and blockage of moving rotors are separated and identified individually through shock strength, vortices, and entropy production. The results show that: 1) For the stator, the shock waves emanating from the trailing edge of the neighboring stator impinging on the later part of the stator’s suction side, creating static pressure fluctuations as large as 20%. 2) For the rotor, the variation of static pressure is synchronized with the rotor passing frequency, but out of phase between the suction and pressure sides. 3) A high entropy region generated by the wake flow from the upstream trailing edge in the rotor passage intensifies and moves towards the rotor hub during the rotor passing periods. 4) Most of the cooling air injected from the rotor leading edge bends towards the suction side, and the cooling air injected from the pressure side turns towards the rotor hub. 5) An increase in the blowing ratio from Br = 0.5 to Br = 1.5 does not affect the pressure fluctuations, but does significantly increase film cooling effectiveness on the rotor pressure side. 6) The mean static pressure on the suction side of the twist blade is lower than a straight blade, indicating the benefit of producing larger torque by using twist rotors.

The Effect of Pulsed Blowing on the Boundary Layer of a Highly Loaded Low Pressure Turbine Blade

2013 ◽  
Author(s):  
Marion Mack ◽  
Roland Brachmanski ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Mach Number ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Trailing Edge ◽  
Low Pressure Turbine ◽  
Wide Range ◽  
Highly Loaded

The performance of the low pressure turbine (LPT) can vary appreciably, because this component operates under a wide range of Reynolds numbers. At higher Reynolds numbers, mid and aft loaded profiles have the advantage that transition of suction side boundary layer happens further downstream than at front loaded profiles, resulting in lower profile loss. At lower Reynolds numbers, aft loading of the blade can mean that if a suction side separation exists, it may remain open up to the trailing edge. This is especially the case when blade lift is increased via increased pitch to chord ratio. There is a trend in research towards exploring the effect of coupling boundary layer control with highly loaded turbine blades, in order to maximize performance over the full relevant Reynolds number range. In an earlier work, pulsed blowing with fluidic oscillators was shown to be effective in reducing the extent of the separated flow region and to significantly decrease the profile losses caused by separation over a wide range of Reynolds numbers. These experiments were carried out in the High-Speed Cascade Wind Tunnel of the German Federal Armed Forces University Munich, Germany, which allows to capture the effects of pulsed blowing at engine relevant conditions. The assumed control mechanism was the triggering of boundary layer transition by excitation of the Tollmien-Schlichting waves. The current work aims to gain further insight into the effects of pulsed blowing. It investigates the effect of a highly efficient configuration of pulsed blowing at a frequency of 9.5 kHz on the boundary layer at a Reynolds number of 70000 and exit Mach number of 0.6. The boundary layer profiles were measured at five positions between peak Mach number and the trailing edge with hot wire anemometry and pneumatic probes. Experiments were conducted with and without actuation under steady as well as periodically unsteady inflow conditions. The results show the development of the boundary layer and its interaction with incoming wakes. It is shown that pulsed blowing accelerates transition over the separation bubble and drastically reduces the boundary layer thickness.

Mach number distribution over the axis of supersonic underexpanded jets

1968 ◽  
Vol 15 (6) ◽  
pp. 1153-1157 ◽  
Author(s):  
Yu. P. Finat'ev ◽  
L. A. Shcherbakov ◽  
N. M. Gorskaya
Keyword(s):  
Mach Number ◽  
Underexpanded Jets ◽  
Number Distribution ◽  
Mach Number Distribution

Shock Wave - Film Cooling Interactions in Transonic Flows

2001 ◽  
Author(s):  
P. M. Ligrani ◽  
C. Saumweber ◽  
A. Schulz ◽  
S. Wittig
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Shock Waves ◽  
Film Cooling ◽  
Cross Flow ◽  
Injection Velocity ◽  
Film Cooling Effectiveness ◽  
Mach Numbers ◽  
Density Ratios ◽  
Film Cooling Holes

Interactions between shock waves and film cooling are described as they affect magnitudes of local and spanwise-averaged adiabatic film cooling effectiveness distributions. A row of three cylindrical holes is employed. Spanwise spacing of holes is 4 diameters, and inclination angle is 30 degrees. Freestream Mach numbers of 0.8 and 1.10–1.12 are used, with coolant to freestream density ratios of 1.5–1.6. Shadowgraph images show different shock structures as the blowing ratio is changed, and as the condition employed for injection of film into the cooling holes is altered. Investigated are film plenum conditions, as well as perpendicular film injection cross-flow Mach numbers of 0.15, 0.3, and 0.6. Dramatic changes to local and spanwise-averaged adiabatic film effectiveness distributions are then observed as different shock wave structures develop in the immediate vicinity of the film-cooling holes. Variations are especially evident as the data obtained with a supersonic Mach number are compared to the data obtained with a freestream Mach number of 0.8. Local and spanwise-averaged effectiveness magnitudes are generally higher when shock waves are present when a film plenum condition (with zero cross-flow Mach number) is utilized. Effectiveness values measured with a supersonic approaching freestream and shock waves then decrease as the injection cross-flow Mach number increases. Such changes are due to altered flow separation regions in film holes, different injection velocity distributions at hole exits, and alterations of static pressures at film hole exits produced by different types of shock wave events.

Design Method for Subsonic and Transonic Cascade With Prescribed Mach Number Distribution

1992 ◽  
Vol 114 (3) ◽  
pp. 553-560 ◽  
Author(s):  
O. Le´onard ◽  
R. A. Van den Braembussche
Keyword(s):  
Mach Number ◽  
Flow Field ◽  
Pressure Distribution ◽  
Velocity Component ◽  
Design Method ◽  
Normal Velocity ◽  
Flow Solver ◽  
Number Distribution ◽  
Mach Number Distribution ◽  
Transonic Cascade

A iterative procedure for blade design, using a time marching procedure to solve the unsteady Euler equations in the blade-to-blade plane, is presented. A flow solver, which performs the analysis of the flow field for a given geometry, is transformed into a design method. This is done by replacing the classical slip condition (no normal velocity component) by other boundary conditions, in such a way that the required pressure or Mach number distribution may be imposed directly on the blade. The unknowns are calculated on the blade wall using the so-called compatibility relations. Since the blade shape is not compatible with the required pressure distribution, a nonzero velocity component normal to the blade wall evolves from the new flow calculation. The blade geometry is then modified by resetting the wall parallel to the new flow field, using a transpiration technique, and the procedure is repeated until the calculated pressure distribution has converged to the required one. Examples for both subsonic and transonic flows are presented and show a rapid convergence to the geometry required for the desired Mach number distribution. An important advantage of the present method is the possibility to use the same code for the design and the analysis of a blade.

CFD Investigation of the Flow of Trailing Edge Cooling Slots

2018 ◽  
Author(s):  
Y. Jiang ◽  
N. Gurram ◽  
E. Romero ◽  
P. T. Ireland ◽  
L. di Mare
Keyword(s):  
Mach Number ◽  
Kinetic Energy ◽  
Flow Separation ◽  
Film Cooling ◽  
High Speed ◽  
Turbine Blades ◽  
Cross Flow ◽  
Turbulent Energy ◽  
Trailing Edge ◽  
Flow Interactions

Slot film cooling is a popular choice for trailing edge cooling in high pressure (HP) turbine blades because it can provide more uniform film coverage compared to discrete film cooling holes. The slot geometry consists of a cut back in the blade pressure side connected through rectangular openings to the internal coolant feed passage. The numerical simulation of this kind of film cooling flows is challenging due to the presence of flow interactions like step flow separation, coolant-mainstream mixing and heat transfer. The geometry under consideration is a cutback surface at the trailing edge of a constant cross-section aerofoil. The cutback surface is divided into three sections separated by narrow lands. The experiments are conducted in a high speed cascade in Oxford Osney Thermo-Fluids Laboratory at Reynolds and Mach number distributions representative of engine conditions. The capability of CFD methods to capture these flow phenomena is investigated in this paper. The isentropic Mach number and film effectiveness are compared between CFD and pressure sensitive paint (PSP) data. Compared to steady k–ω SST method, Scale Adaptive Simulation (SAS) can agree better with the measurement. Furthermore, the profiles of kinetic energy, production and shear stress obtained by the steady and SAS methods are compared to identify the main source of inaccuracy in RANS simulations. The SAS method is better to capture the unsteady coolant-hot gas mixing and vortex shedding at the slot lip. The cross flow is found to affect the film significantly as it triggers flow separation near the lands and reduces the effectiveness. The film is non-symmetric with respect to the half-span plane and different flow features are present in each slot. The effect of mass flow ratio (MFR) on flow pattern and coolant distribution is also studied. The profiles of velocity, kinetic energy and production of turbulent energy are compared among the slots in detail. The MFR not only affects the magnitude but also changes the sign of production.

scholarly journals Exploring the role of cosmological shock waves in the Dianoga simulations of galaxy clusters

2021 ◽  
Author(s):  
S Planelles ◽  
S Borgani ◽  
V Quilis ◽  
G Murante ◽  
V Biffi ◽  
...  
Keyword(s):  
Mach Number ◽  
Shock Waves ◽  
Galaxy Clusters ◽  
Cluster Mass ◽  
Mach Number Distribution ◽  
Disturbed Systems ◽  
Shock Mach Number ◽  
Formation And Evolution ◽  
Massive Clusters

Abstract Cosmological shock waves are ubiquitous to cosmic structure formation and evolution. As a consequence, they play a major role in the energy distribution and thermalization of the intergalactic medium (IGM). We analyse the Mach number distribution in the Dianoga simulations of galaxy clusters performed with the SPH code GADGET-3. The simulations include the effects of radiative cooling, star formation, metal enrichment, supernova and active galactic nuclei feedback. A grid-based shock-finding algorithm is applied in post-processing to the outputs of the simulations. This procedure allows us to explore in detail the distribution of shocked cells and their strengths as a function of cluster mass, redshift and baryonic physics. We also pay special attention to the connection between shock waves and the cool-core/non-cool core (CC/NCC) state and the global dynamical status of the simulated clusters. In terms of general shock statistics, we obtain a broad agreement with previous works, with weak (low-Mach number) shocks filling most of the volume and processing most of the total thermal energy flux. As a function of cluster mass, we find that massive clusters seem more efficient in thermalising the IGM and tend to show larger external accretion shocks than less massive systems. We do not find any relevant difference between CC and NCC clusters. However, we find a mild dependence of the radial distribution of the shock Mach number on the cluster dynamical state, with disturbed systems showing stronger shocks than regular ones throughout the cluster volume.

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