GRCop-42 is a high conductivity, high-strength dispersion strengthened copper-alloy for use in high heat flux applications such as liquid rocket engine combustion devices. This alloy is part of the family of NASAdeveloped GRCop, copper-chrome-niobium alloys. GRCop alloys were developed for harsh environments specific to regeneratively-cooled combustion chambers and nozzles with good oxidation resistance. Significant development was completed on the GRCop-84 and GRCop-42 alloys in the extruded wrought form demonstrating feasibility for combustion chambers. NASA has recently developed a process for additive manufacturing, specifically Powder Bed Fusion (PBF) or Selective Laser Melting (SLM), of GRCop-42 to establish parameters, characterize the material, and complete testing of components with complex internal features. This evolution of the GRCop-42 was based on the successful predecessor development work using GRCop-84 with the motivation of establishing a new copper-alloy option for use in NASA, government, and industry programs with SLM. A few advantages have been shown with the GRCop-42 that include higher conductivity and faster build speeds over the GRCop-84, and a simplified powder supply chain. Initial property development has shown that it is possible to produce high density builds with strengths equivalent to wrought GRCop-42 and a conductivity greater than GRCop-84. The GRCop-42 has completed process development and initial properties have been established. Several demonstrator combustion chambers have also been fabricated with the SLM GRCop-42 that include integral channels and closeouts. Additional test units have been fabricated and are completing substantial hot-fire testing to demonstrate performance of the material, process, and design.

No abstract

Abstract-The fluidic packaging of Power MEMS devices such as the MIT microengine and microrocket requires the fabrication of hermetic seals capable of withstanding temperature in the range 20-600 C and pressures in the range 100-300 atm. We describe an approach to such packaging by attaching Kovar metal tubes to a silicon device using glass seal technology. Failure due to fracture of the seals is a significant reliability concern in the baseline process: microscopy revealed a large number of voids in the glass, pre-cracks in the glass and silicon, and poor wetting of the glass to silicon. The effects of various processing and materials parameters on these phenomena were examined. A robust procedure, based on the use of metal-coated silicon substrates, was developed to ensure good wetting. The bending strength of single-tube specimens was determined at several temperatures. The dominant failure mode changed from fracture at room temperature to yielding of the glass and Kovar at 600 C. The strength in tension at room temperature was analyzed using Weibull statistics; these results indicate a probability of survival of 0 99 at an operational pressure of 125 atm at room temperature for single tubes and a corresponding probability of 0 9 for a packaged device with 11 joints. The residual stresses were analyzed using the method of finite elements and recommendations for the improvement of packaging reliability are suggested.[933]Index Terms-Kovar silicon seal, microengine, microfluidic packaging, Power MEMS.

No abstract

To support the mission for the NASA Vision for Space Exploration, the NASA Marshall Space Flight Center conducted a program in 2005 to improve the capability to predict local thermal compatibility and heat transfer in liquid propellant rocket engine combustion devices. The ultimate objective was to predict and hence reduce the local peak heat flux due to injector design, resulting in a significant improvement in overall engine reliability and durability. Such analyses are applicable to combustion devices in booster, upper stage, and in-space engines, as well as for small thrusters with few elements in the injector. In this program, single element and three-element injectors were hot-fire tested with liquid oxygen and ambient temperature gaseous hydrogen propellants a t The Pennsylvania State University Cryogenic Combustor Laboratory from May to August 2005. Local heat fluxes were measured in a 1-inch internal diameter heat sink combustion chamber using Medtherm coaxial thermocouples and Gardon heat flux gauges. Injectors were tested with shear coaxial and swirl coaxial elements, including recessed, flush and scarfed oxidizer post configurations, and concentric and non-concentric fuel annuli. This paper includes general descriptions of the experimental hardware, instrumentation, and results of the hot-fire testing for three of the single element injectors -recessed-post shear coaxial with concentric fuel, flush-post swirl coaxial with concentric fuel, and scarfed-post swirl coaxial with concentric fuel. Detailed geometry and test results will be published elsewhere to provide well-defined data sets for injector development and model validatation.https://ntrs.nasa.gov/search.jsp?R=20060047645 2018-05-13T07:45:23+00:00Z

Additive Manufacturing (AM) of metals is a processing technology that has significantly matured over the last decade. For liquid propellant rocket engines, the advantages of AM for replacing conventional manufacturing of complicated and expensive metallic components and assemblies are very attractive. AM can significantly reduce hardware cost, shorten fabrication schedules, increase reliability by reducing the number of joints, and improve hardware performance by allowing fabrication of designs not feasible by conventional means. The NASA Marshall Space Flight Center (MSFC) has been involved with various forms of metallic additive manufacturing for use in liquid rocket engine component design, development, and testing since 2010. The AM technique most often used at the NASA MSFC has been powder-bed fusion or selective laser melting (SLM), although other techniques including laser directed energy deposition (DED), arc-based deposition, and laser-wire cladding techniques have also been used to develop several components. The purpose of this paper is to discuss the various internal programs at the NASA MSFC using AM to develop combustion devices hardware. To date at the NASA MSFC, combustion devices component hardware ranging in size from 100 lbf to 35,000 lbf have been designed and manufactured using SLM and deposition-based AM processes, and many of these pieces have been hot-fire tested. Combustion devices component hardware have included thrust chamber injectors, injector components such as faceplates, regeneratively-cooled combustion chambers, regeneratively-cooled nozzles, gas generator and preburner hardware, and augmented spark igniters. Ongoing and future developments for combustion devices have also included design of components sized for boost-class engines. Several design and hot-fire test iterations have been completed on these subscale and larger scale components, and a summary of these results will be presented as well.

This paper presents the fabrication, testing and analysis of the MIT micro-rocket combustion chamber structure. The structure is a rocket chamber with a closed throat and pressure feeds. It is fabricated with the same processes as the full micro-rocket builds. Mechanical test results are correlated with inspection of failed devices for flaws and with finite element modeling of the test condition. This analysis provides recommendations for modifications to improve the strength of the micro-rocket chamber. These recommendations are discussed in the context of the current status of the micro-rocket.

No abstract