Film Cooling Effects on Wall Heat Flux of a Liquid Propellant Combustion Chamber

2006 ◽  
Author(s):  
Jong-Gyu Kim ◽  
Kwang-Jin Lee ◽  
Seonghyeon Seo ◽  
Yeoung-Min Han ◽  
Hong-Jip Kim ◽  
...  
Keyword(s):  
Heat Flux ◽  
Combustion Chamber ◽  
Film Cooling ◽  
Wall Heat Flux ◽  
Liquid Propellant ◽  
Propellant Combustion ◽  
Cooling Effects

scholarly journals Large-Eddy Simulations for the Wall Heat Flux Prediction of a Film-Cooled Single-Element Combustion Chamber

2020 ◽  
pp. 223-234
Author(s):  
R. Olmeda ◽  
P. Breda ◽  
C. Stemmer ◽  
M. Pfitzner
Keyword(s):  
Heat Flux ◽  
Combustion Chamber ◽  
Film Cooling ◽  
Structural Integrity ◽  
Cooling System ◽  
Rocket Engine ◽  
Large Eddy Simulations ◽  
Wall Heat Flux ◽  
Single Element ◽  
Large Eddy

Abstract In order for modern launcher engines to work at their optimum, film cooling can be used to preserve the structural integrity of the combustion chamber. The analysis of this cooling system by means of CFD is complex due to the extreme physical conditions and effects like turbulent fluctuations damping and recombination processes in the boundary layer which locally change the transport properties of the fluid. The combustion phenomena are modeled by means of Flamelet tables taking into account the enthalpy loss in the proximity of the chamber walls. In this work, Large-Eddy Simulations of a single-element combustion chamber experimentally investigated at the Technical University of Munich are carried out at cooled and non-cooled conditions. Compared with the experiment, the LES shows improved results with respect to RANS simulations published. The influence of wall roughness on the wall heat flux is also studied, as it plays an important role for the lifespan of a rocket engine combustors.

Stability rating tests for the length-optimization of baffles in a liquid propellant combustion chamber using a pulse gun

2008 ◽  
Vol 12 (3) ◽  
pp. 214-222 ◽  
Author(s):  
Hong Jip Kim ◽  
Seonghyeon Seo ◽  
Kwang Jin Lee ◽  
Yeoung Min Han ◽  
Soo Yong Lee ◽  
...  
Keyword(s):  
Combustion Chamber ◽  
Liquid Propellant ◽  
Propellant Combustion ◽  
Stability Rating ◽  
Length Optimization

scholarly journals Influence of Polyisobutylene Kerosene Additive on Combustion Efficiency in a Liquid Propellant Rocket Engine

Aerospace ◽  
2019 ◽  
Vol 6 (12) ◽  
pp. 129 ◽  
Author(s):  
Igor Borovik ◽  
Evgeniy Strokach ◽  
Alexander Kozlov ◽  
Valeriy Gaponov ◽  
Vladimir Chvanov ◽  
...  
Keyword(s):  
Combustion Chamber ◽  
Film Cooling ◽  
Rocket Engine ◽  
Combustion Efficiency ◽  
Liquid Propellant ◽  
Gaseous Oxygen ◽  
Wall Film ◽  
Measurement Results ◽  
Liquid Propellant Rocket Engine ◽  
Propellant Rocket

The combustion of kerosene with the polymer additive polyisobutylene (PIB) was experimentally investigated. The aim of the study was to measure the effect of PIB kerosene on the efficiency of combustion chamber cooling and the combustion efficiency of the liquid propellant for a rocket engine operating on kerosene and gaseous oxygen (GOX). The study was conducted on an experimental rocket engine using kerosene wall film cooling in the combustion chamber. Fire tests showed that the addition of polyisobutylene to kerosene had no significant effect on the combustion efficiency. However, analysis of the wall temperature measurement results showed that the use of PIB kerosene is more effective for film cooling than pure kerosene, which can increase the efficiency of combustion chamber cooling and subsequently increase its reliability and reusability. Thus, the findings of this study are expected to be of use in further investigations of wall film cooling efficiency.

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.

Impact of Radiation on the Wall Heat Load at a Test Bench Gas Turbine Combustion Chamber: Measurements and CFD Simulation

2007 ◽  
Author(s):  
R. Dannecker ◽  
K.-U. Schildmacher ◽  
B. Noll ◽  
R. Koch ◽  
M. Hase ◽  
...  
Keyword(s):  
Heat Flux ◽  
Combustion Chamber ◽  
Gas Turbine ◽  
Heat Load ◽  
Thermal Design ◽  
Wall Heat Flux ◽  
Radiative Properties ◽  
Numerical Work ◽  
Gas Turbine Combustion ◽  
The Impact

Experimental and numerical work has been carried out to determine the wall heat load at the liner structure of a model gas turbine combustion chamber. Measured cross-sectional profiles of the velocity and temperature field inside the chamber could be used to validate various CFD calculations of the combustion flow. It turned out that only a special treatment of the thermal boundary conditions at all liner walls would actually lead to appropriate values of the wall heat flux. Radiation modeling included two radiative properties models (SG single gray gas and WSSG weighted sum of gray gases) and three radiation transport models (P1, DT discrete transfer, MC Monte Carlo). The performance of the WSGG model has been assessed with charts and the impact of the radiation on the liner wall temperature distribution has been studied. The experimental values are matched within 3% deviation with the best combination of transport and radiation property models. The radiation contributes to 20-30% of the total wall heat flux. The present approach enables Siemens PG to access the thermal design of combustors more precisely.

Film Cooling and Heat Transfer in Nozzles

1988 ◽  
Vol 110 (1) ◽  
pp. 57-65 ◽  
Author(s):  
J. Stoll ◽  
J. Straub
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Film Cooling ◽  
Wall Heat Flux ◽  
Turbulent Heat Transfer ◽  
Turbulent Heat ◽  
Film Stability ◽  
Hot Air ◽  
Theoretical Investigations ◽  
Favorable Pressure Gradient

In this paper experimental and theoretical investigations on heat transfer and cooling film stability in a convergent–divergent nozzle are presented. Compressed air is injected into hot air in the inlet region of the nozzle and the influence of the strong favorable pressure gradient in the nozzle on turbulent heat transfer and mixing is examined. The experiments cover measurements of wall pressures, wall temperature, and wall heat flux. Calculations with parabolic finite difference boundary layer code have been performed using a well-known k–ε-turbulence model with an extension paying regard to acceleration. As a result the calculated wall heat flux is compared with the measured heat flux.

Phantom Cooling Effects on Rotor Blade Surface Heat Flux in a Transonic Full Scale 1+1/2 Stage Rotating Turbine

2020 ◽  
Author(s):  
Richard J. Anthony ◽  
John Finnegan ◽  
John P. Clark
Keyword(s):  
Heat Flux ◽  
High Pressure ◽  
Air Force ◽  
Film Cooling ◽  
Surface Heat Flux ◽  
Rotor Blade ◽  
Surface Heat ◽  
High Pressure Turbine ◽  
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 1+1/2 stage rotating high pressure turbine 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 nonuniform 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 aero numerical simulation demonstrate how variations in inlet total temperature and incidence angle can also contribute to circumferentially non-uniform rotor heat flux. Better understanding from this investigation aids modelling and design efforts in optimizing film cooling performance in real high pressure turbine flow fields. Understanding the behavior of such non-uniform circumferential rotor phantom cooling effects can be critical to optimize the efficiency, fuel consumption, range, and durability of advanced turbomachines.

Detailed Analysis of the Thermal Wall Heat Load in Annular Combustors

1999 ◽  
Author(s):  
Werner Krebs ◽  
Günther Walz ◽  
Stefan Hoffmann ◽  
Hans Judith
Keyword(s):  
Heat Flux ◽  
Combustion Chamber ◽  
Heat Load ◽  
Thermal Power ◽  
Wall Heat Flux ◽  
Navier Stokes ◽  
Ceramic Tiles ◽  
Combustion Gases ◽  
Annular Combustor ◽  
Cooling Air

A detailed thermal analysis involving both measurements and calculations has been carried out in order to determine the wall heal load and to optimize the amount of cooling air for an annular combustor. In calculations, the convective wall heat flux has been detemined by application of a 3D Navier-Stokes Code. Furthermore, the radiation exchange between the hot combustion gases and the liner has been calculated using a multidimensional spectral approach. Although a quite high thermal power density is found within the combustion chamber the wall heat load is at a low level. Values are well below 80 kW/m2, due to the application of ceramic tiles which have a low thermal conductivity. The wall heat load is dominated by radiation emitted in the lower gas radiation bands (λ < 2.9 μm). The convective wall heat flux is nearly balanced out by the sealing air which is discharged through gaps between the ceramic tiles. The cooling effect of the sealing air, however, is strongly influenced by the 3D near wall flow field in the combustion chamber.

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