Methane Transcritical Behavior in the Cooling System of the HYPROB-BREAD LOX/LCH4 Demonstrator Rocket Engine

2015 ◽  
Author(s):  
D. Ricci ◽  
F. Battista ◽  
M. Ferraiuolo ◽  
V. Salvatore ◽  
M. Fragiacomo
Keyword(s):  
Cooling System ◽  
Fluid Dynamic ◽  
Rocket Engine ◽  
Liquid Oxygen ◽  
Research Centre ◽  
Channel Height ◽  
Rocket Engines ◽  
Combustion Gases ◽  
Cooling Jacket ◽  
Cfd Analyses

The HYPROB Program, developed by the Italian Aerospace Research Centre, has the aim of increasing the Italian system design and manufacturing capabilities on liquid oxygen-hydrocarbon rocket engines; the most important activity is represented by the development and testing of a ground engine demonstrator of three tons thrust based on methane as propellant. The demonstrator baseline concept is featured by 18 injectors and is regeneratively cooled by using liquid methane. The cooling system has a counter-flow architecture and is made by 96 axial channels; methane enters the channels in the nozzle region in supercritical liquid condition, is heated by the combustion gases along the cooling jacket and then is injected into the combustion chamber as a supercritical gas. The goal of the present paper is to describe the activities supporting the cooling jacket design, aiming at identifying the optimal configuration of the cooling channels. 3-D CFD analyses have been performed on different cooling channel arrangements, in terms of channel height and rib width. Moreover, simulations described the thermo-fluid dynamic behavior of methane by means of NIST real gas modeling and they were necessary to give the proper input to the thermo-structural analyses in order to verify the most critical sections of the cooling jacket.

Thermo-Structural and Thermo-Fluid Dynamics Analyses Supporting the Design of the Cooling System of a Methane Liquid Rocket Engine

2014 ◽  
Author(s):  
M. Ferraiuolo ◽  
D. Ricci ◽  
F. Battista ◽  
P. Roncioni ◽  
V. Salvatore
Keyword(s):  
Single Channel ◽  
Cooling System ◽  
Fluid Dynamic ◽  
Thermal Analyses ◽  
Liquid Oxygen ◽  
Research Centre ◽  
Plastic Behavior ◽  
Rocket Engines ◽  
Cooling Jacket ◽  
D Model

The HYPROB Program, developed by the Italian Aerospace Research Centre, has the aim to increase the system design and manufacturing capabilities on liquid oxygen-methane rocket engines. It foresees the designing, manufacturing and testing of a ground engine demonstrator of three tons thrust. The demonstrator baseline concept is featured by 18 injectors and it is regeneratively cooled by using liquid methane. In particular, the cooling system is made by a constant number of axial channels and the counter-flow architecture has been chosen; methane enters the channels in the nozzle region in supercritical liquid condition, is heated by the combustion gases along the cooling jacket and then injected into the combustion chamber as a supercritical gas by means of the injection head. The goal of this paper is to describe the thermo-structural and the thermo-fluid dynamic analyses that have been performed in order to support the design activities aiming at identifying the optimal configuration of the cooling jacket in terms of number of channels, rib height and width. In fact, a fully 3-D model, regarding a single channel, heated by the design input heat flux has been considered in order to perform CFD simulations aiming at describing the thermo-fluid dynamic behavior of methane. The results in terms of convective heat transfer coefficients have been taken into account as inputs for the thermo-structural simulations on the most critical sections of the cooling jacket. The thermo-structural activity has been conducted on the demonstrator by means of a Finite Element Method code taking into account the visco-plastic behavior of the adopted materials. In particular, transient thermal analyses and static structural analyses have been performed using ANSYS code on a 2-D model. These analyses have demonstrated that the cooling jacket can withstand the design goal of 5 thermo-mechanical cycles with a safety factor equal to 4 considering a firing time equal to 30 seconds.

scholarly journals Numerical Investigation on the Thermal Behaviour of a LOx/LCH4 Demonstrator Cooling System

Aerospace ◽  
2021 ◽  
Vol 8 (6) ◽  
pp. 151
Author(s):  
Daniele Ricci ◽  
Francesco Battista ◽  
Manrico Fragiacomo
Keyword(s):  
Thermal Behaviour ◽  
Cooling System ◽  
Thermal Analyses ◽  
Operating Conditions ◽  
Liquid Oxygen ◽  
Rocket Engines ◽  
Successful Operation ◽  
Combustion Gases ◽  
Cooling Jacket ◽  
Liquid Rocket

Reliability of liquid rocket engines is strictly connected with the successful operation of cooling jackets, able to sustain the impressive operative conditions in terms of huge thermal and mechanical loads, generated in thrust chambers. Cryogenic fuels, like methane or hydrogen, are often used as coolants and they may behave as transcritical fluids flowing in the jackets: after injection in a liquid state, a phase pseudo-change occurs along the chamber because of the heat released by combustion gases and coolants exiting as a vapour. Thus, in the development of such subsystems, important issues are focused on numerical methodologies adopted to simulate the fluid thermal behaviour inside the jackets, design procedures as well as manufacturing and technological process topics. The present paper includes the numerical thermal analyses regarding the cooling jacket belonging to the liquid oxygen/liquid methane demonstrator, realized in the framework of the HYPROB (HYdrocarbon PROpulsion test Bench) program. Numerical results considering the nominal operating conditions of cooling jackets in the methane-fuelled mode and the water-fed one are included in the case of the application of electrodeposition process for manufacturing. A comparison with a similar cooling jacket, realized through the conventional brazing process, is addressed to underline the benefits of the application of electrodeposition technology.

Experimental Investigation on the Transcritical Behaviour of Methane and Numerical Rebuilding Activity in the Frame of the HYPROB-BREAD Project

2015 ◽  
Author(s):  
D. Ricci ◽  
P. Natale ◽  
F. Battista ◽  
V. Salvatore
Keyword(s):  
Experimental Data ◽  
Thermal Characteristics ◽  
Electrical Heating ◽  
Liquid Oxygen ◽  
Research Centre ◽  
Bottom Surface ◽  
Rocket Engines ◽  
Fluid Temperature ◽  
Regenerative Cooling ◽  
Channel Bottom

The HYPROB Program, led by the Italian Aerospace Research Centre, has the aim to increase the Italian capabilities in the design and manufacture of liquid oxygen-methane rocket engines; in particular, the line, named HYPROB-BREAD, has the final goal of developing and testing a ground demonstrator of three-tons-thrust class, regeneratively cooled by liquid methane. Some intermediate breadboards have been conceived, realized and tested to deepen some technical issues: among them a specific breadboard, named MTP-B (Methane Thermal Properties Breadboard), has been designed and tested to investigate thermal characteristics of methane as a coolant, and perform the design validation activity by collecting experimental data. The concept is based on the electrical heating of a conductive material that transfers a thermal load, similar to those experienced in the regenerative cooling chambers, to a channel, having dimensions typical of regenerative cooling jackets. The test campaign has been successfully accomplished by collecting data, in terms of inlet/outlet fluid temperature and pressure, temperature at different stations and depth from the channel bottom surface. A numerical rebuilding activity has been planned to verify the numerical thermal models and engineering tools, adopted for the design of the final demonstrator. Thus, simulations on a fully 3-D model, including inlet and outlet interfaces, were performed on some test conditions after completing a preliminary activity on cold flow tests. The results of the numerical rebuilding are satisfactory since very good comparisons with the experimental data were observed.

CFD Modelling of Thermoacoustic Oscillations Inside an Atmospheric Test Rig Generated by a DLN Burner

2004 ◽  
Author(s):  
Stefano Tiribuzi
Keyword(s):  
Fluid Dynamic ◽  
Maximum Size ◽  
Geometrical Model ◽  
Operating Conditions ◽  
Test Facility ◽  
Research Centre ◽  
Cfd Modelling ◽  
Cfd Analyses ◽  
Time Required ◽  
Strong Control

The ENEL Produzione Research Centre of Pisa is deeply involved in the study of flame instabilities which could arise in particular operating conditions in the gas turbine equipped with Dry Low NOx (DLN) lean premixed combustors. An atmospheric pressure test facility, named TAO (Turbogas ad Accesso Ottico), is presently used to test, under scaled conditions, the onset of self-sustained thermoacoustic instabilities during the operation of a typical DLN burner. To support this activity, Computation Fluid Dynamic (CFD) analyses are carried out by means of KIEN, an in-house Reynold Averaged Navier Stokes (RANS) code, for simulating 3D instationary reactive flows. A 3D geometrical model, extending from the air plenum upstream the burner up to the end of the long exhaust tube, is adopted. The great extension of the calculation domain, coupled with the long real time required for the spontaneous onset of oscillations, could became computationally very onerous for Large Eddy Simulation (LES) codes, which require a strong control on the cell maximum size throughout the entire domain, while it can be acceptable for an instationary RANS code, which could use a coarser grid. This allows the simulation of long real transients with overnight runs. Starting from uniform no-motion conditions, the same fluctuating behaviour, detected during TAO experiments, spontaneously onsets in the numerical simulations too. The experimental and numerical frequencies are nearly the same and the amplitude of the pressure oscillations is very close. Supported by the congruence with experimental available data, computed results are utilized to get a more detailed description of the evolution of many thermofluiddynamics quantities during these instabilities. A wide sample of information that can be derived directly from the output of the instationary RANS code or by postprocessing of its results is provided.

Conceptual Regenerative Nozzle Cooling Design for a Hydroxyl-Terminated Polybutadiene and Oxygen Hybrid Rocket Engine

2017 ◽  
Author(s):  
Luis R. Robles ◽  
Johnny Ho ◽  
Bao Nguyen ◽  
Geoffrey Wagner ◽  
Jeremy Surmi ◽  
...  
Keyword(s):  
Combustion Chamber ◽  
High Temperatures ◽  
Cooling System ◽  
Rocket Engine ◽  
Combustion Efficiency ◽  
Rocket Engines ◽  
Hybrid Rocket ◽  
Gaseous Oxygen ◽  
Rocket Nozzle ◽  
Regenerative Cycle

Regenerative rocket nozzle cooling technology is well developed for liquid fueled rocket engines, but the technology has yet to be widely applied to hybrid rockets. Liquid engines use fuel as coolant, and while the oxidizers typically used in hybrids are not as efficient at conducting heat, the increased renewability of a rocket using regenerative cycle should still make the technology attractive. Due to the high temperatures that permeate throughout a rocket nozzle, most nozzles are predisposed to ablation, supporting the need to implement a nozzle cooling system. This paper presents a proof-of-concept regenerative cooling system for a hybrid engine which uses hydroxyl-terminated polybutadiene (HTPB) as its solid fuel and gaseous oxygen (O2) as its oxidizer, whereby a portion of gaseous oxygen is injected directly into the combustion chamber and another portion is routed up through grooves on the exterior of a copper-chromium nozzle and, afterwards, injected into the combustion chamber. Using O2 as a coolant will significantly lower the temperature of the nozzle which will prevent ablation due to the high temperatures produced by the exhaust. Additional advantages are an increase in combustion efficiency due to the heated O2 being used for combustion and an increased overall efficiency from the regenerative cycle. A computational model is presented, and several experiments are performed using computational fluid dynamics (CFD).

Laser Ignition of Liquid-Oxygen–Gaseous-Hydrogen Fuel in a Large-Scale Combustion Chamber

2019 ◽  
pp. 104-114
Author(s):  
S.G. Rebrov ◽  
V.A. Golubev ◽  
Y.P. Kosmachev ◽  
V.P. Kosmacheva
Keyword(s):  
Combustion Chamber ◽  
Large Scale ◽  
Rocket Engine ◽  
Laser System ◽  
Optical Breakdown ◽  
Gaseous Hydrogen ◽  
Liquid Oxygen ◽  
Research Centre ◽  
Laser Ignition ◽  
Experimental Chamber

The article presents a review of the results of studies of laser ignition of a cryogenic mixture (gaseous hydrogen and liquid oxygen) in an experimental combustion chamber, carried out at the bench testing facility of KBKhA (Voronezh). A laser ignition module specially designed at the Keldysh Research Centre and with parameters optimized for use in the rocket engine launch system was used during the experiments. Fuel ignition by the laser system occurred directly in the experimental chamber without the use of an ignition device or pre-chamber. To implement this ignition method, inflammation of the fuel in the chamber was carried out by focusing the laser radiation into the mixture, with the initiation of a spark of optical breakdown in the selected area with conditions favorable for the start of combustion. The results of the experiments confirmed the efficiency of the laser module during both standalone and firing tests, including multiple launches of the propulsion unit operated on a cryogenic mixture (gaseous hydrogen and liquid oxygen).

Adjoint-Based Optimisation of Rocket Engine Turbine Blades

2021 ◽  
Author(s):  
Bhupinder Singh Sanghera ◽  
Nitish Anand ◽  
Louis Souverein ◽  
Loïc Penin ◽  
Matteo Pini
Keyword(s):  
Shock Waves ◽  
Fluid Dynamic ◽  
Rocket Engine ◽  
Pressure Ratio ◽  
Turbine Blades ◽  
Shape Optimisation ◽  
Rocket Engines ◽  
Gas Generator ◽  
Axial Turbine ◽  
Loss Mechanisms

Abstract Axial turbine stages for gas generator cycle type rocket engines typically employ highly-loaded supersonic stator vanes. Consequently, the flow pattern downstream of the vanes is characterized by shock waves which induce high-frequency excitation on the subsequent rotor. For this reason, the optimal design of the stator is crucial in the context of development of the next generation of high-performance rocket engines, where reusability is a principal design criterion. A thorough comprehension of the loss mechanisms combined with the adoption of automated optimisation techniques can therefore enable new stator designs that may provide large benefits in terms of overall turbine performance and lifespan. The scope of this study stems from these considerations and its objective is twofold, namely i) the shape optimisation of a supersonic stator for rocket engines and ii) the investigation of the loss mechanisms in supersonic axial turbine stator vanes at on- and off-design conditions. The investigation is performed on stator vanes that are under development for the first turbine stage of a gas generator cycle type rocket engine. The stator vanes are therefore optimised in order to reduce the profile losses by exploiting a novel adjoint optimisation framework for turbomachinery implemented in the open-source code SU2. The effect on the resulting flow field and loss sources is finally investigated. Results show that entropy based loss coefficient gains of 6% can be achieved via shape optimisation and that the fluid-dynamic performance of these vanes is less sensitive to changes in pressure ratio compared to the performance provided by the baseline configuration. Eventually, shock-waves remain the primary loss source.

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