Experimental Investigations of Film Cooling in a Conical Nozzle Under Rocket-Engine-Like Flow Conditions

AIAA Journal ◽  
2019 ◽  
Vol 57 (3) ◽  
pp. 1172-1183 ◽  
Author(s):  
S. Ludescher ◽  
H. Olivier
Keyword(s):  
Film Cooling ◽  
Rocket Engine ◽  
Experimental Investigations ◽  
Conical Nozzle ◽  
Flow Conditions

scholarly journals Numerical Simulation of Reactive Flows in Overexpanded Supersonic Nozzle with Film Cooling

2015 ◽  
Vol 2015 ◽  
pp. 1-15 ◽  
Author(s):  
Mohamed Sellam ◽  
Amer Chpoun
Keyword(s):  
Film Cooling ◽  
Structural Integrity ◽  
Safety Issue ◽  
Rocket Engine ◽  
Supersonic Nozzle ◽  
Nozzle Geometry ◽  
Nozzle Flow ◽  
Rocket Engines ◽  
Flow Conditions ◽  
Reactive Flows

Reignition phenomena occurring in a supersonic nozzle flow may present a crucial safety issue for rocket propulsion systems. These phenomena concern mainly rocket engines which use H2gas (GH2) in the film cooling device, particularly when the nozzle operates under over expanded flow conditions at sea level or at low altitudes. Consequently, the induced wall thermal loads can lead to the nozzle geometry alteration, which in turn, leads to the appearance of strong side loads that may be detrimental to the rocket engine structural integrity. It is therefore necessary to understand both aerodynamic and chemical mechanisms that are at the origin of these processes. This paper is a numerical contribution which reports results from CFD analysis carried out for supersonic reactive flows in a planar nozzle cooled with GH2film. Like the experimental observations, CFD simulations showed their ability to highlight these phenomena for the same nozzle flow conditions. Induced thermal load are also analyzed in terms of cooling efficiency and the results already give an idea on their magnitude. It was also shown that slightly increasing the film injection pressure can avoid the reignition phenomena by moving the separation shock towards the nozzle exit section.

Boundary-layer velocity profiles in a swirling convergent flow field

1972 ◽  
Vol 52 (2) ◽  
pp. 357-367 ◽  
Author(s):  
T. M. Houlihan ◽  
D. J. Hornstra
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Swirling Flow ◽  
Theoretical Calculations ◽  
Experimental Investigations ◽  
Conical Nozzle ◽  
Flow Conditions ◽  
Boundary Layer Flows ◽  
Layer Velocity ◽  
Convergent Flow

Velocity distributions within the boundary layer of a swirling flow of incompressible fluid in a convergent conical nozzle have been investigated. Theoretical calculations with boundary conditions more appropriate to physically existent situations discounted the existence of 'super-velocities’ within the boundary layer. Parallel experimental investigations demonstrated an interdependence of core and boundary-layer flows which precluded the maintenance of the flow conditions required by the analysis.

scholarly journals Film Cooling in Rocket Nozzles

2020 ◽  
pp. 65-78
Author(s):  
Sandra Ludescher ◽  
Herbert Olivier
Keyword(s):  
Experimental Data ◽  
Parametric Study ◽  
Film Cooling ◽  
Rocket Engine ◽  
Cooling Efficiency ◽  
Conical Nozzle ◽  
Hot Gas ◽  
Dual Bell ◽  
Film Cooling Efficiency ◽  
Cooling Model

Abstract In this project supersonic, tangential film cooling in the expansion part of a nozzle with rocket-engine like hot gas conditions was investigated. Therefore, a parametric study in a conical nozzle was conducted revealing the most important influencing parameter on film cooling for the presented setup. Additionally, a new axisymmetric film cooling model and a method for calculating the cooling efficiency from experimental data was developed. These models lead to a satisfying correlation of the data. Furthermore, film cooling in a dual-bell nozzle performing in altitude mode was investigated. The aim of these experiments was to show the influence of different contour inflection geometries on the film cooling efficiency in the bell extension.

Experimental Investigation of Film-Cooling Effectiveness of a Highly Loaded Turbine Blade Under Steady and Periodic Unsteady Flow Conditions

2017 ◽  
Vol 139 (7) ◽  
Author(s):  
Ali Nikparto ◽  
Meinhard T. Schobeiri
Keyword(s):  
Unsteady Flow ◽  
Film Cooling ◽  
Flow Condition ◽  
Leading Edge ◽  
Experimental Investigations ◽  
Pressure Side ◽  
Flow Conditions ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Highly Loaded

This paper describes the experimental investigations of film-cooling effectiveness on a highly loaded low-pressure turbine blade under steady and unsteady wake flow conditions. The cascade facility in Turbomachinery Performance and Flow Research Lab (TPFL) at the Texas A&M University was used to simulate the periodic flow condition inside gas turbine engines. Moving wakes, originated from upstream stator blades, are simulated inside the cascade facility by moving rods in front of the blades. The flow coefficient is maintained at 0.8 and the incoming wakes have a reduced frequency of 3.18. A total of 617 holes on the blade are distributed along 13 different rows. Six rows cover the suction side, six other rows cover the pressure side, and one last row feeds the leading edge. Each row has a twin row on the other side of the blade with exact same number of holes and arrangement (except for leading edge). They both are connected to the same cavity. Coolant is injected from either sides of the blade through cavities to form a uniform distribution along the span of the blade. Film-cooling effectiveness under periodic unsteady flow condition was studied using pressure-sensitive paint. Experiments were performed at Reynolds number of 150,000 and blowing ratio of one, based on equal mass flux distribution. Experimental investigations were performed to determine the effect of flow separation and pressure gradient on film-cooling effectiveness. Moreover, the effect of impinging wakes on the overall film coverage of blade surfaces was studied. It was found that heat transfer coefficient (HTC) and film-cooling effectiveness (FCE) in majority of regions behave in opposite ways. This can be justified from turbulence intensity and velocity fluctuation point of view. Also, unsteady wakes imposed on top of film injection have opposite effects on suction and pressure side of the blade. This is more clearly seen in region near leading edge.

A study into the feasibility of using the oxygen-hydrocarbon engine 11D58M as a basis for development of a high-performance multifunctional gas-generatorless rocket engine with oxygen cooling

2019 ◽  
pp. 67-80
Author(s):  
Boris A. Sokolov ◽  
Nikolay N. Tupitsyn
Keyword(s):  
Film Cooling ◽  
High Performance ◽  
Rocket Engine ◽  
Hydrocarbon Fuel ◽  
Gas Generator ◽  
Automatic Equipment ◽  
The Us ◽  
Upper Stage ◽  
Pump Assembly ◽  
Turbo Pump

The paper presents results of engineering studies and research and development efforts at RSC Energia to analyze and prove the feasibility of using the mass-produced oxygen-hydrocarbon engine 11D58M with 8.5 ton-force thrust as a basis for development of a high-performance multifunctional rocket engine with oxygen cooling and 5 ton-force thrust, which is optimal for upper stages (US), embodying a system that does not include a gas generator. The multi-functionality of the engine implies including in it additional units supporting some functions that are important for US, such as feeding propellant from US tanks to the engine after flying in zero gravity, autonomous control of the engine automatic equipment to support its firing, shutdown, adjustments during burn and emergency protection in case of off-nominal operation, as well as generating torques for controlling the US attitude and stabilizing it during coasting, etc. Replacing conventional engine chamber cooling that uses high-boiling hydrocarbon fuel with the innovative oxygen cooling makes it possible to get rid of the internal film cooling circuits and eliminate their attendant losses of fuel, while the use of the oxygen gasified in the cooling circuit of the chamber to drive the turbo pump assembly permits to design an engine that does not have a gas generator. Key words: Multifunctional rocket engine, oxygen cooling, gas-generatorless design, upper stage.

Effusion-Cooling Performance at Gas Turbine Combustor Representative Flow Conditions

2012 ◽  
Author(s):  
Vinod U. Kakade ◽  
Steven J. Thorpe ◽  
Miklós Gerendás
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Free Stream ◽  
Flow Conditions ◽  
Cooling Performance ◽  
Free Stream Turbulence ◽  
High Turbulence ◽  
Adiabatic Effectiveness

The thermal management of aero gas turbine engine combustion systems commonly employs effusion-cooling in combination with various cold-side convective cooling schemes. The combustor liner incorporates many small holes which are usually set in staggered arrays and at a shallow angle to the cooled surface; relatively cold compressor delivery air is then allowed to flow through these holes to provide the full-coverage film-cooling effect. The efficient design of such systems requires robust correlations of film-cooling effectiveness and heat transfer coefficient at a range of aero-thermal conditions, and the use of appropriately validated computational models. However, the flow conditions within a combustor are characterised by particularly high turbulence levels and relatively large length scales. The experimental evidence for performance of effusion-cooling under such flow conditions is currently sparse. The work reported here is aimed at quantifying typical effusion-cooling performance at a range of combustor relevant free-stream conditions (high turbulence), and also to assess the importance of modeling the coolant to free-stream density ratio. Details of a new laboratory wind-tunnel facility for the investigation of film-cooling at high turbulence levels are reported. For a typical combustor effusion geometry that uses cylindrical holes, spatially resolved measurements of adiabatic effectiveness, heat transfer coefficient and net heat flux reduction are presented for a range of blowing ratios (0.48 to 2), free-stream turbulence conditions (4 and 22%) and density ratios (0.97 and 1.47). The measurements reveal that elevated free-stream turbulence impacts on both the adiabatic effectiveness and heat transfer coefficient, although this is dependent upon the blowing ratio being employed and particularly the extent to which the coolant jets detach from the surface. At low blowing ratios the presence of high turbulence levels causes increased lateral spreading of the coolant adjacent to the injection points, but more rapid degradation in the downstream direction. At high blowing ratios, high turbulence levels cause a modest increase in effectiveness due to turbulent transport of the detached coolant fluid. Additionally, the augmentation of heat transfer coefficient caused by the coolant injection is seen to be increased at high free-stream turbulence levels.

Cooling Effectiveness on Film Cooled Turbine Blades

2001 ◽  
Author(s):  
M. Derrar ◽  
J. Nagler ◽  
W. W. Koschel
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Separation Bubble ◽  
Turbine Blades ◽  
Suction Side ◽  
Experimental Simulation ◽  
Flow Conditions ◽  
Cooling Effectiveness ◽  
Boundary Interaction ◽  
Simultaneous Modeling

Abstract This paper presents experiments on the cooling effectiveness obtained for two different injection locations on the suction side of a turbine blade at transonic flow conditions. Previous results of a computational analysis and flow visualization indicated that a separation bubble is present on the suction side at a location x/L = 0.43 and the location x/L = 0.575 corresponds to a shock-boundary interaction zone [9]. The scientific interest is primarily focused on the realization of high film cooling efficiencies and its relevant parameters under these flow conditions. Streamwise aligned as well as inclined angled film coolant hole configurations have been investigated for each location. Due to the high number of interacting parameters the experimental simulation of turbine blade film cooling is extremely complex, which can only be solved by a simultaneous modeling using the experimentally measured results. Test rig, instrumentation and data analysis are described in detail. The goal of the investigations is to determine the optimum location of the film coolant injection.

The Effects of Inlet Boundary Layer Condition and Periodically Incoming Wakes on Secondary Flow in a Low Pressure Turbine Cascade

2020 ◽  
Author(s):  
Tobias Schubert ◽  
Silvio Chemnitz ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Secondary Flow ◽  
High Speed ◽  
Low Pressure ◽  
Experimental Investigations ◽  
Turbine Cascade ◽  
Flow Conditions ◽  
Low Pressure Turbine ◽  
Speed Flow ◽  
Unsteady Inflow

Abstract A particular turbine cascade design is presented with the goal of providing a basis for high quality investigations of endwall flow at high-speed flow conditions and unsteady inflow. The key feature of the design is an integrated two-part flat plate serving as a cascade endwall at part-span, which enables a variation of the inlet endwall boundary layer conditions. The new design is applied to the T106A low pressure turbine cascade for endwall flow investigations in the High-Speed Cascade Wind Tunnel of the Institute of Jet Propulsion at the Bundeswehr University Munich. Measurements are conducted at realistic flow conditions (M2th = 0.59, Re2th = 2·105) in three cases of different endwall boundary layer conditions with and without periodically incoming wakes. The endwall boundary layer is characterized by 1D-CTA measurements upstream of the blade passage. Secondary flow is evaluated by Five-hole-probe measurements in the turbine exit flow. A strong similarity is found between the time-averaged effects of unsteady inflow conditions and the effects of changing inlet endwall boundary layer conditions regarding the attenuation of secondary flow. Furthermore, the experimental investigations show, that all design goals for the improved T106A cascade are met.

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.

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