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Friday, July 10, 2020 | History

3 edition of Thermal analysis and design of a cooling system for a Mach 14 nozzle found in the catalog.

Thermal analysis and design of a cooling system for a Mach 14 nozzle

Thermal analysis and design of a cooling system for a Mach 14 nozzle

final report covering period February, 1986 to January, 1987

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  • 28 Currently reading

Published by Thermodynamics Facilities Branch, NASA Ames Research Center, National Technical Information Service, distributor in Moffett Field, CA, [Springfield, Va .
Written in English


Edition Notes

Statementprepared by Ronald Mullisen, principal investigator ; Keith Kaste
SeriesNASA contractor report -- NASA CR-180098
ContributionsKaste, Keith, Ames Research Center. Thermodynamics Facilities Branch
The Physical Object
FormatMicroform
Pagination1 v
ID Numbers
Open LibraryOL14985432M

Thermal Management in a Scramjet-Powered Hypersonic Cruise Vehicle by Christopher D. Marley A dissertation submitted in partial fulfillment of the requirements for the degree of. Thermal Design And Analysis Of Regeneratively Cooled Thrust Chamber Of Cryogenic Rocket Engine T. Vinitha1, S. Senthilkumar2, K. Manikandan3 PG scholar1, Research scholar 2,Research scholar3 1, 2,3Department of Aeronautical Engineering, Nehru institute of engineering and technology Abstract The thrust chamber of the cryogenic rocket engine.

  The influencing factors of the cooling system are studied to give guidance on the practical design of R.C./F.C. cooled scramjet engine with hydrocarbon fuel used as propellant. 2. Working scheme and physical model. The working process of the R.C./F.C. cooling system for the hydrocarbon fueled scramjet engine has been shown in Fig. 1. The fuel. Thermal Analysis | Thermal Modeling is a key part of any thermal design. These days, one can model almost anything using the state-of-the-art analysis and modeling tools available in the market. When used correctly, these tools can give you accurate results quickly and cost-effectively.

Adiabatic wall temperature, Mach number, friction factor Adiabatic, critical Velocity, underground car park Adsorption, Adsorbent, Natural Zeolite, Cooling Application Adsorption, Humic Substances, Batch Stirred Tank Reactor, Coconut Copra Aerodynamics, front cascade wing, down-force, drag-force, CFD analysis. Design of a Solar-Driven Ejector Cooling System S. du Clou and M.J. Brooks University of KwaZulu-Natal, Durban, , South Africa Centre for Renewable and Sustainable Energy StudiesAbstractThe Pulse Refrigeration System, or PRS, is a heat transfer and cooling system .


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Thermal analysis and design of a cooling system for a Mach 14 nozzle Download PDF EPUB FB2

Get this from a library. Thermal analysis and design of a cooling system for a Mach 14 nozzle: final report covering period February, to January, [Ronald Mullisen; Keith Kaste; Ames Research Center. Thermodynamics Facilities Branch.]. 1. Introduction. Cooling flow of a first stage high pressure turbine nozzle guide vane (HPNGV) may account for up to 13% of core flow in modern gas turbine engines [].This is of great importance regarding the fact that 1% of cooling air is a loss of approximately 1% specific fuel consumption [].Therefore efficient utilization and control of turbine cooling flows are essential for Author: Hadi Yavari, Ali Khavari, Mohammad Alizadeh, Behrad Kashfi, Hiwa Khaledi.

THERMAL ANALYSIS AND DESIGN OF A COOLING SYSTEM FOR THE NASA AMES MACH 14 WIND TUNNEL NOZZLE This report provides the analysis and design of a Mach 14 converging diverging nozzle wall liner. indicates that: no fin on the coolant side of the nozzle wall is optimum, the thermal stresses are dominant, and the critical area is very near the throat.

The analysis and design of a Mach 14 converging diverging nozzle wall liner is provided. The analysis indicates that: no fin on the coolant side of the nozzle wall is optimum, the thermal stresses are dominant, and the critical area is very near the : Keith Kaste and Ronald Mullisen.

Thermal design includes the method of cooling, selection of coolant, flow rate of coolant, selection of configuration and dimensions of the channels and material selection. In order to analyze the thrust chamber cooling system, it is necessary to understand both dimensions and shape of the cooling : T.

Vinitha, S. Senthilkumar, K. Manikandan. The cascade was tested at a high inlet turbulence intensity level (Tu 1 = 9%) and at a constant inlet Mach number of and nominal cooling condition.

Aero-thermal characterization of vane platform was obtained through 5-hole probe and end wall adiabatic film cooling. propulsion system design, analysis and simulation effort. The program aids in unifying the nozzle, chamber and injector portions of a rocket propulsion system design effort quickly and efficiently using a streamlined graphical user interface (GUI).

The program also allows for the. The present work aims to provide insights into the design of an effective cooling system for the nozzle and other components of the hypersonic long-run wind tunnel.

Due In this work, the two-dimensional, three-dimensional and axisymmetric Mach nozzles at an inlet total temperature of K and a total pressure of MPa were studied with.

An external cooling scheme including several rows of fan-shaped and cylindrical cooling holes has been designed. By testing different cooling flow rates at a NGV exit Reynolds number of E+06 and Mach number ofdetailed aerodynamic and heat transfer values were obtained destined to assess the design tools for film cooled platforms.

A cooling system for a hypersonic aircraft is disclosed which enables hypersonic flight using non-cryogenic fuels.

The cooling system positions a primary heat exchanger at an external location on the aircraft which remains relatively cool during hypersonic flight. A working fluid is passed through the primary heat exchanger to the hot parts of a supersonic combustion ram jet engine.

Stage 1 Blade Pitch-Line Mach Number Distributton. External Heat Transfer Coefficient. Stage 1 Blade Cooling System. Stage 1 Blade Tip Cap Cooling Design. Stage 1 Blade Flow Characteristics.

Stage 1 Blade Pitch-Line Temperature Distribution st Steady- State Takeoff. Stage 1 Blade Transient Thermal Analysis, Stage 1 Shroud - Cooling Geometry. The revealed trends would be useful to guide the design of thermal management systems for kerosene-fueled scramjets.

condition of Mach 7 with active cooling for a coolant mass flow rate of 0. help of fluid management system or suction system on this 16 spray nozzle array, the heat transfer was improved on the average by 30 W/cm 2 for similar values of superheat above 5 °C.

cooling system is the amount of adequate cooling. developed to design a rocket nozzle cooling sy stem and to predict the effectiveness of the perform thermal analysis on two engines: the. Recent research suggests that the cooling flow at transonic turbine tip has several unique flow features such as the strong interaction with the base flow, the acceleration at div.

Engines like SSME, F-1, J-2, RS, Vulcain 2, RD and RD use film cooling technique for combustion chamber cooling.

Several open-cycle rocket engines have turbine exhaust gas (TEG) delivered to the nozzle for film cooling, including the F1 engine and J2 engine of the United States, the upgraded LE5 engine of Japan and Vulcain 2 of the EADS Astrium.

Coolant-to-mainstream blowing ratio (BR), density ratio (DR), main flow isentropic exit Mach number (Ma 2 is), and turbulence intensity level (Tu 1) were the considered parameters. The cascade was tested in an atmospheric wind tunnel at Ma 2 is values ranging from towith an inlet turbulence intensity level of % and 9%, at variable.

Abstract. This paper will describe the thermal analysis techniques used to predict temperatures in the film-cooled ablative rocket nozzle used on the MC-1 60K rocket engine.

A model was developed that predicts char and pyrolysis depths, liner thermal gradients, and temperatures of the bondline between the overwrap and liner. Correlation of the. In the current study, a numerical analysis was performed for the heat transfer over the surface of nozzle guide vanes (NGVs) using three-dimensional computational fluid dynamics (CFD) models.

The investigation has taken place in two stages:. NASA/CR RAE Design, Fabrication, and Testing of an Auxiliary Cooling System for Jet Engines Kevin Leamy and Jim Griffiths GE Aircraft Engines, Cincinnati, Ohio. Heat transfer and pressure drop for a representative part of a turbine active cooling system were numerically investigated by means of an in-house code.

This code has been developed in the framework of an internal research program and has been validated by experiments and CFD. The analysed system represents the classical open bird cage arrangement that consists of an air supply pipe with a.A standalone solar PV system suitable to run the mist cooling system of PSI line pressure is designed.

The detailed design and selection of appropriate components are discussed in this article. The precooled combined cycle engines were proposed to overcome the limitation of Mach number due to high-temperature inlet. However, there has been li.