scholarly journals Influence of propellant leakage from pump area into turbine area on turbo-pump operation stability

2021 ◽  
Vol 27 (1) ◽  
pp. 97-102
Author(s):  
M.V. Andriievskyi ◽  
◽  
Yu.O. Mitikov ◽  
Keyword(s):  
Combustion Engine ◽  
Specific Impulse ◽  
Liquid Oxygen ◽  
Stable Operation ◽  
Rocket Engines ◽  
Staged Combustion ◽  
Pump Operation ◽  
Operation Stability ◽  
The Stability ◽  
Propellant Rocket

There is an increasing trend to liquid-propellant rocket engines which run on eco-friendly storable propellant. This trend is mostly dictated by the refusal to use traditional toxic storable propellant in many countries. The most widespread eco-friendly storable propellant is hydrogen peroxide with kerosene. Though, this propellant has a lower specific impulse in comparison with traditional liquid oxygen with kerosene. To compensate the loss of specific impulse, there is a reason to design a staged combustion engine. Evidently, the turbopump is the most complicated system in the staged combustion propulsion system. This fact makes research devoted to turbo-pumps a top priority. The paper aims to determine the influence of propellant leakage from the pump area into the turbine area and create recommendations which would allow organizing the stable operation of turbopump. As a result of turbopump staged combustion cycle testing, a conclusion had been made that leakage, which opens during the test, significantly influences the stability of turbopump operation. Depending on the amount of leakage, the turbine generated power drop was between 20 and 45%, which led to a decrease in rotation speed and outlet pressure of the pump. During the R&D process, a way of leakage influence elimination had been offered. Formulated recommendations may be used during the design process of the turbopump for staged combustion liquid propulsion systems.

ACOUSTIC CAVITIES DESIGN PROCEDURES

2007 ◽  
Vol 6 (2) ◽  
pp. 27
Author(s):  
A. Santana Jr. ◽  
M. S. Silva ◽  
P. T. Lacava ◽  
L. C. S. Góes
Keyword(s):  
Solid Propellant ◽  
Rocket Engines ◽  
Combustion Chambers ◽  
High Frequencies ◽  
Design Procedures ◽  
Solid Propellant Rocket ◽  
Acoustic Cavities ◽  
The Stability ◽  
Propellant Rocket ◽  
Damping Systems

Combustion instability is recognized as one of the major problems frequently faced by engineers during the development of either liquid or solid propellant rocket engines. The performance of the engine can be highly affected by these high frequencies instabilities, possibly leading the rocket to an explosion. The main goal while studying combustion chambers instability, either by means of baffles or acoustic absorbers, is to achieve the stability needed using the simplest possible manner. This paper has the purpose of studying combustion chambers instabilities, as well as the design of acoustic absorbers capable of reducing their eigenfrequencies. Damping systems act on the chamber eigenfrequency, which has to be, therefore, previously known.

Increasing the specific impulse of liquid oxygen + kerosene rocket engines by introducing hydrogen into the combustion chamber

2019 ◽  
Author(s):  
S.A. Orlin ◽  
A.V. Orlov
Keyword(s):  
Combustion Chamber ◽  
Thermodynamic Analysis ◽  
Specific Impulse ◽  
Launch Vehicle ◽  
Liquid Oxygen ◽  
Rocket Engines ◽  
State Technical ◽  
Analytical Studies ◽  
Technical University ◽  
Engine Type

The investigation carried out at the Bauman Moscow State Technical University is aimed at establishing whether it may be possible to increase the specific impulse of liquid oxygen + kerosene rocket engines. It involved analytical studies of increasing specific impulse by introducing hydrogen into the oxygen/kerosene propellant. We confirm that, in the case of the oxygen/kerosene propellant used in the first stage engines, introducing hydrogen into the combustion chamber may increase its specific impulse. The results of our thermodynamic analysis show that the specific impulse increase is a function of the mass of hydrogen introduced. This enables the same engine type to be used for the first and second stages of a launch vehicle, which makes the whole system considerably less expensive and more reliable.

scholarly journals Recent Experimental Efforts on High-Pressure Supercritical Injection for Liquid Rockets and Their Implications

2012 ◽  
Vol 2012 ◽  
pp. 1-31 ◽  
Author(s):  
Bruce Chehroudi
Keyword(s):  
High Pressure ◽  
Critical Point ◽  
Space Shuttle ◽  
Specific Impulse ◽  
Liquid Oxygen ◽  
Liquid Jets ◽  
Rocket Engines ◽  
Liquid Rocket Engines ◽  
Liquid Rocket ◽  
Air Force Research Laboratory

Pressure and temperature of the liquid rocket thrust chambers into which propellants are injected have been in an ascending trajectory to gain higher specific impulse. It is quite possible then that the thermodynamic condition into which liquid propellants are injected reaches or surpasses the critical point of one or more of the injected fluids. For example, in cryogenic hydrogen/oxygen liquid rocket engines, such as Space Shuttle Main Engine (SSME) or Vulcain (Ariane 5), the injected liquid oxygen finds itself in a supercritical condition. Very little detailed information was available on the behavior of liquid jets under such a harsh environment nearly two decades ago. The author had the opportunity to be intimately involved in the evolutionary understanding of injection processes at the Air Force Research Laboratory (AFRL), spanning sub- to supercritical conditions during this period. The information included here attempts to present a coherent summary of experimental achievements pertinent to liquid rockets, focusing only on the injection of nonreacting cryogenic liquids into a high-pressure environment surpassing the critical point of at least one of the propellants. Moreover, some implications of the results acquired under such an environment are offered in the context of the liquid rocket combustion instability problem.

scholarly journals Comparison of cylindrical and non-cylindrical grain internal ballistic behavior of hybrid rocket engines and solid rocket motors

2021 ◽  
Author(s):  
Zhiliu Lu
Keyword(s):  
Specific Impulse ◽  
Propulsion System ◽  
Simulation Program ◽  
Rocket Engines ◽  
Hybrid Rocket ◽  
Ballistic Performance ◽  
Flow Process ◽  
Rocket Motors ◽  
Internal Ballistic ◽  
Propellant Rocket

Hybrid rocket engines (HREs) are a chemical propulsion system that nominally combine the advantages of liquid-propellant rocket engines (LREs) and solid-propellant rocket motors (SRMs). HREs in some cases can have a higher specific impulse and better controllability than SRMs, and lower cost and engineering complexity than LREs. For HREs and SRMs, both kinds of rocket engine employ a solid fuel grain, and the chosen grain configuration is a crucial point of their design. Different grain configurations have different internal ballistic behavior, which in turn can deliver different engine performance. A cylindrical grain design is a very common design for SRMs and HREs. A non-cylindrical-grain is a more complex grain configuration (than cylindrical) that has been used in many SRMs, and is also a choice for some HREs. However, while an HRE and an SRM can employ the same fuel grain configuration, the resulting internal ballistic behavior would not be expected to be the same. Pressure-dependent burning tends to dominate in SRMs, while axial flow-dependent burning tends to dominate in HREs. To help demonstrate in a more direct manner the influence of the differing combustion processes on the same fuel grain configuration used by an HRE and SRM, a number of internal ballistic simulations are undertaken for the present study. For the reference SRM cases looked at, an internal ballistic simulation program that has the capability of predicting head-end pressure and thrust as a function of time into a simulated firing is utilized for the present investigation; for the corresponding HRE cases, a simulation program is used to simulate the burning and flow process of these engines. For the present investigation, the two simulation programs are used to simulate the internal ballistic performance of various HREs and SRMs employing comparable cylindrical and non-cylindrical fuel grain configurations. The predicted performance results, in terms of pressure and thrust, are consistent with expectations that one would have for both propulsion system types.

Spray and Combustion Modeling in High Pressure Cryogenic Jet Flames

2012 ◽  
Author(s):  
Maria Grazia De Giorgi ◽  
Aldebara Sciolti ◽  
Antonio Ficarella
Keyword(s):  
Supercritical Pressure ◽  
Liquid Oxygen ◽  
Rocket Engines ◽  
Liquid Propellant ◽  
Combustion Modeling ◽  
Discrete Phase ◽  
Real Behavior ◽  
Shear Coaxial Injector ◽  
Coaxial Injector ◽  
Propellant Rocket

The aim of the present work is the investigation of the combustion phenomenon in liquid-propellant rocket engines. The combustion of liquid oxygen and gaseous methane in a shear coaxial injector under supercritical pressure was analyzed. To realize an efficient numerical description of the phenomena, it is important to treat the LOx jet in a manner which takes into account its real behavior. In the present work different kinetics, combustion models and thermodynamics approaches were used in association with the description of the jet as a discrete phase. For all the approaches used, a comparison with experimental data from literature was performed.

scholarly journals Comparison of cylindrical and non-cylindrical grain internal ballistic behavior of hybrid rocket engines and solid rocket motors

2021 ◽  
Author(s):  
Zhiliu Lu
Keyword(s):  
Specific Impulse ◽  
Propulsion System ◽  
Simulation Program ◽  
Rocket Engines ◽  
Hybrid Rocket ◽  
Ballistic Performance ◽  
Flow Process ◽  
Rocket Motors ◽  
Internal Ballistic ◽  
Propellant Rocket

Hybrid rocket engines (HREs) are a chemical propulsion system that nominally combine the advantages of liquid-propellant rocket engines (LREs) and solid-propellant rocket motors (SRMs). HREs in some cases can have a higher specific impulse and better controllability than SRMs, and lower cost and engineering complexity than LREs. For HREs and SRMs, both kinds of rocket engine employ a solid fuel grain, and the chosen grain configuration is a crucial point of their design. Different grain configurations have different internal ballistic behavior, which in turn can deliver different engine performance. A cylindrical grain design is a very common design for SRMs and HREs. A non-cylindrical-grain is a more complex grain configuration (than cylindrical) that has been used in many SRMs, and is also a choice for some HREs. However, while an HRE and an SRM can employ the same fuel grain configuration, the resulting internal ballistic behavior would not be expected to be the same. Pressure-dependent burning tends to dominate in SRMs, while axial flow-dependent burning tends to dominate in HREs. To help demonstrate in a more direct manner the influence of the differing combustion processes on the same fuel grain configuration used by an HRE and SRM, a number of internal ballistic simulations are undertaken for the present study. For the reference SRM cases looked at, an internal ballistic simulation program that has the capability of predicting head-end pressure and thrust as a function of time into a simulated firing is utilized for the present investigation; for the corresponding HRE cases, a simulation program is used to simulate the burning and flow process of these engines. For the present investigation, the two simulation programs are used to simulate the internal ballistic performance of various HREs and SRMs employing comparable cylindrical and non-cylindrical fuel grain configurations. The predicted performance results, in terms of pressure and thrust, are consistent with expectations that one would have for both propulsion system types.

scholarly journals Theoretical Analysis of Ammonium-perchlorate Based Composite Propellants with RDX Containing Small Size Particles of Beryllium

2020 ◽  
Author(s):  
Paulo Alexandre Rodrigues de Vasconcelos Figueiredo ◽  
Francisco Miguel Ribeiro Proença Brojo
Keyword(s):  
Ammonium Perchlorate ◽  
Rocket Engine ◽  
Specific Impulse ◽  
Combustion Products ◽  
Rocket Engines ◽  
Oxygen Balance ◽  
Metallic Fuels ◽  
Ignition And Combustion ◽  
Operational Envelope ◽  
Propellant Rocket

Rocket engines have been developed for at least the last six decades. There is a need to improve the actual solid propellant grain for rocket engines through the addiction of metallic fuels in the mixture as well as the addiction of energetic binders to stabilize the combustion. The rocket industry expects the launchers to be reliable, to be faster, stable and to have longer times of operation for the most possible payload weight (operational envelope). New propellants should have optimized ignition and combustion time rates reducing the possibility of negative oxygen balance thus reducing detonation process. Deflagration process should be optimized for best performance of the rocket. In this evolution, small quantities of explosives have been used in the propellant in order to increase the operational burning time, hence, the specific impulse. Adding metallic fuels such as aluminum, boron or beryllium on double based composite propellants and ammonium perchlorate are expected to increase the propellant density over chemical stability and aging resistance. The study of heterogeneous propellants containing large amounts of fine beryllium and ammonium perchlorate,   it is necessary to understand the combustion products so to a proper evaluation of specific impulse, Mach number and mass flow of the mixture. In this study a mixture with nitramides (RDX – Cyclotrimethylene trinitramide) and ammonium perchlorate was analyzed with and without the addiction of small size particles of beryllium using a numerical algorithm. Therefore, this study relates the influence of beryllium  in the performance parameters of ammonium perchlorate based composite propellants. Keywords: Propellant, Rocket engine, RDX, Ammonium perclorate

The anaerobic treatment of a ligno-cellulosic substrate offering little natural pH buffering capacity

1998 ◽  
Vol 38 (4-5) ◽  
pp. 29-35 ◽  
Author(s):  
C. J. Banks ◽  
P. N. Humphreys
Keyword(s):  
Ph Control ◽  
Anaerobic Treatment ◽  
Buffering Capacity ◽  
Single Stage ◽  
Stable Operation ◽  
Reactor Performance ◽  
Two Stage ◽  
Ph Buffering ◽  
Stage System ◽  
The Stability

The stability and operational performance of single stage digestion with and without liquor recycle and two stage digestion were assessed using a mixture of paper and wood as the digestion substrate. Attempts to maintain stable digestion in both single stage reactors were unsuccessful due to the inherently low natural buffering capacity exhibited; this resulted in a rapid souring of the reactor due to unbuffered volatile fatty acid (VFA) accumulation. The use of lime to control pH was unsatisfactory due to interference with the carbonate/bicarbonate equilibrium resulting in wide oscillations in the control parameter. The two stage system overcame the pH stability problems allowing stable operation for a period of 200 days without any requirement for pH control; this was attributed to the rapid flushing of VFA from the first stage reactor into the second stage, where efficient conversion to methane was established. Reactor performance was judged to be satisfactory with the breakdown of 53% of influent volatile solids. It was concluded that the reactor configuration of the two stage system offers the potential for the treatment of cellulosic wastes with a sub-optimal carbon to nitrogen ratio for conventional digestion.

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