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Hydrogen oxygen rocket engine two phase liquid hydrogen pump capability and hydrodynamic design, analyzing constant-quality flow, acoustic effects, compressible flow and cavitation
Hydrogen oxygen rocket engine two phase liquid hydrogen pump capability and hydrodynamic design, analyzing constant-quality flow, acoustic effects, compressible flow and cavitation
Liquid propellant rocket engine combustion instabilities, describing unsteady combustor flow by single nonlinear wave equation with stable and unstable finite amplitude limit cycles
When a Rocket Engine shuts down under a fuel rich environment, a significant amount of unburned RP-1 is trapped In the engine. It is necessary to clean the residual RP-1 prior to subsequent firing to avoid any explosion due to detonation. The conventional method is to dry RP-1 with inert gas such as Nitrogen or Helium. It is difficult to estimate the drying time unless the engine is adequately equipped with instruments to measure the trace of RP-1 during the drying process. Such instrumentation in flight hardware is often impractical and costly. On the other hand numerical modeling of the drying process can provide a good insight for a satisfactory operation of the process. A numerical model can provide answer to questions such as a) how long it takes to dry, b) which fluid is a better dryer for RP-1, c) how to reduce drying time etc. The purpose of the present paper is to describe a numerical model of drying RP-1 trapped in a cavity with flowing nitrogen or helium. The numerical model assumes one dimensional flow of drying fluid in contact with liquid pool of RP-1. An evaporative mass transfer takes place across the contact surface.
A rocket engine igniter has been created that uses a pneumatically driven hammer that, by specialized geometry, is induced into an oscillatory state that can be used to either repeatedly impact a piezoelectric crystal with sufficient force to generate a spark capable of initiating combustion, or can be used with any other system capable of generating a spark from direct oscillatory motion. This innovation uses the energy of flowing gaseous propellant, which by means of pressure differentials and kinetic motion, causes a hammer object to oscillate. The concept works by mass flows being induced through orifices on both sides of a cylindrical tube with one or more vent paths. As the mass flow enters the chamber, the pressure differential is caused because the hammer object is supplied with flow on one side and the other side is opened with access to the vent path. The object then crosses the vent opening and begins to slow because the pressure differential across the ball reverses due to the geometry in the tube. Eventually, the object stops because of the increasing pressure differential on the object until all of the kinetic energy has been transferred to the gas via compression. This is the point where the object reverses direction because of the pressure differential. This behavior excites a piezoelectric crystal via direct impact from the hammer object. The hammer strikes a piezoelectric crystal, then reverses direction, and the resultant high voltage created from the crystal is transferred via an electrode to a spark gap in the ignition zone, thereby providing a spark to ignite the engine. Magnets, or other retention methods, might be employed to favorably position the hammer object prior to start, but are not necessary to maintain the oscillatory behavior. Various manifestations of the igniter have been developed and tested to improve device efficiency, and some improved designs are capable of operation at gas flow rates of a fraction of a gram per second (0.001 lb/s) and pressure drops on the order of 30 to 50 kilopascal (a few psi). An analytical model has been created and tested in conjunction with a precisely calibrated reference model. The analytical model accurately captures the overall behavior of this innovation. The model is a simple "volume-orifice" concept, with each chamber considered a single temperature and pressure "node" connected to adjacent nodes, or to vent paths through flow control orifices. Mass and energy balances are applied to each node, with gas flow predicted using simple compressible flow equations.
Rocket engine turbopump assembly test capability design, facilities, cryogenic systems, control requirements, computer simulation, activation and checkout
The issues of Chemical Rocket Engine Control are broadly covered. The basic feedback information and control variables used in expendable and reusable rocket engines, such as Space Shuttle Main Engine, are discussed. The deficiencies of current approaches are considered and a brief introduction to Intelligent Control Systems for rocket engines (and vehicles) is presented.
In the past, the design of rocket engines has primarily relied on the cold flow/hot fire test, and the empirical correlations developed based on the database from previous designs. However, it is very costly to fabricate and test various hardware designs during the design cycle, whereas the empirical model becomes unreliable in designing the advanced rocket engine where its operating conditions exceed the range of the database. The main goal of the 2nd Generation Reusable Launching Vehicle (GEN-II RLV) is to reduce the cost per payload and to extend the life of the hardware, which poses a great challenge to the rocket engine design. Hence, understanding the flow characteristics in each engine components is thus critical to the engine design. In the last few decades, the methodology of computational fluid dynamics (CFD) has been advanced to be a mature tool of analyzing various engine components. Therefore, it is important for the CFD design tool to be able to properly simulate the hot flow environment near the liquid injector, and thus to accurately predict the heat load to the injector faceplate. However, to date it is still not feasible to conduct CFD simulations of the detailed flowfield with very complicated geometries such as fluid flow and heat transfer in an injector assembly and through a porous plate, which requires gigantic computer memories and power to resolve the detailed geometry. The rigimesh (a sintered metal material), utilized to reduce the heat load to the faceplate, is one of the design concepts for the injector faceplate of the GEN-II RLV. In addition, the injector assembly is designed to distribute propellants into the combustion chamber of the liquid rocket engine. A porosity mode thus becomes a necessity for the CFD code in order to efficiently simulate the flow and heat transfer in these porous media, and maintain good accuracy in describing the flow fields. Currently, the FDNS (Finite Difference Navier-Stakes) code is one of the CFD codes which are most widely used by research engineers at NASA Marshall Space Flight Center (MSFC) to simulate various flow problems related to rocket engines. The objective of this research work during the 10-week summer faculty fellowship program was to 1) debug the framework of the porosity model in the current FDNS code, and 2) validate the porosity model by simulating flows through various porous media such as tube banks and porous plate.
This paper describes how commercial pump technology was modified for rocket engine application by trading long life for low weight. The component life of typical turbopump components is presented, and the modes of failure will be discussed. A comparison of the basic commercial and rocket engine pump requirements such as flow, pressure, etc. is made. It is concluded that rocket engine pump technology can be applied directly only if similar basic requirements exist in emergency or stand-by equipment. The specific design requirements for rocket engine turbopumps are presented and the need for high speeds is explained by the weight-speed relationship and the optimum NPSH of the tank and pump system. The trend of required rocket engine pump discharge pressure and future operational requirements is predicted. The need for inducers and boost pumps in rocket engines is explained, and the results of the flat-plate-inducer cavitation therein are briefly presented. Design solutions and material selections for high-pressure pump housings, ultra-high tip speed impellers, and pump packaging are demonstrated. As a typical example of utilizing rocket engine technology for commercial application, a helicopter stand-by engine i shown. This standby engine is designed from existing rocket engine components, using JATO rockets as the prime energy source.
Regenerative cooling of rocket engines considering heat transfer, chamber wall, coolant tube and pressure drop
A device for installing rocket engines is reported that is supported at a cant relative to vertical by an axially extensible, tiltable pedestal. A lifting platform supports the rocket engine at its thrust chamber exit, including a mount having a concentric base characterized by a concave bearing surface, a plurality of uniformly spaced legs extended radially from the base, and an annular receiver coaxially aligned with the base and affixed to the distal ends of said legs for receiving the thrust chamber exit. The lifting platform rests on a seat concentrically related to the pedestal and affixed to an extended end portion thereof having a convex bearing surface mated in sliding engagement with the concave bearing surface of the annular base for accommodating a rocking motion of the platform.
This book intends to build a bridge for the student and the young engineer: to link the rocket propulsion fundamentals and elements (which are well covered in the literature) with the actual rocket engine design and development work as it is carried out in industry (which is very little, if at all covered in literature). The book attempts to further the understanding of the realistic application of liquid rocket propulsion theories, and to help avoid or at least reduce time and money consuming errors and disappointments. In so doing, it also attempts to digest and consolidate numerous closely related subjects, hitherto often treated as separate, bringing them up to date at the same time. The book was written "on the job" for use by those active in all phases of engine systems design, development, and application, in industry as well as government agencies. Since it addresses itself to human beings set out to create new machines, rather than describing machines about to dominate man, the language chosen may not always be "functional" in the strict sense of the word. The book presents sufficient detail to familiarize and educate thoroughly those responsible for various aspects of liquid propellant rocketry, including engine systems design, engine development, and flight vehicle application. It should enable the rocket engineer to conduct, independently, complete or partial engine systems preliminary detail designs and to understand and judge the activities in, and the problems, limitations, and "facts of life~ of the various subsystems making up a complete engine system. It also attempts to educate those ultimately interested in specialized subsystems and component design (thrust chamber, turbopump, control valves, etc.) about their own as well as neighboring subsystems and about the complete engine system. This should enable the student to prepare realistic analytical calculations and design layouts with a long headstart toward the final specialized designs for subsystem production release. Special emphasis has been placed on engine flight application to stimulate engine systems and subsystem designers to think in these terms from the outset. The book is intended as a textbook, with specific consideration of the teacher without industry experience. We hope it will stimulate those desiring to specialize in the area of a rocket engine subsystem by supplying adequate information to enable them to benefit fully from the specialized literature. Thus, it provides a realistic expert introduction for those joining the liquid propellant rocket engine field.
Protective coating systems for refractory metal used in rocket engine nozzles - rocket engine test firing
During static testing of multi-million pound thrust rocket engines areas adjacent to the test site have been subjected to the noise generated by rocket engines. Structural damage claims and subjective complaints were filed by those who alleged that the noise levels were excessive. The statistical analysis of these claims and complaints which were filed during these rocket engine development programs led to the determination of a relationship between claims and overall sound pressure level. Community exposure criteria are then assessed based on what can be considered allowable acoustic environments from large rocket engines.
Film and transpiration cooled baffles for M-1 engine
Vibration environment of liquid propellant rocket engines
Two RL10A-3-3A rocket engines comprise the main propulsion system for the Centaur upper stage vehicle. Centaur is used with bod Titan and Atlas launch vehicles, carrying military and civilian payloads from high altitudes into orbit and beyond. The RL10 has delivered highly reliable service for the past 30 years. Recently, however, there have been two in-flight failures which have refocused attention on the RL10. This heightened interest has sparked a desire for an independent RL10 modeling capability within NASA and th Air Force. Pratt & Whitney, which presently has the most detailed model of the RL10, also sees merit in having an independent model which could be used as a cross-check with their own simulations. The Space Propulsion Technology Division (SPTD) at the NASA Lewis Research Center has developed a computer model of the RL10A-3-3A. A project team was formed, consisting of experts in the areas of turbomachinery, combustion, and heat transfer. The overall goal of the project was to provide a model of the entire RL10 rocket engine for government use. In the course of the project, the major engine components have been modeled using a combination of simple correlations and detailed component analysis tools (computer codes). The results of these component analyses were verified with data provided by Pratt & Whitney. Select modeling results and test data curves were then integrated to form the RL10 engine system model The purpose of this report is to introduce the reader to the RL10 rocket engine and to describe the engine system model. The RL10 engine and its application to U.S. launch vehicles are described first, followed by a summary of the SPTD project organization, goals, and accomplishments. Simulated output from the system model are shown in comparison with test and flight data for start transient, steady state, and shut-down transient operations. Detailed descriptions of all component analyses, including those not selected for integration with the system model, are included as appendices.