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All sourcesNASA NTRSDOE OSTI
NASA NTRS1979

Symposium /International/ on Combustion, 17th, Leeds University, Leeds, England, August 20-25, 1978, Proceedings

The Symposium focused on deflagration to detonation transition, coal combustion, turbulent-combustion interactions, kinetics, furnace combustion, inhibition and ignition, flame structure and chemistry, combustion studies, measurement techniques, fire and explosion, engine combustion, soot, and propellants and explosives. Papers were presented on numerical modeling of the deflagration-to-detonation transition, the interaction between turbulence and combustion, turbulent flame propagation in premixed gases, spray evaporation in recirculating flow, dissociation of nitric oxide in shock waves, pollutant emissions from partially mixed turbulent flames, energy transfer and quenching rates of laser-pumped electronically excited alkalis in flames, a study of flammability limits using counterflow flames, the unified theory of explosions with fuel consumption, and the dynamics and radiant intensity of large hydrogen flames.

Source record↗
NASA NTRS1989

Hybrid propulsion technology program. Volume 2: Technology definition package

A concept design study was performed to configure two sizes of hybrid boosters; one which duplicates the advanced shuttle rocket motor vacuum thrust time curve and a smaller, quarter thrust level booster. Two sizes of hybrid boosters were configured for either pump-fed or pressure-fed oxygen feed systems. Performance analyses show improved payload capability relative to a solid propellant booster. Size optimization and fuel safety considerations resulted in a 4.57 m (180 inch) diameter large booster with an inert hydrocarbon fuel. The preferred diameter for the quarter thrust level booster is 2.53 m (96 inches). The demonstration plan would culminate with test firings of a 3.05 m (120 inch) diameter hybrid booster.

Jensen, Gordon E.↗
DOE OSTI2021

Final Report on Aerosol Pretreatment Technology Performance and Benchmarking

Solvent-based post-combustion CO 2 capture (PCC) technology remains one of the leading methods to combat global CO 2 emissions produced from large-scale coal-fired power production. Advanced solventbased PCC technology has made significant improvements in design and performance that reduce capital and operating costs to enable its commercial use. Key to low cost, manageable logistics, and environmentally safe operation of solvent-based PCC technology are minimal solvent losses from the process through the treated gas stream exiting PCC plant absorbers. High flue gas aerosol particle concentrations (>10 5 particles/cm 3 ) for particles in the range of 70-200 nm have been shown to cause significant amine solvent losses for solvent-based PCC processes through several mechanisms including absorption of solvent and water into growing aerosol particles. Flue gas aerosol pretreatment technology is the only realistic and economically attractive method to reduce very high aerosol particle concentrations (>10 7 particles/cm 3 ) to enable solvent-based PCC for existing power plants lacking sufficient particle removal systems, such as baghouses. The overall goal of this project was to design, construct, independently test, and evaluate three flue gas aerosol pretreatment technologies identified to significantly reduce high aerosol particle concentrations (>10 7 particles/cm 3 ) in the 70-200 nm particle size range: (1) a novel, high-velocity water injection spray concept developed by RWE and tested by Linde, (2) an innovative electrostatic precipitator (ESP) device with optimized operating conditions developed by Washington University in St. Louis (WUSTL), and (3) a non-regenerative sorbent-based filter technology developed by InnoSepra for SO x and NO x removal from coal-fired power plant flue gas. Each technology has been validated with tests on 500-1000 scfm of actual coal-fired flue gas and evaluated in terms of particle removal efficiency (%), cost competitiveness, and environmental impact. Aerosol measurements were performed upstream and downstream of each aerosol reduction unit during independent testing of each technology using advanced instrumentation and analytical methods provided by WUSTL. To perform aerosol measurements, isokinetic probes were inserted into flanged pipes attached to the flue gas piping, and a small suction pump was used to sample gas containing aerosol particles. Aerosol particle number concentrations (# particles/cm 3 ) and size distributions were then measured using a scanning mobility particle sizer (SMPS, TSI Inc.) for very fine particles (<1,000 nm) and a particle counter manufactured by GRIMM for particles larger than 1,000 nm. This report summarizes the aerosol removal performance results from pilot scale testing of each technology. Performance results have been benchmarked against pre-defined targets and other flue gas aerosol pretreatment technologies with documented performance. Linde Gas North America LLC has been the prime contractor to DOE responsible for overall project management and provided the design for the high-velocity water spray-based aerosol removal technology based on a design concept developed by the German utility company RWE.

20 FOSSIL-FUELED POWER PLANTS↗
NASA NTRS2012

Benefits of a Single-Person Spacecraft for Weightless Operations

Historically, less than 20 percent of crew time related to extravehicular activity (EVA) is spent on productive external work.1 A single-person spacecraft with 90 percent efficiency provides productive new capabilities for maintaining the International Space Station (ISS), exploring asteroids, and servicing telescopes or satellites. With suits, going outside to inspect, service or repair a spacecraft is time-consuming, requiring pre-breathe time, donning a fitted space suit, and pumping down an airlock. For ISS, this is between 12.5 and 16 hours for each EVA, not including translation and work-site set up. The work is physically demanding requiring a day of rest between EVAs and often results in suit-induced trauma with frequent injury to astronauts fingers2. For maximum mobility, suits use a low pressure, pure oxygen atmosphere. This represents a fire hazard and requires pre-breathing to reduce the risk of decompression sickness (bends). With virtually no gravity, humans exploring asteroids cannot use legs for walking. The Manned Maneuvering Unit offers a propulsive alternative however it is no longer in NASA s flight inventory. FlexCraft is a single person spacecraft operating at the same cabin atmosphere as its host so there is no risk of the bends and no pre-breathing. This allows rapid, any-time access to space for repeated short or long EVAs by different astronauts. Integrated propulsion eliminates hand-over-hand translation or having another crew member operate the robotic arm. The one-size-fits-all FlexCraft interior eliminates the suit part inventory and crew time required to fit all astronauts. With a shirtsleeve cockpit, conventional displays and controls are used and because the work is not strenuous no rest days are required. Furthermore, there is no need for hand tools because manipulators are equipped with force multiplying end-effectors that can deliver the precise torque for the job.

Griffin, Brand N.↗
NASA NTRS2010

J-2X Turbopump Cavitation Diagnostics

The J-2X is the upper stage engine currently being designed by Pratt & Whitney Rocketdyne (PWR) for the Ares I Crew Launch Vehicle (CLV). Propellant supply requirements for the J-2X are defined by the Ares Upper Stage to J-2X Interface Control Document (ICD). Supply conditions outside ICD defined start or run boxes can induce turbopump cavitation leading to interruption of J-2X propellant flow during hot fire operation. In severe cases, cavitation can lead to uncontained engine failure with the potential to cause a vehicle catastrophic event. Turbopump and engine system performance models supported by system design information and test data are required to predict existence, severity, and consequences of a cavitation event. A cavitation model for each of the J-2X fuel and oxidizer turbopumps was developed using data from pump water flow test facilities at Pratt & Whitney Rocketdyne (PWR) and Marshall Space Flight Center (MSFC) together with data from Powerpack 1A testing at Stennis Space Center (SSC) and from heritage systems. These component models were implemented within the PWR J-2X Real Time Model (RTM) to provide a foundation for predicting system level effects following turbopump cavitation. The RTM serves as a general failure simulation platform supporting estimation of J-2X redline system effectiveness. A study to compare cavitation induced conditions with component level structural limit thresholds throughout the engine was performed using the RTM. Results provided insight into system level turbopump cavitation effects and redline system effectiveness in preventing structural limit violations. A need to better understand structural limits and redline system failure mitigation potential in the event of fuel side cavitation was indicated. This paper examines study results, efforts to mature J-2X turbopump cavitation models and structural limits, and issues with engine redline detection of cavitation and the use of vehicle-side abort triggers to augment the engine redline system.

Santi, I. Michael↗
DOE OSTI2024

Field Study of 120-volt Heat Pump Water Heaters in the Big Easy

To meet long-term goals for reducing carbon emissions and lessen the impacts of climate change, the U.S. plans to decarbonize its building stock, which includes the replacement of fossil fuel-burning end uses with energy efficient electric alternatives. As part of this strategy, the replacement of fossil fuel water heaters with heat pump water heaters (HPWH) has the potential to avoid substantial carbon emissions. An estimated 10-15 million single-family homes with fossil fuel water heaters do not have the electrical panel capacity to install a 240-volt, 30-amp HPWH. For these homes, a technology recently introduced to the market, the 120-volt plug-in HPWH, can provide energy efficient electrification of water heating without an expensive panel and wiring retrofit. This paper presents the results of an ongoing 120-volt HPWH field study conducted in 17 homes in New Orleans, LA. Key installation scenarios are profiled for retrofitting from gas-fired water heaters to 120-volt HPWHs, accounting for space, air volume, air temperature, condensate drainage, and electrical. Each HPWH had its surrounding air temperature, relative humidity, inlet and outlet water temperature, flow, and energy consumption monitored. Using these data, hot water delivery and energy efficiency performance were analyzed based on home characteristics. Hot water run outs were investigated to understand how hot water usage, inlet water temperature, and surrounding air temperature impact the HPWH’s ability to meet load. From these results, best practices for 120-volt HPWH siting, sizing, and installation were developed. In addition, results are explored further to add insight for electrification policies and product development.

heat pump water heater, 120-volt, Residential buil↗
NASA NTRS2010

A History of Welding on the Space Shuttle Main Engine (1975 to 2010)

The Space Shuttle Main Engine (SSME) is a high performance, throttleable, liquid hydrogen fueled rocket engine. High thrust and specific impulse (Isp) are achieved through a staged combustion engine cycle, combined with high combustion pressure (approx.3000psi) generated by the two-stage pump and combustion process. The SSME is continuously throttleable from 67% to 109% of design thrust level. The design criteria for this engine maximize performance and weight, resulting in a 7,800 pound rocket engine that produces over a half million pounds of thrust in vacuum with a specific impulse of 452/sec. It is the most reliable rocket engine in the world, accumulating over one million seconds of hot-fire time and achieving 100% flight success in the Space Shuttle program. A rocket engine with the unique combination of high reliability, performance, and reusability comes at the expense of manufacturing simplicity. Several innovative design features and fabrication techniques are unique to this engine. This is as true for welding as any other manufacturing process. For many of the weld joints it seemed mean cheating physics and metallurgy to meet the requirements. This paper will present a history of the welding used to produce the world s highest performance throttleable rocket engine.

Zimmerman, Frank R.↗
DOE OSTI2024

Investigation of powertrain system decarbonization using electrically assisted turbocharging and hybridization in off-road vehicles

In recent years, the causes, effects, and existential threat of a global anthropogenic climate shift have drawn significant attention and stimulated mitigation efforts from all genres of scientific, political, and industrial bodies. The greenhouse effect and the detrimental environmental impact of excess greenhouse gas (GHG) emissions, like carbon dioxide, is well studied and targeted as the primary culprit for aversive action. However, some fields exist where reducing GHG emissions is met by considerable challenges. One such field is the production and operations of off-road vehicles. Applications of these vehicles are highly diverse and are often characterized by rugged, especially transient, and power intensive duty cycles that make one-fits-all vehicle configuration solutions impractical. This research surveys existing literature to identify the modern technologies packages and challenges that face development and configuration of powertrain systems which currently navigate the regulated emissions and unique duty cycle requirements of this market space. A novel powertrain concept is proposed and evaluated with respect to the pinnacle objectives of load performance improvement, criteria pollutant reduction, and reducing the life cycle GHG intensity of its operation. A sure-fire path to vehicular GHG reduction is through improving the fuel efficiency of internal combustion engines (ICEs), an entrenched component of the off-road vehicle sector. Unfortunately, this is more easily said than done. An approach that has proven successful in this endeavor is engine downsizing and turbocharging, where a larger engine is replaced with a smaller one with added air system boosting via turbocharger to reduce frictional and pumping losses while also enabling access to additional fuel energy. However, practical realization of these potential benefits is often impeded by the transient response capability of the smaller engine across the operating space. For this reason, downsized ICE powertrains have turned to electrified forced induction systems (EFISs) for a decoupling of exhaust energy and engine speed from boost capability. Platformed on 48V hybrid technology, these systems introduce the need for more sophisticated controls around the engine gas-exchange process for the management of boost performance and exhaust gas emissions. On this account, simulation studies utilizing a GT-POWER model of a turbocharged 4.5 L engine outfit with an EFIS using an electrically driven compressor (eBooster®) are conducted to provide insight into the performance of this air handling architecture on an off-road engine. The results show that improvement in transient torque response time in sync with reduction in engine-out soot and NOx emissions are possible with an engine recalibration that leverages the transient air-fuel ratio authority of the eBooster®. Benefits are further demonstrated when duty cycle simulations of the 48V mild-hybrid engine concept are exercised, showing an acute decrease in cumulative fuel usage and soot production. Powertrain hybridization is another technological pathway achieving pronounced GHG reduction successes in modern on-road vehicle applications through integration of Li-ion battery technology. In the off-road vehicle segment, a review of available literature concludes that hybridized architectures are present but generally lack the depth of technologies that have both high specific energy and power capabilities, and thus are limited in their inclusion of Li-ion batteries for ICE assistance and enhanced energy storage capability. Therefore, building on the mild-hybrid engine results, the downsized and eBoosted engine concept was integrated into a larger high-voltage battery-hybrid series-electric powertrain system. Hybrid powertrain parameter sensitivity studies were carried out in a numerical charge-sustaining framework, providing novel insights into power flows between the battery and the engine and how their respective capabilities and operation contribute to GHG and criteria pollutant emissions of diverse duty cycles. Application of supervisory power management introduced robustness into the power sourcing and battery SOC control process and showed that optimum specification of battery properties can yield synergies between GHG emissions and battery energy capacity. Furthermore, examination of recent literature on Li-ion batteries has shown that pack manufacturing is a highly energy intensive process, thereby producing considerable quantities of GHGs that scale with energy storage capacity. Also scaling with a battery’s energy storage capacity is its investment cost. In consideration of these factors, an inclusive technoeconomic and GHG life cycle analysis is conducted. This analysis systematically compares the carbon footprint and total cost of ownership associated with the proposed hybrid powertrain concept to reference and alternative powertrain configurations, facilitating a thorough evaluation of the decarbonization effectiveness and economic viability.

99 GENERAL AND MISCELLANEOUS↗
NASA NTRS1999

Materials Selection and Their Characteristics as Used in Rapid Prototyping

NASA's Marshall Space Flight Center (MSFC) conducted a program to evaluate six technologies used in Rapid Prototyping (RP) to produce investment casting patterns. In this paper, RP refers to the collective additive fabrication technologies known as Solid Free-Form Fabrication. Such technologies are being used with increasing frequency in manufacturing applications, due in part to their rapidly expanding capabilities to fabricate models from many types of materials. This study used ABS plastic, polycarbonate, TrueForm PM6, epoxy resin, paper, starch, and wax. The baseline model was a semi-complex prototype fuel pump housing, intended for use in the X-33 reusable launch vehicle. All models were shelled in a production- grade colloidal silica ceramic. Primary coats were zircon-base flour with zircon backup, while secondary coats were silica grains with a tabular alumina backup. Each model was shelled in an identical manner, using the same atmospheric conditions and drying times, as well as the same number of layers. Bake-outs and firing cycles were consistent with the leach ability of each material. Preheat and bath temperatures were also kept consistent. All molds were cast in vacuum using a hydrogen-resistant superalloy (NASA- 23) that was developed in-house. The final technical evaluation included detailed measurements of the model and the final casting, in order to determine any dimensional changes caused by different pattern materials, as well as documentation of all defects and any obvious refractory/model reactions. Prototype production costs were estimated for each method and taken into consideration during trade-off analysis.

Cooper, K.↗
NASA NTRS1960

Ignition of a Combustible Atmosphere by Incandescent Carbon Wear Particles

A study was made to determine whether carbon wear particles from carbon elements in sliding contact with a metal surface were sufficiently hot to cause ignition of a combustible atmosphere. In some machinery, electric potential differences and currents may appear at the carbon-metal interface. For this reason the effect of these voltages and currents on the ability of carbon wear particles to cause ignition was evaluated. The test specimens used in the investigation were carbon vanes taken from a fuel pump and flat 21-inch-diameter 2 metal disks (440-C stainless steel) representing the pump housing. During each experiment a vane was loaded against a disk with a 0.5-pound force, and the disk was rotated to give a surface speed of 3140 feet per minute. The chamber of the apparatus that housed the vane and the disk was filled with a combustible mixture of air and propane. Various voltages and amperages were applied across the vane-disk interface. Experiments were conducted at temperatures of 75, 350, 400, and 450 F. Fires were produced by incandescent carbon wear particles obtained at conditions of electric potential as low as 106 volts and 0.3 ampere at 400 F. Ignitions were obtained only with carbon wear particles produced with an electric potential across the carbon-vane-disk interface. No ignitions were obtained with carbon wear particles produced in the absence of this potential; also, the potential difference produced no ignitions in the absence of carbon wear particles. A film supplement showing ignition by incandescent wear particles is available.

Buckley, Donald H.↗
NASA NTRS2011

Application of Probabilistic Methods to Assess Risk Due to Resonance in the Design of J-2X Rocket Engine Turbine Blades

The LOX-Hydrogen J-2X Rocket Engine, which is proposed for use as an upper-stage engine for numerous earth-to-orbit and heavy lift launch vehicle architectures, is presently in the design phase and will move shortly to the initial development test phase. Analysis of the design has revealed numerous potential resonance issues with hardware in the turbomachinery turbine-side flow-path. The analysis of the fuel pump turbine blades requires particular care because resonant failure of the blades, which are rotating in excess of 30,000 revolutions/minutes (RPM), could be catastrophic for the engine and the entire launch vehicle. This paper describes a series of probabilistic analyses performed to assess the risk of failure of the turbine blades due to resonant vibration during past and present test series. Some significant results are that the probability of failure during a single complete engine hot-fire test is low (~1%) because of the small likelihood of resonance, but that the probability increases to around 30% for a more focused turbomachinery-only test because all speeds will be ramped through and there is a greater likelihood of dwelling at more speeds. These risk calculations have been invaluable for use by program management in deciding if risk-reduction methods such as dampers are necessary immediately or if the test can be performed before the risk-reduction hardware is ready.

Brown, Andrew M.↗
NASA NTRS1992

Lidar Measurements of Metallic Species in Mesopause Region

Sodium density distributions were measured over more than 20 nights since Jul. 1991 at Tokyo Metropolitan University. Our sodium lidar specifications are shown. The lidar consists of a dye laser that is pumped by a pulsed Nd:YAG laser. The dye laser is tuned to the D2 resonant absorption line of Na at the wavelength of 589 nm. A small fraction of the output beam is directed to a hollow-cathode lamp. The precise tuning of the Na D2 line is accomplished by monitoring the signal obtained from the hollow-cathode lamp and controlling the grating in the dye laser oscillator cavity. The raw data for each profile were obtained by firing the laser 5000 times for approximately 8.5 minutes. We have observed the sporadic sodium layers three times during the observation period. Recently, lidar measurements of iron atoms were also performed. Other aspects of this investigation are presented.

Nagasawa, Chikao↗
NASA NTRS1985

Flow dynamic environment data base development for the SSME

The fluid flow-induced vibration of the Space Shuttle main engine (SSME) components are being studied with a view to correlating the frequency characteristics of the pressure fluctuations in a rocket engine to its operating conditions and geometry. An overview of the data base development for SSME test firing results and the interactive computer software used to access, retrieve, and plot or print the results selectively for given thrust levels, engine numbers, etc., is presented. The various statistical methods available in the computer code for data analysis are discussed. Plots of test data, nondimensionalized using parameters such as fluid flow velocities, densities, and pressures, are presented. Results are compared with those available in the literature. Correlations between the resonant peaks observed at higher frequencies in power spectral density plots with pump geometry and operating conditions are discussed. An overview of the status of the investigation is presented and future directions are discussed.

Sundaram, C. V.↗
NASA NTRS2006

Liquid Bismuth Feed System for Electric Propulsion

Operation of Hall thrusters with bismuth propellant has been shown to be a promising path toward high-power, high-performance, long-lifetime electric propulsion for spaceflight missions. For example, the VHITAL project aims td accurately, experimentally assess the performance characteristics of 10 kW-class bismuth-fed Hall thrusters - in order to validate earlier results and resuscitate a promising technology that has been relatively dormant for about two decades. A critical element of these tests will be the precise metering of propellant to the thruster, since performance cannot be accurately assessed without an accurate accounting of mass flow rate. Earlier work used a pre/post-test propellant weighing scheme that did not provide any real-time measurement of mass flow rate while the thruster was firing, and makes subsequent performance calculations difficult. The motivation of the present work was to develop a precision liquid bismuth Propellant Management System (PMS) that provides real-time propellant mass flow rate measurement and control, enabling accurate thruster performance measurements. Additionally, our approach emphasizes the development of new liquid metal flow control components and, hence, will establish a basis for the future development of components for application in spaceflight. The design of various critical components in a bismuth PMS are described - reservoir, electromagnetic pump, hotspot flow sensor, and automated control system. Particular emphasis is given to material selection and high-temperature sealing techniques. Open loop calibration test results are reported, which validate the systems capability to deliver bismuth at mass flow rates ranging from 10 to 100 mg/sec with an uncertainty of less than +/- 5%. Results of integrated vaporizer/liquid PMS tests demonstrate all of the necessary elements of a complete bismuth feed system for electric propulsion.

Markusic, T. E.↗
NASA NTRS1976

Liquid-metal magnetohydrodynamic system evaluation

The present study emphasizes a direct coal-fired design using a bubbly two-component flow of sodium and argon in the MHD generator and a Rankine steam-bottoming plant. Two basic cycles were studied, corresponding to argon temperatures of 922 and 1089 K at the duct inlet. The MHD duct system consisted of multiple ducts arranged in clusters and separated by iron magnet pole pieces. The ducts, each with an output of about 100 MW, were parallel to the flow, but were connected in series electrically to provide a higher MHD voltage. With channel efficiencies of 80%, a pump efficiency of 90%, and a 45% efficient steam-bottoming plant, the overall efficiency of the 1089 K liquid-metal MHD power plant was 43%.

Holman, R. R.↗
NASA NTRS2022

Design of Hypervelocity Flow Generator (HFG) and Its Diagnostics

Ground facilities for hypersonic research are a key element for successful development of aerodynamically proven hypervelocity vehicles. Design concepts and diagnostics of a hypervelocity flow generator (HFG) were made as a test platform for hypersonic/hypervelocity spacecraft models at the NASA Langley Research Center. The HFG is a hypersonic flow field generator using optically heated gas which is blown into an 80 cu m vacuum chamber. The vacuum chamber is kept at a stable vacuum pressure with a combination of three large vacuum pumps, while the HFG is in the test mode. The HFG provides a relatively small test section with approximately a 20 cm window. This facility was designed to generate 2.45 km/sec of flow speed, and potentially generate a continuous flow with the nozzle and vacuum system. The window in test section provides a direct view of the shock wave around a model in order to measure temperature, pressure, and density profile within the shock layer. One of the key test goals under this project is to understand why the emission spectra from the standing shockwave plasma predicted by Lora-Loran codes are significantly different from the measured emission spectra from the Flight Investigation Reentry Environment (FIRE II) Flight. The correct estimate of the thermal loading on the leading edge of hypersonic vehicles greatly affects the aerodynamic design, the material selection for the vehicle, and the cooling requirement and can be obtained by the precise modeling of emission spectra from shockwave plasma. However, the estimation of thermal loading is not an easy task due to complex non-equilibrium radiative process within high temperature shock layers that still falls into a category of cold plasma. Direct flight experiments are the most desirable, but not a cost-effective approach. Analysis by computational fluid dynamics (CFD) offers many test flexibilities. However, the CFD codes must be fully tested and validated with experimental data before the codes are effectively used for practical design. Large discrepancies between experiments and codes appear in hypersonic/hypervelocity flow regimes at high altitudes of 60 km ~ 90 km. This HFG facility offers some important parameters for CFD code validation, such as collision cross-sections, relaxation times, reaction rate coefficients and transportation coefficients. The HFG test facility is based on the ejection flow of high temperature gas heated over to 3500 K through a nozzle. The tungsten gas chamber of the HFG is heated up to a desired temperature by a 60 kW optical power beam source. This system consists of an optical power source, a thermal chamber, an expansion nozzle, a test section, and an 80 cu m vacuum tank. 60 kW optical input power is obtained from the 150 kW Vortek arc lamp system (by Vortek Industries, Vancouver, Canada). This optical beam is focused to heat the gas chamber within which a flow media is heated. The maximum achievable temperature of the flow medium reached approximately 3500 K or even higher but is limited by the melting point of the chamber material used. The exhaust velocity through the nozzle was determined by the stagnation temperature and the molecular weight of the working medium at the test section. To provide design parameters, a NASA Chemical Equilibrium with Application (CEA) computer program is used for the simulation of aerothermal data. This CEA program can calculate chemical equilibrium and properties of complex mixtures using shock tube parameters. For nitrogen gas at 2666 K stagnation temperature, the maximum achievable velocity at the test section is approximately 2.45 km/sec which is within the range of the thermal velocity of 8000 K shock layer. Based on the calculation through the CEA program, the design parameters of a HFG were determined and implemented for the test section that includes an expansion nozzle. The installed test section of the experimental facility can sustain a condition of a re-entry vehicle from the Space at an altitude of 60-90 km. After installation of the HFG, the system was fully tested and its operational parameters were measured. An 80 cu m vacuum chamber of HFG was set at 1 torr level to keep a stable downstream condition. The pumping time to reach the minimum vacuum pressure (~ 1 torr) at the test chamber from atmospheric pressure was approximately an hour using Kinney (MBV-14000/MB -1600/KT-300) and Stokes (Model 1772 and 412) pumps. Such a setup condition allowed a continued stable operation of the HFG experiment with flowrates through 1-, 3-, or 5-mm diameters of nozzle throat. The flow characteristics of the HFG for various operating conditions were performed using a focused Schlieren method. At 0.8 torr chamber pressure, a barrel shock was observed at the test section. With a sphere obstacle of 12 mm diameter in the flow, a bow shock (~ 2-mm thickness) was observed by a focused Schlieren visualization method.

Hypervelocity flow↗
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 NTRS1993

Test Results of the RS-44 Integrated Component Evaluator Liquid Oxygen/Hydrogen Rocket Engine

An advanced LOX/LH2 expander cycle rocket engine, producing 15,000 lbf thrust for Orbital Transfer Vehicle missions, was tested to determine ignition, transition, and main stage characteristics. Detail design and fabrication of the pump fed RS44 integrated component evaluator (ICE) was accomplished using company discretionary resources and was tested under this contracted effort. Successful demonstrations were completed to about the 50 percent fuel turbopump power level (87,000 RPM), but during this last test, a high pressure fuel turbopump (HPFTP) bearing failed curtailing the test program. No other hardware were affected by the HPFTP premature shutdown. The ICE operations matched well with the predicted start transient simulations. The tests demonstrated the feasibility of a high performance advanced expander cycle engine. All engine components operated nominally, except for the HPFTP, during the engine hot-fire tests. A failure investigation was completed using company discretionary resources.

Sutton, R. F.↗