An experimental and computational investigation of film cooling effects on an interceptor forebody at Mach 10

1990 ◽  
Author(s):  
J. MAJESKI ◽  
H. MORRIS
Keyword(s):  
Film Cooling ◽  
Computational Investigation ◽  
Cooling Effects ◽  
Mach 10

Cross Flow Effects on Endwall Heat Transfer in Film-Cooled Turbine Nozzle Guide Vanes

2013 ◽  
Author(s):  
Rebecca Reviol ◽  
Roman Franze ◽  
Martin Böhle ◽  
Kenichiro Takeishi ◽  
Alexander Wiedermann
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Cross Flow ◽  
Turbulence Models ◽  
Secondary Flows ◽  
Stagnation Region ◽  
Naphthalene Sublimation ◽  
Sublimation Method ◽  
Cooling Effects ◽  
Nozzle Guide

Film cooling effects on endwalls in the stagnation point region are of special interest since the heat transfer is influenced drastically by secondary flows. Additionally, a complex vortex structure exists along the stagnation streamline which influences heat transfer on the endwall. The flow phenomenon is described and discussed in the open literature but it is still difficult to predict the heat transfer on the endwall and the turbine profile by CFD methods with sufficient accuracy. In this paper it is examined how the flow field in the stagnation region should be simulated using CFD. The effect of meshes with various grid resolutions and turbulence models as k-ε-, k-ω-SST- and DES-turbulence models have been investigated. The CFD-data are compared with the experimental results obtained by Naphthalene Sublimation Method, Pressure Sensitive Paint, Laser Induced Fluorescence and Particle Image Velocimetry. Three cases, namely film cooling on a flat plate, the endwall flow near a symmetrical airfoil and the symmetrical airfoil with endwall film cooling, are examined in detail.

Experimental/Computational Investigation of Drag Reduction by Electric-Arc Airspikes at Mach 10

2004 ◽  
Author(s):  
Christopher Hartley ◽  
Tom Portwood ◽  
Mathew Filippelli ◽  
Leik Myrabo ◽  
Henry Nagamatsu ◽  
...  
Keyword(s):  
Drag Reduction ◽  
Electric Arc ◽  
Computational Investigation ◽  
Mach 10

The Effects of Vane Showerhead Injection Angle and Film Compound Angle on Nozzle Endwall Cooling (Phantom Cooling)

2014 ◽  
Vol 137 (2) ◽  
Author(s):  
Luzeng Zhang ◽  
Juan Yin ◽  
Hee Koo Moon
Keyword(s):  
Film Cooling ◽  
Suction Side ◽  
Nitrogen Gas ◽  
Injection Angle ◽  
Test Models ◽  
Cooling Hole ◽  
Cooling Effects ◽  
Film Cooling Holes ◽  
Compound Angle ◽  
Endwall Cooling

The effects of airfoil showerhead (SH) injection angle and film-cooling hole compound angle on nozzle endwall cooling (second order film-cooling effects, also called "phantom cooling") were experimentally investigated in a scaled linear cascade. The test cascade was built based on a typical industrial gas turbine nozzle vane. Endwall surface phantom cooling film effectiveness measurements were made using a computerized pressure sensitive paint (PSP) technique. Nitrogen gas was used to simulate cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness can be obtained by the mass transfer analogy. Two separate nozzle test models were fabricated, which have the same number and size of film-cooling holes but different configurations. One had a SH angle of 45 deg and no compound angles on the pressure and suction side (SS) film holes. The other had a 30 deg SH angle and 30 deg compound angles on the pressure and SS film-cooling holes. Nitrogen gas (cooling air) was fed through nozzle vanes, and measurements were conducted on the endwall surface between the two airfoils where no direct film cooling was applied. Six cooling mass flow ratios (MFRs, blowing ratios) were studied, and local (phantom) film effectiveness distributions were measured. Film effectiveness distributions were pitchwise averaged for comparison. Phantom cooling on the endwall by the SS film injections was found to be insignificant, but phantom cooling on the endwall by the pressure side (PS) airfoil film injections noticeably helped the endwall cooling (phantom cooling) and was a strong function of the MFR. It was concluded that reducing the SH angle and introducing a compound angle on the PS injections would enhance the endwall surface phantom cooling, particularly for a higher MFR.

Flow Migration and Mixing in a High Pressure Turbine Rotor Passage With Axisymmetrically Contoured Endwall and Leakage Flow

2015 ◽  
Author(s):  
Reema Saxena ◽  
Arya Ayaskanta ◽  
Terrence W. Simon ◽  
Hee-Koo Moon ◽  
Luzeng J. Zhang
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Film Cooling ◽  
Thermal Loading ◽  
Turbine Rotor ◽  
Secondary Flows ◽  
Leakage Flow ◽  
Flow Rates ◽  
Flow Mixing ◽  
Cooling Effects

The flow field in the passage of a high pressure gas turbine is quite complex, involving strong secondary flows, transverse pressure gradients and strong streamwise acceleration. This complexity may have an adverse effect on cooling of the hub endwall, which is subjected to high thermal loading due to the flat combustor exit temperature profile of modern low-NOx systems. Therefore, given material limitations, better cooling management techniques that can be included with certainty in new gas turbine designs are needed. In the present study, film cooling has been investigated experimentally in a stationary linear cascade. The flow is representative of a high pressure gas turbine rotor with combustor liner coolant introduced to the approach flow. Focus is on the endwall axisymmetric contouring and the cooling effect of leakage flow bled from the compressor through the stator-rotor disc cavity. Two endwall contours, ‘shark nose’ (gradual slope over a larger distance) and ‘dolphin nose’ (steep slope over a shorter distance), are considered and comparison is made under conditions of three mass flow rates (MFR) of leakage, 0.5%, 1.0% and 1.5% of the approach flow rate. The performance of both endwall contours is compared at different streamwise locations in terms of adiabatic effectiveness values over the endwall. This study gives enhanced insight into the physics of coolant flow mixing, migration and subsequent coverage over the endwall. The results show the cooling effects of the contoured shapes over a range of leakage flow rates in the strong secondary flow environment. It is found that the leakage flow plays a crucial role in enhancing coolant coverage over the endwall. To add to our knowledge of mixing effects, detailed thermal field data are taken in the leakage flow discharge region. Doing so helps explain the behavior of the flow as it is ejected into the passage and interacts with the mainstream flow.

PHANTOM COOLING EFFECTS ON ROTOR BLADE SURFACE HEAT FLUX IN A TRANSONIC FULL SCALE 1+1/2 STAGE ROTATING TURBINE

2021 ◽  
pp. 1-13
Author(s):  
Richard J. Anthony ◽  
John Finnegan ◽  
John Clark
Keyword(s):  
Heat Flux ◽  
Air Force ◽  
Film Cooling ◽  
Surface Heat Flux ◽  
Rotor Blade ◽  
Surface Heat ◽  
Pressure Side ◽  
Blade Surface ◽  
Leakage Flows ◽  
Cooling Effects

Abstract An experimental and numerical investigation of phantom cooling effects on cooled and uncooled rotating high pressure turbine blades in a full scale 1+1/2 stage turbine test is carried out. Objectives set to capture, separate, and quantify the effects of upstream vane film-cooling and leakage flows on the downstream rotor blade surface heat flux. Multiple series of tests were carried out in the Air Force Research Laboratory, Turbine Research Facility, at Wright-Patterson Air Force Base, Ohio. A non-proprietary research turbine test article is uniquely instrumented with high frequency double-sided thin film heat flux gauges custom made at AFRL. High bandwidth, time resolved surface heat flux is measured on multiple film-cooled and non-film-cooled HPT rotor blades downstream of both film-cooled and non-film-cooled vane sectors. Upstream wake passing and heat flux is characterized on both rotor pressure and suction side surfaces, along with quantifying rotor phantom cooling effects from non-uniform 1st stage vane film cooling and leakage flows. Fast response heat flux measurements quantify how rotor phantom cooling impacts the blade pressure side greatest; increasing along the pressure side towards the trailing edge. It is discovered upstream vane film-cooling alone can account for 50% of the rotor blade cooling effect, and even outweigh the rotor blade film cooling effect far from the blade showerhead holes. Added unsteady numerical simulation demonstrates how variations in inlet total temperature and incidence angle can also contribute to circumferentially non-uniform rotor heat flux.

Implementation of Vortex Generator and Ramp to Improve Film Cooling Effectiveness on Blade Endwall

2021 ◽  
Author(s):  
Sadam Hussain ◽  
Xin Yan
Keyword(s):  
Numerical Methods ◽  
Film Cooling ◽  
Vortex Generator ◽  
Leading Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Baseline Case ◽  
Cooling Hole ◽  
Cooling Effects ◽  
Endwall Film Cooling

Abstract With the arrangements of vortex generators (VG) and ramp, film cooling effects on endwall near leading edge were numerically investigated at two blowing ratios (i.e. M = 0.5 and M = 1). To determine suitable numerical methods, mesh independency analysis and turbulence model selection were carried out based on the existing experimental data and LES results. With the numerical methods, flow fields near the leading edge were visualized to illustrate the influence of VG and ramp on coolant coverage on blade endwall. Film cooling effectiveness distributions on endwall and coolant trajectories near leading edge were compared among five different configurations with VG and ramp. The results show that the attachment of coolant on blade endwall is improved with the implement of VG between shaped-hole and leading edge. With the implementation of ramp on endwall between cooling hole and leading edge, the coolant spreads wider on endwall along pitchwise direction than the baseline case. With the implementation of VG and ramp, film cooling effect on endwall near leading edge is significantly improved as compared with the only ramp and only VG cases. Compared with the baseline case, pitchwise-averaged film cooling effectiveness on blade endwall near leading edge is increased by about 9%, and the film cooling effectiveness distributions on endwall along pitchwise direction become much uniform, for the case with both ramp and VG at M = 1.

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