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Liquid Rocket Engine Design for Additive Manufacturing
Abstract
This paper discusses the processes involved in the additive manufacturing of a regenerative and film-cooled liquid rocket engine with a thrust of 10 kN using Inconel 718, while detailing validation techniques. A description of the objectives and design constraints provide the context and motivations. Computational Fluid Dynamics (CFD) models were developed and provided the expected pressure and thermal regimes under regenerative and film cooling. Additionally, Finite Element (FE) models were used to predict the capabilities of the engine structure. A description of 3D printing methods highlights the benefits and limitations of the technology, specifically the influence the design of liquid rocket engines. A pintle injector is used, printed as a separate, easily removable and replaceable component. Issues related to overhangs, surface roughness, and shrinkage; aspects related to post-print processing and the need to minimize machining are discussed. Results from the CT scans of the engine and its components are presented. The paper also outlines the series of tests that will be performed on this engine to verify its performance and provide design basis for future works This engine will be used to power the reusable flight vehicle that is under development at the Kyushu Institute of Technology in Japan. The student-led Liquid Propulsion Laboratory at the University of Southern California is responsible for the work detailed below.
... Dressler [1], Mercieca [7] or Vasques [16], T M R can also be found as the inverse radialto-axial ratio, such as in Sakaki [17], Austin [18] or Fessla [19]. ...
The pintle injector is distinguished by its unique geometry and injection characteristics, compared to the impinging or coaxial distributed-element injectors typically used on liquid bipropellant rocket engines. The pintle injector design can deliver high combustion efficiency (typically 96-99%) and enables some unique operating features, such as deep throttling and injector face shut-off. Its design simplicity makes it ideally suited for use on low cost engines [1]. The thesis work is focused on the development and validation of analytical models and software for the design of a pintle injector for a LOX-CH4 liquid rocket engine. A numerical tool has been developed to preliminary design the injector and give insights on the achievable performance of the system. It defines a layout of the injector orifices then provides the thermal analysis of the components exposed to the heat of the combustion chamber. Once defined a configuration that achieves both high performance and withstands the thermal loads, computational fluid dynamics has been employed to support the detailed design of the components, to identify weaknesses in the flow field and to optimize its topology. A technology demonstrator has been designed and manufactured, to be ready for a cold flow characterization campaign. Significantly lower development costs has been achieved optimizing the model for additive layer manufacturing, which is imposing as a major breakthrough in propulsion technology and allows to realize complex systems near net shape and with a lower number of parts.
... Fessl et al. 11 used additive manufacturing (AM) to print the liquid rocket engine they had designed. They used an EOS M290 metal printer that uses Selective Laser Sintering (SLS) to print with materials such as Inconel 718. ...
The paper deals with the problem of using prototyping in developing rocket and space technology products and their production systems as one of the components of the “Prototyping, manufacturing, testing of rocket and space technology products” field within the formation of an end-to-end design and production and experimental environment. The study shows the role of the prototyping technology in the system of model-based development of both the products themselves and their production systems; gives the theoretical basis and characteristics of the area of application of prototyping methods at the initial stages of design; demonstrates methodical approaches to choosing the subject of prototyping and the procedure for working with it for launch vehicles and spacecraft within the framework of students’ work on the development of rocket and space systems for space missions. We analyzed and introduced the experience of using the technology in the educational process and identified next stages for developing and implementing prototyping tools to further implement the end-to-end prototyping technology in solving practical problems of the rocket and space industry.
Metal additive manufacturing involves manufacturing techniques that add material to produce metallic components, typically layer by layer. The substantial growth in this technology is partly driven by its opportunity for commercial and performance benefits in the aerospace industry. The fundamental opportunities for metal additive manufacturing in aerospace applications include: significant cost and lead-time reductions, novel materials and unique design solutions, mass reduction of components through highly efficient and lightweight designs, and consolidation of multiple components for performance enhancement or risk management, e.g. through internal cooling features in thermally loaded components or by eliminating traditional joining processes. These opportunities are being commercially applied in a range of high-profile aerospace applications including liquid-fuel rocket engines, propellant tanks, satellite components, heat exchangers, turbomachinery, valves, and sustainment of legacy systems. This paper provides a comprehensive review of metal additive manufacturing in the aerospace industry (from industrial/popular as well as technical literature). This provides a current state of the art, while also summarizing the primary application scenarios and the associated commercial and technical benefits of additive manufacturing in these applications. Based on these observations, challenges and potential opportunities are highlighted for metal additive manufacturing for each application scenario.
The effective use of additive manufacturing requires careful design methods, specifically targeting subsequent post-processing and machining efforts. This paper covers the design, manufacturing and post-processing considerations that went into the development of an additively manufactured, bi-propellant, pressure-fed liquid rocket engine. Special considerations include initial constraints on engine design related to DMLS printing requirements, as well as post-printing ultrasonic cleaning, heat treatment, wire EDM and machining processes. Unexpected design related challenges and complications were found and solved throughout the manufacturing process. Solutions, such as developing custom machining tools, were devised to overcome these difficulties and ultimately allowed the completion of a hot-fire readiness testing campaign. This project was designed and completed by the student-led Liquid Propulsion Laboratory at the University of Southern California.
High speed sintering is a novel additive manufacturing technology that uses inkjet printing and infra-red energy to selectively sinter polymeric powder. The research presented here investigates the effect of build orientation on dimensional accuracy, density, mechanical properties and surface roughness of high speed sintered parts. Tensile specimens were built through seven different angles between and including the XY (horizontal) and ZY (vertical) planes and analysed. The effect of the PUSh™ process was also investigated across this range of build orientations. The results show that build orientation does infuence the properties of the parts. A number of mechanical properties showed a relationship with build orientation. Density was seen to decrease as the angle increased from XY towards ZY. This increase in angle was shown to increase surface roughness while ultimate tensile strength and elongation at break decreased. At all build orientations, the PUSh™ process significantly reduces surface roughness, mildly increases part density and had a small effect on ultimate tensile strength whilst showing a small but consistent increase in elongation at break.
The pintle injector rocket engine is fundamentally different from other rocket engines, which nearly universally employ a series of separate propellant injection orifices distributed across the diameter of the headend of the combustion chamber. The pintle’s central, singular injection geometry results in a combustion chamber flowfield that varies greatly from that of conventional rocket engines. These differences result in certain operational characteristics of great benefit to rocket engine design, performance, stability, and test flexibility. The mid-1950’s origin of the pintle injector concept and the subsequent early development work and applications in rocket engines are reviewed. The pintle engine’s key design and operational features are compared to conventional rocket engines. Pintle injector design refinements and associated recent applications are discussed. The presentation includes photographs and summaries of many different rocket engines that TRW has developed and successfully flown, each of which used the pintle injector.
A thermal stability and heat transfer investigation was conducted using five common hydrocarbon fuels: JP-7, JP-8, JP-8+100, JP-10, and RP-1. Tests were conducted at the NASA Glenn Research Center Heated Tube Facility using resistively heated tube sections to simulate conditions encountered in regeneratively cooled rocket engines. Six tests were conducted for each fuel using a halffactorial test matrix with fuel average velocities set at 25 ft/sec and 75 ft/sec, wall temperatures of 750 °F and 1000 °F, and OFE 101 copper and stainless steel 304 as tube materials. Test section average pressure and fuel bulk outlet temperature were held constant at 1000 psi and 500 °F respectively, while target run times for all tests were 20 minutes. Carbon deposition rates for all tests were comparable (≈100 μg/cm2-hr) with the exception of the JP-8/JP-8+100 copper tests, which yielded deposition rates an order of magnitude greater (≈1000 μg/cm2-hr). The heat transfer data collected from all tests could be well correlated by a standard pipe flow equation, which yielded an R2 value of 0.96. © 2002 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
The present publication introduces the fundamental principles of liquid-propellant rocket engines that are required for actual design applications. After an introduction to the gas-flow properties, performance parameters, and propellant types for these rocket engines, engine requirements are set forth for such indicators as duration, weight, envelope, and thrust level. Sample calculations are given for A-1 through A-4 stage engines, and the design of thrust chambers and combustion devices is reviewed for factors such as injectors, ignition, cooling, and instability. Other elements that are discussed in detail are: gas-pressurized and turbopump-propellant feed systems; control and condition-monitoring conditions; interconnecting components and mounts; propellant tanks; and designs for specific space applications.
Results of a study to determine the applicability of film cooling to rocket engines in the 10to 1000-lb-thrust range using earth-storable, space-storable, and cryogenic propellant combinations are presented. The accuracy of the analytical model was verified with test data using the propellant combinations of nitrogen tetroxide/monomethylhydrazine and fluorine/monomethylhydrazine (N2O4/MMH and F2/MMH). Data are presented which illustrate the interrelational effects of mixture ratio, specific impulse, wall temperature, and the percent of film cooling. The studies show that N2O4/MMH and C1F5/MMH and other similar earth-storable propellant combinations are readily adaptable to small film-cooled spacecraft engines. The space-storable and deep-cryogenic systems OF2/MMH, OF2/B2H6, OF2/CH4(LPG's), F2/MMH, O2/H2, and F2/H2 will require the use of film cooling and high-temperature materials if the potentially high performance they possess is to be realized. Otherwise, the systems have to be operated fuel rich with high percentages of fuel film cooling, which will compromise the deliverable specific impulse and system weight. For example, if a 100-lb-thrust F2/H2 engine operating at 100-psia chamber pressure was needed to provide 400 sec of /sp at a mixture ratio of 8.0 or greater, the minimum percent of film cooling and minimum wall temperature capability of the chamber material must be 40% and 5000QF, respectively.
We have developed surrogate mixture models to represent the thermophysical properties of two kerosene rocket propellants, RP-1 and RP-2. The surrogates were developed with a procedure that incorporated experimental data for the density, sound speed, viscosity, thermal conductivity, and the advanced distillation curves for samples of the two fuels. The surrogate for RP-1 contains four components (α-methyldecalin, n-dodecane, 5-methylnonane, and heptylcyclohexane), and the surrogate for RP-2 contains five components (α-methyldecalin, n-dodecane, 5-methylnonane, 2,4-dimethylnonane, and heptylcyclohexane). Comparisons with experimental data demonstrate that the models are able to represent the density, sound speed, viscosity, and thermal conductivity of both fuels to within (at a 95% confidence level) 0.4, 2, 2, and 4%, respectively. The volatility behavior, as measured by the advanced distillation curves, is reproduced to within 0.5%.
Liquid rocket thrust chambers aspects of modeling, analysis, and design
- M Habiballah
- V Yang
Habiballah, M., & Yang, V. (2004). Liquid
rocket thrust chambers aspects of modeling, analysis,
and design. Reston, Va.: American Institute of
Aeronautics and Astronautics, Inc.
Liquid Rocket Engine Component Water-Flow Test Stand
- M Moruzzi
- J Fessl
Moruzzi, M. Fessl, J., "Liquid Rocket Engine
Component Water-Flow Test Stand," Journal of
Propulsion & Power (not yet published).