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NASA NTRS2009

100-lbf LO2/CH4 RCS Thruster Testing and Validation

100 pound thrust liquid Oxygen-Methane thruster sized for RCS (Reaction Control System) applications. Innovative Design Characteristics include: a) Simple compact design with minimal part count; b) Gaseous or Liquid propellant operation; c) Affordable and Reusable; d) Greater flexibility than existing systems; e) Part of NASA'S study of "Green Propellants." Hot-fire testing validated performance and functionality of thruster. Thruster's dependence on mixture ratio has been evaluated. Data has been used to calculate performance parameters such as thrust and Isp. Data has been compared with previous test results to verify reliability and repeatability. Thruster was found to have an Isp of 131 s and 82 lbf thrust at a mixture ratio of 1.62.

Barnes, Frank↗
NASA NTRS2003

Design and Testing of Non-Toxic RCS Thrusters for Second Generation Reusable Launch Vehicle

The current NASA Space Shuttle auxiliary propulsion system utilizes nitrogen tetroxide (NTO) and monomethylhydrazine (MMH), hypergolic propellants. This use of these propellants has resulted in high levels of maintenance and precautions that contribute to costly launch operations. By employing alternate propellant combinations, those less toxic to humans, the hazards and time required between missions can be significantly reduced. Use of alternate propellants can thereby increase the efficiency and lower the cost in launch operations. In support of NASA's Space Launch Initiative (SLI), TRW proposed a three-phase project structured to significantly increase the technology readiness of a high-performance reaction control subsystem (RCS) thruster using non-toxic propellant for an operationally efficient and reusable auxiliary propulsion system (APS). The project enables the development of an integrated primary/vernier thruster capable of providing dual-thrust levels of both 1000-lbf class thrust and 25-lbf thrust. The intent of the project is to reduce the risk associated with the development of an improved RCS flight design that meets the primary NASA objectives of improved safety and reliability while reducing systems operations and maintenance costs. TRW proposed two non-toxic auxiliary propulsion engine designs, one using liquid oxygen and liquid hydrogen and the other using liquid oxygen and liquid ethanol, as candidates to meet the goals of reliability and affordability at the RCS level. Both of these propellant combinations offer the advantage of a safe environment for maintenance, while at the same time providing adequate to excellent performance for a conventional liquid propulsion systems. The key enabling technology incorporated in both TRW thrusters is the coaxial liquid on liquid pintle injector. This paper will concentrate on only the design and testing of one of the thrusters, the liquid oxygen (LOX) and liquid hydrogen (LH2) thruster. The LOX/LH2 thruster design includes a LOX-centered pintle injector, consisting of two rows of slots that create a radial spoke spray pattern in the combustion chamber. The main fuel injector creates a continuous sheet of LH2 originating upstream of the LOX pintle injector. The two propellants impinge at the pintle slots, where the resulting momentum ratio and spray pattern determines the combustion efficiency and thermal effects on the hardware. Another enabling technology used in the design of this thruster is fuel film cooling through a duct, lining the inner wall of the combustion chamber barrel section. The duct is also acts as a secondary fuel injection point. The variation in the amount of LH2 used for the duct allows for adjustments in the cooling capacity for the thruster. The Non-Toxic LOX-LH2 RCS Workhorse Thruster was tested at the NASA Marshall Space Flight Center's Test Stand 500. Hot-fire tests were conducted between March 08, 2002 and April 05, 2002. All testing during the program base period were performed at sea-level conditions. During the test program, 7 configurations were tested, including 2 combustion chambers, 3 LOX injector pintle tips, and 4 LH2 injector stroke settings. The operating conditions that were surveyed varied thrust levels, mixture ratio and LH2 duct cooling flow. The copper heat sink chamber was used for 16 burns, each burn lasting from 0.4 to 10 seconds, totaling 51.4 seconds, followed by Haynes chamber testing ranging from 0.9 to 120 seconds, totaling 300.9 seconds. The total accumulated burn time for the test program is 352.3 seconds. C* efficiency was calculated and found to be within expectable limits for most operating conditions. The temperature on the Haynes combustion chamber remained below established material limits, with the exception of one localized hot spot. The test results demonstrate that both the coaxial liquid-on-liquid pintle injector design and fuel duct concepts are viable for the intended application. The thruster head-e design maintained cryogenic injection temperatures while firing, which validates the concept for minimal heat soak back. By injecting fuel into the duct, the throat temperatures were manageable, yet the split of fuel through the cooling duct does not compromise the overall combustion efficiency, which indicates that, provided proper design refinement, such a concept can be applied to a high-performance version of the thruster. These hot fire tests demonstrate the robustness of the duct design concept and good capability to withstand off-nominal operating conditions without adversely impacting the thermal response of the engine, a key design feature for a cryogenic thruster.

Calvignac, Jacky↗
NASA NTRS2006

Visualization of a Capsule Entry Vehicle Reaction-Control System (RCS) Thruster

Planar laser-induced fluorescence (PLIF) was used to visualize the reaction control system (RCS) jet flow emanating from the aft-body of an Apollo-geometry capsule test article in the NASA Langley Research Center 31-Inch Mach 10 Air wind tunnel. The RCS jet was oriented normal to the aft surface of the model and had a nominal Mach number of 2.94. The composition of the jet gas by mass was 95% nitrogen (N2) and 5% nitric oxide (NO). The RCS jet flowrate varied between zero and 0.5 standard liters per minute and the angle of attack and tunnel stagnation pressure were also varied. PLIF was used to excite the NO molecules for flow visualization. These flow visualization images were processed to determine the trajectory and to quantify the flapping of the RCS jet. The spatial resolution of the jet trajectory measurement was about 1 mm and the single-shot precision of the measurement was estimated to be 0.02 mm in the far field of the jet plume. The jet flapping, measured by the standard deviation of the jet centerline position was as large as 0.9 mm, while the jet was 1.5-4 mm in diameter (full width at half maximum). Schlieren flow visualization images were obtained for comparison with the PLIF. Surface pressures were also measured and presented. Virtual Diagnostics Interface (VIDI) technology developed at NASA Langley was used to superimpose and visualize the data sets. The measurements demonstrate some of the capabilities of the PLIF method while providing a test case for computational fluid dynamics (CFD) validation.

Danehy, P. M.↗
NASA NTRS1973

Exhaust plume and contamination characteristics of a bipropellant (MMH/N2O4) RCS thruster

Results are presented for three recent tests in a series of thruster contamination experiments made in liquid helium-cooled environmental facility. The contaminating effects encountered on various materials, surfaces, and components, due to the exhaust products from a 5-pound thrust, bipropellant (MMH/N2O4) thruster are investigated. The angular distribution of plume effects around the periphery of the thruster established by transmittance changes of quartz samples over the wavelength range from 0.2 to 2.0 micrometer is studied, along with mass deposition rates at a specific location measured with a quartz crystal microbalance for three different experiments. Quadrupole mass spectrometer measurements of the exhaust products over the mass number range from 12 to 75; infrared transmittance measurements of contaminated samples for the wavelength range from 2.5 to 15 microns; and infrared transmittance measurements of residue from the thruster nozzle are also considered.

Spisz, E. W.↗
NASA NTRS1973

Plume mass flow and optical damage distributions for an MMH/N2O4 RCS thruster

The data obtained from two recent experiments conducted in a continuing series of experiments at the Lewis Research Center into the contamination characteristics of a 5-pound thrust MMH/N2O4 engine are presented. The primary objectives of these experiments were to establish the angular distribution of condensible exhaust products within the plume and the corresponding optical damage angular distribution of transmitting optical elements attributable to this contaminant. The plume mass flow distribution was measured by five quartz crystal microbalances (QCM's) located at the engine axis evaluation. The fifth QCM was located above the engine and 15 deg behind the nozzle exit plane. The optical damage was determined by ex-situ transmittance measurements for the wavelength range from 0.2 to 0.6 microns on 2.54 cm diameter fused silica discs also located at engine centerline elevation. Both the mass deposition and optical damage angular distributions followed the expected trend of decreasing deposition and damage as the angle between sensor or sample and the nozzle axis increased. A simple plume gas flow equation predicted the deposition distribution reasonably well for angles of up to 55 degrees. The optical damage measurements also indicated significant effects at large angles.

Spisz, E. W.↗
NASA NTRS2003

Design and Testing of Non-Toxic RCS Thrusters for Second Generation Reusable Launch Vehicle

Under NASA sponsorship, Northrop Grumman Space Technology (NGST) designed, built and tested two non-toxic, reaction control engines, one using liquid oxygen (LOX) and liquid hydrogen (LH2) and the other using liquid oxygen and ethanol. This paper presents the design and testing of the LOX/LH2 thruster. The two key enabling technologies are the coaxial liquid-on-liquid pintle injector and the fuelcooling duct. The workhorse thruster was hotfire tested at the NASA Marshall Space Flight Center Test Stand 500 in March and April of 2002. All tests were performed at sea-level conditions. During the test program, 7 configurations were tested, including 2 combustion chambers, 3 LOX injector pintle tips, and 4 LHp injector settings. The operating conditions surveyed were 70 to 100% thrust levels, mixture ratios from 3.27 to 4.29, and LH2 duct cooling from 18.0 to 25.5% fuel flow. The copper heat sink chamber was used for 16 burns, each burn lasting from 0.4 to 10 seconds, totaling 51.4 seconds, followed by Haynes chamber testing ranging from 0.9 to 120 seconds, totaling 300.9 seconds. The performance of the engine reached 95% C* efficiency. The temperature on the Haynes chamber remained well below established material limits, with the exception of one localized hot spot. These results demonstrate that both the coaxial liquid-on-liquid pintle injector design and fuel duct concepts are viable for the intended application. The thruster headend design maintained cryogenic injection temperatures while firing, which validates the selected injector design approach for minimal heat soak-back. Also, off -nominal operation without adversely impacting the thermal response of the engine showed the robustness of the duct design, a key design feature for this application. By injecting fuel into the duct, the throat temperatures are manageable, yet the split of fuel through the cooling duct does not compromise the overall combstion efficiency, which indicates that, provided proper design refinement, such a concept could be applied to a high-performance version of the thruster.

Calvignac, Jacky↗
NASA NTRS2004

Causes and Mitigation of Fuel Pilot Operated Valve Pilot Seal Extrusion in Space Shuttle Orbiter Primary RCS Thrusters

Extrusion of a polytetrafluoroethylene (PTFE) pilot seal located in the Space Shuttle Orbiter Primary Reaction Control Subsystem (PRCS) thruster fuel valve has been implicated in 68 ground and on-orbit fuel valve failures. A rash of six extrusion-related in-flight anomalies over a six-mission span from December 2001 to October 2002 led to heightened activity at various NASA centers, and the formation of a multidisciplinary team to solve the problem. Empirical and theoretical approaches were used. For example, thermomechanical analysis (TMA) and exposure tests showed that some extrusion is produced by thermal cycling; however, a review of thruster service histories did not reveal a strong link between thermal cycling and extrusion. Calculations showed that the amount of observed extrusion often exceeded the amount allowed by thermally-induced stress relief. Failure analysis of failed hardware also revealed the presence of fuel-oxidizer reaction product (FORP) inside the fuel valve pilot seal cavity, and differential scanning calorimetry (DSC) showed that the FORP was intimately associated with the pilot seal material. Component-level exposure tests showed that FORP of similar composition could be produced by adjacent oxidizer valve leakage in the absence of thruster firing. Specific gravity data showed that extruded fuel valve pilot seals were less dense than new pilot seals or oxidizer valve pilot seals, indicating permanent modification of the PTFE occurred during service. It is concluded that some thermally-induced extrusion is unavoidable; however, oxidizer leakage-induced extrusion is mostly avoidable and can be mitigated. Several engineering level mitigation strategies are discussed.

Waller, Jess M.↗
NASA NTRS2012

Plume-Free Stream Interaction Heating Effects During Orion Crew Module Reentry

During reentry of the Orion Crew Module (CM), vehicle attitude control will be performed by firing reaction control system (RCS) thrusters. Simulation of RCS plumes and their interaction with the oncoming flow has been difficult for the analysis community due to the large scarf angles of the RCS thrusters and the unsteady nature of the Orion capsule backshell environments. The model for the aerothermal database has thus relied on wind tunnel test data to capture the heating effects of thruster plume interactions with the freestream. These data are only valid for the continuum flow regime of the reentry trajectory. A Direct Simulation Monte Carlo (DSMC) analysis was performed to study the vehicle heating effects that result from the RCS thruster plume interaction with the oncoming freestream flow at high altitudes during Orion CM reentry. The study was performed with the DSMC Analysis Code (DAC). The inflow boundary conditions for the jets were obtained from Data Parallel Line Relaxation (DPLR) computational fluid dynamics (CFD) solutions. Simulations were performed for the roll, yaw, pitch-up and pitch-down jets at altitudes of 105 km, 125 km and 160 km as well as vacuum conditions. For comparison purposes (see Figure 1), the freestream conditions were based on previous DAC simulations performed without active RCS to populate the aerodynamic database for the Orion CM. Other inputs to the analysis included a constant Orbital reentry velocity of 7.5 km/s and angle of attack of 160 degrees. The results of the study showed that the interaction effects decrease quickly with increasing altitude. Also, jets with highly scarfed nozzles cause more severe heating compared to the nozzles with lower scarf angles. The difficulty of performing these simulations was based on the maximum number density and the ratio of number densities between the freestream and the plume for each simulation. The lowest altitude solutions required a substantial amount of computational resources (up to 1800 processors) to simulate approximately 2 billion molecules for the refined (adapted) solutions.

Marichalar, J.↗
NASA NTRSDate not supplied

Application of OpenFOAM to Plume Impingement in Space Environments

After 30 years of continuous human presence in low-earth orbit, NASA is returning to the moon and eventually will go to Mars. Travelling beyond low earth orbit requires NASA to learn how humans can live in Deep Space environments – beyond the protection of Earth’s magnetosphere and at distances from Earth that prevent a quick return in case of trouble. To this end, NASA is constructing the Lunar Gateway, an ISS-like space station to be put in orbit around the moon to act as a home base for Lunar exploration for NASA astronauts. The Gateway Lunar outpost will be built incrementally, via modules which will arrive at separate times and dock to the existing structure. The incremental addition of Gateway modules, and the docking of visiting vehicles, is achieved via a sequence of firings from the approaching body’s onboard reaction control system (RCS) thrusters to achieve the required approach trajectory. The typical hypergolic chemical RCS thrusters work by firing hot gases to produce adverse thrust and the needed change in velocity to safely finish the docking process. The exhaust gas from the RCS thrusters form plumes that expand into the vacuum of space and can impinge onto the outer surfaces of the Lunar Gateway, causing unwanted forces and moments, heat loads, sediment deposition, and in extreme cases, even surface erosion - all mechanisms that can damage the Lunar Gateway and must be minimized. Both permanent and visiting modules will have this RCS thruster exhaust impingement problem. This research aims to establish existing OpenFOAM solvers as a methodology for improving simulation techniques of rocket exhaust plume impingement in space environments. The flow structure of a plume in a space environment is complex; a plume that originates from a hypergolic chemical RCS thruster and expands into a vacuum will experience several regimes of rarefication. This range includes the continuum flow in the rocket nozzle through the fully rarefied free molecular flow further from the nozzle. The flow physics is different at these two extremes, and as such, the simulation approach for plumes is generally divided into a traditional computational fluid dynamics (CFD) simulation in and near the nozzle which is coupled to a subsequent direct simulation Monte Carlo (DSMC) simulation. At this time, the scope of this research is developing, verifying, and validating a method using existing solvers in the OpenFOAM framework for performing coupled CFD/DSMC calculations to determine the extent of plume impingement loading on generic space structures. This presentation will detail code-to-code comparisons between the hyStrath dsmcFoam+ solver, developed using OpenFOAM and available as open-source, and NASA’s in-house DSMC Analysis Code (DAC). Comparisons to several open-source publication findings using DAC [3,4] are presented, and advantages of using an OpenFOAM based solver are also discussed. The presentation concludes with a discussion of future work, and a plan for coupling the dsmcFoam+ solver with CFD simulations of chemical rocket engines for unified coupled plume simulation.

DSMC↗
NASA NTRS1991

Control of a varying thrust spacecraft for autonomous space rendezvous

Before the use of autonomous rendezvous will be allowed as a substitute for man-in-the-loop control, adequate safety and mission performance will have to be guaranteed. Most autopilots for autonomous rendezvous of spacecraft assume constant thrust reaction control system (RCS) thrusters. This assumption implies either true constant thrust RCS thrusters or thrusters whose thrust levels vary very slowly. The ongoing work described in this presentation examines the autonomous rendezvous problem when varying thrust RCS thrusters are inherent in the system equations of motion.

Dawson, Travis E., Jr.↗
NASA NTRS2011

Metallographic Preparation of Space Shuttle Reaction Control System Thruster Electron Beam Welds for Electron Backscatter Diffraction

A Space Shuttle Reaction Control System (RCS) thruster failed during a firing test at the NASA White Sands Test Facility (WSTF), Las Cruces, New Mexico. The firing test was being conducted to investigate a previous electrical malfunction. A number of cracks were found associated with the fuel closure plate/injector assembly (Fig 1). The firing test failure generated a flight constraint to the launch of STS-133. A team comprised of several NASA centers and other research institutes was assembled to investigate and determine the root cause of the failure. The JSC Materials Evaluation Laboratory was asked to compare and characterize the outboard circumferential electron beam (EB) weld between the fuel closure plate (Titanium 6Al-4V) and the injector (Niobium C-103 alloy) of four different RCS thrusters, including the failed RCS thruster. Several metallographic challenges in grinding/polishing, and particularly in etching were encountered because of the differences in hardness, ductility, and chemical resistance between the two alloys and the bimetallic weld. Segments from each thruster were sectioned from the outboard weld. The segments were hot-compression mounted using a conductive, carbon-filled epoxy. A grinding/polishing procedure for titanium alloys was used [1]. This procedure worked well on the titanium; but a thin, disturbed layer was visible on the niobium surface by means of polarized light. Once polished, each sample was micrographed using bright field, differential interference contrast optical microscopy, and scanning electron microscopy (SEM) using a backscatter electron (BSE) detector. No typical weld anomalies were observed in any of the cross sections. However, areas of large atomic contrast were clearly visible in the weld nugget, particularly along fusion line interfaces between the titanium and the niobium. This prompted the need to better understand the chemistry and microstructure of the weld (Fig 2). Energy Dispersive X-Ray Spectroscopy (EDS) was used to confirm the chemical composition of the variations in contrast in these areas. Niobium alloys generally require exposure to more aggressive chemical reagents than titanium alloys for etching because of niobium s chemical resistance; therefore, the titanium portion of the sample was etched first. A five second immersion in Kroll s reagent revealed a general microstructure on the titanium portion of the sample; however, the titanium heat affected zone closest to the weld, was over-etched due to higher concentrations of refined grains and an increase in eta-phase. The Kroll s etchant also revealed some microstructure in the weld nugget itself; the niobium portion of the sample remained unetched.

Martinez, James↗
NASA NTRS2007

Laboratory Reproduction and Failure Analysis of Cracked Orbiter Reaction Control System Niobium Thruster Injectors

A viewgraph presentation describing cracks and failure analysis of an orbiter reaction control system is shown. The topics include: 1) Endeavour STS-113 Landing; 2) RCS Thruster; 3) Thruster Cross-Section; 4) RCS Injector; 5) RCS Thruster, S/N 120l 6) Counterbore Cracks; 7) Relief Radius Cracks; 8) RCS Thruster Cracking History; 9) Thruster Manufacturing Timelines; 10) Laboratory Reproduction of Injector Cracking; 11) The Brownfield Specimen; 12) HF EtchantTests/Specimen Loading; 13) Specimen #3 HF + 600F; 14) Specimen #3 IG Fracture; 15) Specimen #5 HF + 600F; 16) Specimen #5 Popcorn ; 17) Specimen #5 Cleaned and Bent; 18) HF Exposure Test Matrix; 19) Krytox143AC Tests; 20) KrytoxTests/Specimen Loading; 21) Specimen #13 Krytox + 600F; and 22) KrytoxExposure Test Matrix.

Jacobs, Jeremy B.↗
NASA NTRS1977

Results of tests using a 0.0125-scale model (70-QT) of the space shuttle vehicle orbiter in the AEDC VKF tunnel B (IA22), volume 2

Tabulated data of an experimental investigation are presented which was conducted in the AEDC/VKF Tunnel B to obtain interaction effects of RCS thruster jet plumes on SSV aerodynamics during staging to simulate RTLS abort. Interaction effects of the orbiter RCS thruster jet plumes on the orbiter and ET aerodynamics were investigated. RCS thruster jet plumes were simulated using both air and a 15 percent argon 85 percent helium gas mixture. The ET angle of attack range was -40 to +25 deg at sideslip angles of 0, 3, and 6 degrees. Orbiter angle of attack was varied from -15 to +10 degrees at sideslip angles of 0 and 3 deg. External tank full scale separation distances simulated were 0 to 1400 in. axially; 0 to 54 in. laterally; and a range of -100 to 1000 in. vertically. Data were also obtained on the ET in the interference-free flow field. Quiescent (no tunnel flow) thruster plume interaction data were obtained on the orbiter and orbiter-ET combination. Tests were conducted at Mach number 6 and a Reynolds number of 0.86 million per foot.

Daileda, J. J.↗
NASA NTRS2018

Final Cassini Propulsion System In-Flight Characterization

Cassini plunged into Saturn’s atmosphere on September 15, 2017, the culmination of two decades of incredible planetary exploration. The Cassini propulsion system performed spectacularly during this lengthy mission, executing 360 propulsive maneuvers successfully. R-4D performance was impeccable during 183 main-engine burns, ranging from 1.3 seconds to 96 minutes in duration. Pressure transducer drift was negligible, latch valve back-pressure relief was nominal, and pressurization system performance was flawless, except for hard-seat prime regulator leakage early in the mission. System health checks based on helium mass calculations (helium mass budgets) demonstrated no discernible helium pressurant or propellant leakage. Premature degradation of prime MR-103H RCS thrusters was noted, but back-up RCS thrusters completed the mission successfully. Final usable propellant margins were about 25% for hydrazine and less than 1% for bipropellant.

Barber, Todd J.↗
NASA NTRS2006

JSC Materials Laboratory Reproduction and Failure Analysis of Cracked Orbiter Reaction Control System Niobium Thruster Injectors

In April 2004 a Space Shuttle Orbiter Reaction Control System (RCS) thruster was found to be cracked while undergoing a nozzle (niobium/C103 alloy) retrofit. As a failure resulting from an in-flight RCS thruster burn-through (initiated from a crack) could be catastrophic, an official Space Shuttle Program flight constraint was issued until flight safety could be adequately demonstrated. This paper describes the laboratory test program which was undertaken to reproduce the cracking in order to fully understand and bound the driving environments. The associated rationale developed to justify continued safe flight of the Orbiter RCS system is also described. The laboratory testing successfully reproduced the niobium cracking, and established specific bounding conditions necessary to cause cracking in the C103 thruster injectors. Each of the following conditions is necessary in combination together: 1) a mechanically disturbed / cold-worked free surface, 2) an externally applied sustained tensile stress near yield strength, 3) presence of fluorine-containing fluids on exposed tensile / cold-worked free surfaces, and 4) sustained exposure to temperatures greater than 400 F. As a result of this work, it was concluded that fluorine-containing materials (e.g. HF acid, Krytox , Brayco etc.) should be carefully controlled or altogether eliminated during processing of niobium and its alloys.

Castner, Willard L.↗
NASA NTRS2016

CryoCart Restoration and Vacuum Pipe Construction

Propulsion systems that utilize hypergolic propellants have been used to power space vehicles since the beginning of the space program. Liquid methane and oxygen propulsion systems have emerged as an alternative and have proven to be more environmentally friendly. The incorporation of liquid methane/liquid oxygen (LOX) into the propulsion system has demonstrated an increase in engine performance, as well as a reduction in the volume, size and complexity of the system. Consequently, reducing the total mass of the vehicle which is a crucial aspect that is considered when planning space missions to both the Moon and Mars [1]. Project Morpheus has made significant advancements in liquid oxygen/liquid methane propulsion system technologies by incorporating a LOX/methane propulsion system to a vertical test bed. The vehicle consisted of a 5,000 lb main engine and four 20 lb remote control system (RCS) engines that utilize liquid methane/LOX as its propellant [1]. The vehicle completed successful flight testing at Kennedy Space Center in 2014 which marked the completion of the Morpheus project. Subsequent projects utilizing Morpheus' vertical test bed have been developed to make further advancements. One of the subsequent projects consisted of the addition of a smaller 2,000 lb main engine and a cold helium heat exchanger which would make it possible for a pressurant tank systems to be send to Mars or the Moon by significantly decreasing the overall mass and volume of the pressurant tank. The hot fire tests of the integrated system with the smaller main engine and cold helium heat exchanger were successful at sea level, but further studies are being conducted to better understand how the vertical test bed will behave under thermal-vacuum conditions. For this reason, the integrated vehicle will be taken to Plum Brook to be tested in a chamber capable of simulating these conditions. To ensure that the vehicle will function properly under vacuum conditions, testing will be first completed at the component level. During this process, the igniter of the main engine and the RCS thrusters will be tested under a vacuum. To complete the testing of the components, the test setup first needed to be finalized. The CryoCart is being used to feed the propellants to the test article. The CryoCart is a movable test set-up that was developed in 2009 to provide a mobile platform for testing oxygen/methane systems with hot-fire capability up to 100 lbf. The CryoCart consists of three different systems: Oxygen, Methane, and liquid Nitrogen. The Oxygen and Methane systems are placed into two different carts while the liquid nitrogen system is mainly located in the methane cart. Over the years, the CryoCart has been utilized for different projects and has undergone deterioration. For this reason, a new phase has been developed to rebuild it to working conditions once again. During my internship, I was aiding in the construction and restoration of the CryoCart. In the initial stages of the process, I updated the fluid and electrical schematics for the oxygen, methane, and test article systems. The original CryoCart consisted of an electrical panel that utilized electromechanical relays and a terminal to drive the igniter power and signal, as well as the main fuel and oxygen valves. This electrical panel connected to the CryoCart through various wire harnesses that could be found exiting from the CryoCart. First, it was determined how these harnesses connected to the electromechanical relays so that they worked correctly. Once the electrical system was understood, an alternative for the electromechanical relays and the Molex connectors used throughout the system was sought since these components can often prove to be unreliable. Solid State relays and MIL connectors were purchased to serve as replacements. Upon arrival of the parts, crimping and wiring was completed to install the new solid state relays and MIL connectors. During the replacement of the relays and connectors, system checks of the electrical system were ran to ensure that the system was working correctly. While completing system checks, the pressure transducers that were not functioning properly were also replaced and any issue with the wiring or signal was addressed. Once the electrical components were replaced, the restoration of the fluid system began. Parts of the tubing in the CryoCart had to be rebuild and often consisted of sizing, cutting, bending, filing, and sanding the tubing to prepare it to be flared. Many components had to be proof-tested to bring their certifications up to date, and several components had to be replaced. Various flex hoses, valves, and fittings were send to the Clean Lab because they were new, dirty, or had gone through proof-testing. Once they arrived from the cleaning lab they had to be put back to the system and leak checks and functional tests were conducted. In the Nitrogen system, the copper tubing located in the Oxygen cart was rebuild and Aerogel insulation was added to this section. A new gaseous nitrogen system was added to the CryoCart to purge the vacuum tube which will serve as the test chamber. Once the CryoCart was completed, construction of parts of the vacuum tube began. A flange was manufactured with welded fittings to hold the line of the vacuum pump as well as some extra fittings which will serve as extra inlets used to introduce fluid lines to the vacuum tube. Stress analysis was ran in this flange to ensure that it would not fail under vacuum conditions. The fluid lines leading from the air side of the vacuum to the test article were also constructed and added to the mount that had already been manufactured. Three different sets of tubing were constructed to accommodate the seven different RCS thruster and the main engine igniter that are going to be tested. Full electrical system checks were completed to ensure that all the wire harnesses and valves were functioning. Upon the completion of the CryoCart and the vacuum tube, hot fire testing for the RCS thrusters and the main engine igniter are going to begin. During this time any issues encountered with the engines or igniter will be addressed to ensure that the components function under vacuum conditions. After successful completion of testing, the vertical test bed, Morpheus, will be rebuilt and prepared to be sent to Plum Brook. In Plum Brook, the vehicle will be tested in the thermal-vacuum chamber to demonstrate that integrated lox-methane propulsion system operation in space-like conditions. This internship has allowed me the opportunity to gain valuable hands on experience and to develop skills that will aid in my education as well as in the workforce, while at the same time helping me determine that I would like to further pursue a career in propulsion engineering.

Chaidez, Mariana↗
NASA NTRS2011

Advanced Development of a Compact 5-15 lbf Lox/Methane Thruster for an Integrated Reaction Control and Main Engine Propulsion System

This paper describes the advanced development and testing of a compact 5 to 15 lbf LOX/LCH4 thruster for a pressure-fed integrated main engine and RCS propulsion system to be used on a spacecraft "vertical" test bed (VTB). The ability of the RCS thruster and the main engine to operate off the same propellant supply in zero-g reduces mass and improves mission flexibility. This compact RCS engine incorporates several features to dramatically reduce mass and parts count, to ease manufacturing, and to maintain acceptable performance given that specific impulse (Isp) is not the driver. For example, radial injection holes placed on the chamber body for easier drilling, and high temperature Haynes 230 were selected for the chamber over other more expensive options. The valve inlets are rotatable before welding allowing different orientations for vehicle integration. In addition, the engine design effort selected a coil-on-plug ignition system which integrates a relay and coil with the plug electrode, and moves some exciter electronics to avionics driver board. The engine injector design has small dribble volumes to target minimum pulse widths of 20 msec. and an efficient minimum impulse bit of less than 0.05 lbf-sec. The propellants, oxygen and methane, were chosen because together they are a non-toxic, Mars-forward, high density, space storable, and high performance propellant combination that is capable of pressure-fed and pump-fed configurations and integration with life support and power subsystems. This paper will present the results of the advanced development testing to date of the RCS thruster and the integration with a vehicle propulsion system.

Hurlbert, Eric A.↗
NASA NTRS1973

Aerothermodynamics of the Space Shuttle reaction control system.

Based on experimental data, the aerothermodynamic surface environment for the Space Shuttle Orbiter reaction control system (RCS) thrusters during entry was established. Using this environment the required thermal protection for the RCS thruster nozzles and surrounding vehicle surfaces on the orbiter nose was determined. Wind tunnel test data was obtained at Mach 8 to measure the heat transfer and pressure distribution in separated flow due to deep, three-dimensional cavities and lateral jet interference using helium and air (hot and cold) injection from a supersonic nozzle. Results of the analysis and correlation of the test data are presented.

Brevig, O.↗