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At least 415 records · Page 23
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NASA NTRSDate not supplied

Three Dimensional Instantaneous Spray Measurements of Additively Manufactured Candidate Injector Schemes at Varying Pressure Drops

Additively manufactured (AM) injectors have the potential to improve upon the combustion performance achieved using traditionally manufactured (TM) injectors. Traditionally manufactured injector combustion performance decreases greatly in off-design conditions. Additively manufactured injectors hold the potential to sustain high combustion performance over a broader range of operating conditions than currently seen in TM injectors. Additively manufactured injectors promise to offer higher resistances to instability and achieve better overall propellant mixing and atomization for all types of combustion devices. This study expands upon previous studies that used 2-D imaging techniques to characterize the injector spray pattern of candidate AM injector elements. In this study, 3-D instantaneous measurements of the injector spray pattern are taken, allowing for deeper insight into the liquid spray breakup, an inherently 3-D phenomena. Stereo imaging is used for three dimensional analysis, which utilizes two cameras at different angles to reconstruct a three dimensional reconstruction. With this information, AM injectors can be designed to improve propellant mixing and atomization and reduce breakup length. As part of this work, a high-pressure, cold flow experimental spray facility was modified to measure each injector’s spray characteristics in terms of 𝑅𝑒 𝐿 and 𝑊𝑒 𝑔 . This facility is operated to capture instantaneous stereo photographic images, mass flow rates, and injector element thrust for pressure drop across the injector face ranging from 100 to 200 psig at a stiffness ratio of 1.0.

Murphy Mitchell↗
NASA NTRS2011

Liquid Methane Conditioning Capabilities Developed at the NASA Glenn Research Center's Small Multi- Purpose Research Facility (SMiRF) for Accelerated Lunar Surface Storage Thermal Testing

Glenn Research Center s Creek Road Cryogenic Complex, Small Multi-Purpose Research Facility (SMiRF) recently completed validation / checkout testing of a new liquid methane delivery system and liquid methane (LCH4) conditioning system. Facility checkout validation was conducted in preparation for a series of passive thermal control technology tests planned at SMiRF in FY10 using a flight-like propellant tank at simulated thermal environments from 140 to 350K. These tests will validate models and provide high quality data to support consideration of LCH4/LO2 propellant combination option for a lunar or planetary ascent stage.An infrastructure has been put in place which will support testing of large amounts of liquid methane at SMiRF. Extensive modifications were made to the test facility s existing liquid hydrogen system for compatibility with liquid methane. Also, a new liquid methane fluid conditioning system will enable liquid methane to be quickly densified (sub-cooled below normal boiling point) and to be quickly reheated to saturation conditions between 92 and 140 K. Fluid temperatures can be quickly adjusted to compress the overall test duration. A detailed trade study was conducted to determine an appropriate technique to liquid conditioning with regard to the SMiRF facility s existing infrastructure. In addition, a completely new roadable dewar has been procured for transportation and temporary storage of liquid methane. A new spherical, flight-representative tank has also been fabricated for integration into the vacuum chamber at SMiRF. The addition of this system to SMiRF marks the first time a large-scale liquid methane propellant test capability has been realized at Glenn.This work supports the Cryogenic Fluid Management Project being conducted under the auspices of the Exploration Technology Development Program, providing focused cryogenic fluid management technology efforts to support NASA s future robotic or human exploration missions.

Bamberger, Helmut H.↗
NASA NTRS2007

Liquid Bismuth Propellant Flow Sensor

Quantifying the propellant mass flow rate in liquid bismuth-fed electric propulsion systems has two challenging facets. First, the flow sensors must be capable of providing a resolvable measurement at propellant mass flow rates on the order of 10 mg/see with and uncertainty of less that 5%. The second challenge has to do with the fact that the materials from which the flow sensors are fabricated must be capable of resisting any of the corrosive effects associated with the high-temperature propellant. The measurement itself is necessary in order to properly assess the performance (thrust efficiency, Isp) of thruster systems in the laboratory environment. The hotspot sensor[I] has been designed to provide the bismuth propellant mass flow rate measurement. In the hotspot sensor, a pulse of thermal energy (derived from a current pulse and associated joule heating) is applied near the inlet of the sensor. The flow is "tagged" with a thermal feature that is convected downstream by the flowing liquid metal. Downstream, a temperature measurement is performed to detect a "ripple" in the local temperature associated with the passing "hotspot" in the propellant. By measuring the time between the upstream generation and downstream detection of the thermal feature, the flow speed can be calculated using a "time of flight" analysis. In addition, the system can be calibrated by measuring the accumulated mass exiting the system as a-function of time and correlating this with the time it takes the hotspot to convect through the sensor. The primary advantage of this technique is that it doesn't depend on an absolute measurement of temperature but, instead, relies on the observation of thermal features. This makes the technique insensitive to other externally generated thermal fluctuations. In this paper, we describe experiments performed using the hotspot flow sensor aimed at quantifying the resolution of the sensor technology. Propellant is expelled onto an electronic scale to provide an independent measure of the propellant mass flow rate as a function of time. In addition, two separate detection schemes are employed. The first uses a thermocouple to directly measure temperature in the fluid. The second involves the ,use of a fiber optic coupled to a photodiode allowing for detection of an increase in light emission from the fluid as the hotspot passes. the detection location.

Polzin, Kurt A.↗
NASA NTRS1999

A Flight Demonstration of Plasma Rocket Propulsion

The Advanced Space Propulsion Laboratory at the Johnson Space Center has been engaged in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One advantageous feature of this rocket concept is the ability to vary its specific impulse so that it can be operated in a mode which maximizes propellant efficiency or a mode which maximizes thrust. This presentation will describe a proposed flight experiment in which a simple version of the rocket will be tested in space. In addition to the plasma rocket, the flight experiment will also demonstrate the use of a superconducting electromagnet, extensive use of heat pipes, and possibly the transfer of cryogenic propellant in space.

Petro, Andrew↗
NASA NTRSDate not supplied

Extreme Environment Hot Fire Durability of Post-Processed Additively Manufactured GRCop-Alloy Combustion Chambers in LOX/Hydrogen and LOX/Methane

Extreme environment survivability of metal additive manufactured (AM) GRCop-alloy thrust chambers has been demonstrated in different bi-propellants at near stoichiometric and even oxygen rich combustion. GRCop-alloy chambers tested at NASA Marshall Space Flight Center (MSFC) have accumulated over 26,000 seconds of hot fire duration and over 500 starts. These chambers are produced using an AM process called laser powder bed fusion (L-PBF). A major feature of this process is high wall surface roughness which can be customized in post processing to leverage various performance advantages. Post-processing can include heat treatment, final machining, polishing, and welding and is key to hardware survivability in extreme environments. Surface finish enhancement techniques were applied to the hot wall and coolant channels to reduce the overall total heat load to the chamber walls. Performance optimization of various thrust class TCA’s is a strategic technology goal of NASA MSFC. Three different chamber geometries using cryogenic methane and de-ionized water as coolants were hot fire tested to obtain their life cycle, pressure drop, and heat load performances. Several 1.2K lbf LOX/H2 chambers, 1K lbf LOX/CH4 chambers, and 7K lbf LOX/CH4 chambers were tested. All post-processed configurations performed remarkably well when subjected to extreme hot fire test conditions. Streaking and blanching due to localized oxygen rich conditions were observed on some test articles in their as-built surface finish state. However, this is actually a function of the injector mixing and only serves to further establish the durability of L-PBF GRCop-alloy thrust chambers. Several different polishing techniques were applied to the hot wall and integrated coolant channels prior to hot fire testing and their performances assessed. Overall, AM produced GRCop chambers are extremely reliable, durable, and customizable to the desired performance metrics.

grcop↗
NASA NTRS2010

Cryogenic Propellant Management Device: Conceptual Design Study

Concepts of Propellant Management Devices (PMDs) were designed for lunar descent stage reaction control system (RCS) and lunar ascent stage (main and RCS propulsion) missions using liquid oxygen (LO2) and liquid methane (LCH4). Study ground rules set a maximum of 19 days from launch to lunar touchdown, and an additional 210 days on the lunar surface before liftoff. Two PMDs were conceptually designed for each of the descent stage RCS propellant tanks, and two designs for each of the ascent stage main propellant tanks. One of the two PMD types is a traditional partial four-screen channel device. The other type is a novel, expanding volume device which uses a stretched, flexing screen. It was found that several unique design features simplified the PMD designs. These features are (1) high propellant tank operating pressures, (2) aluminum tanks for propellant storage, and (3) stringent insulation requirements. Consequently, it was possible to treat LO2 and LCH4 as if they were equivalent to Earth-storable propellants because they would remain substantially subcooled during the lunar mission. In fact, prelaunch procedures are simplified with cryogens, because any trapped vapor will condense once the propellant tanks are pressurized in space.

Wollen, Mark↗
NASA NTRS1993

Design, analysis, and fabrication of oxide-coated iridium/rhenium combustion chambers

Iridium-coated rhenium (Ir/Re) combustion chambers provide high temperature, oxidation-resistant operation for radiation-cooled liquid-fueled rocket engines. A 22-N (5-lb(sub f)) chamber has been operated for 15 hours at 2200 C (4000 F) using nitrogen tetroxide/monomethyl hydrazine (NTO/MMH) propellant, with negligible internal erosion. The oxidation resistance of these chambers could be further increased by the addition of refractory oxide coatings, providing longer life and/or operation in more oxidizing and higher temperature environments. The oxide coatings would serve as a thermal and diffusion barrier for the iridium coating, lowering the temperature of the iridium layer while also preventing the ingress of oxygen and egress of iridium oxides. This would serve to slow the failure mechanisms of Ir/Re chambers, namely the diffusion of rhenium to the inner surface and the oxidation of iridium. Such protection could extend chamber lifetimes by tens or perhaps hundreds of hours, and allow chamber operation on stoichiometric or higher mixture ratio oxygen/hydrogen (O2/H2) propellant. Extensive thermomechanical, thermochemical, and mass transport modeling was performed as a key material/structure design tool. Based on the results of these analyses, several 22-N oxide-coated Ir/Re chambers were fabricated and delivered to NASA Lewis Research Center for hot-fire testing.

Jang, Q.↗
NASA NTRS1993

Generic propellant transfer unit: A project update

Currently, most satellite programs use dedicated hypergolic propellant ground servicing equipment. These 'carts' are typically program specific (non-interchangeable), costly, and non-Department of Transportation (DOT) compliant. They require safety waivers when used on government property and must be emptied before off-site transport. The emptying process usually generates large amounts of Hazardous Wastes. Fuel carts are usually loaded from multiple 208-liter (55 gallon) drums and oxidizer carts from 909-kg (2000 lb) cylinders. This further increases operating costs due to additional container samples and cart-filling transfers. NASA, Kennedy Space Center has developed reusable, generic, DOT compliant propellant delivery containers as possible replacements for the propellant-tank portion of servicing carts. Our ultimate goal is to reduce operating expenses by minimizing propellant transfers, samples, and waste generation through use of bulk propellant operations, minimizing drum or cylinder operations. This has the added benefit of increasing operating safety by decreasing overall propellant operations.

Davis, C. L.↗
NASA NTRS1982

Effect of storable propellants on single-stage earth-to-orbit vehicles

A comparison is made between the storable-propellant engine and an equivalent dual-expander engine using oxygen, propane, and hydrogen propellants. The propane and hydrogen dual-expander engine is then compared with previous results with separate propane and hydrocarbon engines. It is shown that the dual-expander reduces vehicle dry mass from 94 to 89 Mg, or 5%. The hydrogen-only thrust of the dual-expander is 0.4 MN, while the separate hydrogen engines have a thrust of 2.3 MN. The low thrust level increases the difficulty of cooling the dual-expander engine, and additional hydrogen flow is added for transpirational cooling. This additional hydrogen flow leads to an increase in the hydrogen tank size. It is also shown that storable propellants increase the dry mass of single-stage earth-to-orbit vehicles by about 20% and the gross mass by about 54%. It is concluded that dual-expander engines should be studied with several thrust splits and thrust levels.

Martin, J. A.↗
NASA NTRS2008

Igniters for Liquid Oxygen/Liquid Methane Technology Development

As part of NASA's technology development of liquid methane / liquid oxygen engines for future exploration missions, two different igniters were recently studied at NASA Marshall Space Flight Center. The first igniter tested was an impinging injection, spark-initiated torch igniter, and the second was a microwave-generated plasma igniter. The purpose of the ignition tests was to define the ignition limits under vacuum conditions and characterize the transient start-up performance as a function of propellant mixture ratio (MR), mass flow rates, inlet temperatures, and pre-ignition chamber pressure. In addition, for the impinging igniter two different spark plugs were tested, and for the microwave igniter the magnetron filament warm-up time and the magnetron input power were both varied. The results gathered from these tests indicated that the impinging igniter is capable of operating over an MR range of 2 - 27, with methane and oxygen inlet temperatures as low as -161 F and -233 F, respectively. The microwave igniter was tested over an MR range of 2 - 9, with methane and oxygen inlet temperatures as low as -90 F and -200 F, respectively. The microwave igniter achieved ignition over this range, although an upper ignition limit was determined for the oxidizer mass flow rate. In general, the torch exhaust temperatures for the microwave igniter were not as high as those attained with the impinging igniter. The microwave igniter, however, was hot-fired 17 times and was still operational, whereas the impinging igniter spark plugs experienced thermal shock and erosion over nine hot-fire tests. It was concluded that for the microwave igniter better mixing of the propellants might be required in order to both raise the torch exhaust temperature and decrease the required magnetron input power, and for the impinging igniter the spark plug position within the igniter chamber should be varied in future tests to identify a more optimal location. All of the igniter tests were supported by the Propulsion & Cryogenics Advanced Development project, which is part of NASA's Exploration Technology Development Program.

Osborne, Robin J.↗
NASA NTRS1974

Development of high temperature materials for solid propellant rocket nozzle applications

Evaluation of tantalum carbide-tungsten fiber composites has been completed as far as weight percent carbon additions and weight percent additions of tungsten fiber. Extensive studies were undertaken concerning Young's Modulus and fracture strength of this material. Also, in-depth analysis of the embrittling effects of the extra carbon additions on the tungsten fibers has been completed. The complete fabrication procedure for the tantalum carbide-tungsten fiber composites with extra carbon additions is given. Microprobe and metallographic studies showed the effect of extra carbon on the tungsten fibers, and evaluation of the thermal shock parameter fracture strength/Young's Modulus is included.

Manning, C. R., Jr.↗
NASA NTRS2018

Additive Manufacturing of Liquid Rocket Engine Combustion Devices: A Summary of Process Developments and Hot-Fire Testing Results

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.

Gradl, Paul↗
NASA NTRSDate not supplied

Slosh-Induced Ullage Collapse Validation by Computational Fluid Dynamics and Application to In-space Propellant Dynamics

Ullage collapse predictions made using Computational Fluid Dynamics (CFD) were validated then made for an in-space cryogenic propellant unsettling application. Ullage collapse is the decrease in propellant tank pressure due to cooling of the ullage gas, which can be significant for cryogenic propulsion systems. Vehicles may be developed that are robust to temperature and pressure changes caused by heat leaks, e.g., propellant tanks with high maximum operation pressures, technologies yielding active thermal conditioning, and systems with ample pressurant for conditioning. These methods are robust because additional mass, cost, and complexity are used to address cryogenic fluid management (CFM). As vehicles are optimized for increased payload capacity, tracking of events such as ullage collapse and self-pressurization in tanks will become increasingly important for propellant inventory calculations. Some ullage collapse events can be studied through ground experimentation, but propellant unsettling in and around in-space coast mission phases must be studied in the absence of an acceleration field. Computational Fluid Dynamics is a powerful tool to address this need, enabling prediction of complex physics prior to flight. Validation of the chosen CFD tool and modeling methodology was done using data from a previously conducted subscale slosh-induced ullage collapse ground test performed in part by the Japan Aerospace Exploration Agency (JAXA). A small tank was partially filled with liquid nitrogen and pressurized with gaseous nitrogen at various temperatures. The tank was oscillated laterally to induce high amplitude slosh dynamics resulting in wave breakup. Successful qualitative validation of slosh wave dynamics and quantitative validation of ullage collapse promoted confidence in extending ullage collapse predictions to a flight application. Mechanisms of ullage collapse caused by main engine cutoff while in-space were identified. Both slosh dynamics leading to engine shutdown and body dynamics following shutdown were shown to contribute to propellant unsettling and ullage collapse.

J M Brodnick↗
NASA NTRS2015

Liquid Hydrogen Propellant Tank Sub-Surface Pressurization with Gaseous Helium

A series of tests were conducted to evaluate the performance of a propellant tank pressurization system with the pressurant diffuser intentionally submerged beneath the surface of the liquid. Propellant tanks and pressurization systems are typically designed with the diffuser positioned to apply pressurant gas directly into the tank ullage space when the liquid propellant is settled. Space vehicles, and potentially propellant depots, may need to conduct tank pressurization operations in micro‐gravity environments where the exact location of the liquid relative to the diffuser is not well understood. If the diffuser is positioned to supply pressurant gas directly to the tank ullage space when the propellant is settled, then it may become partially or completely submerged when the liquid becomes unsettled in a microgravity environment. In such case, the pressurization system performance will be adversely affected requiring additional pressurant mass and longer pressurization times. This series of tests compares and evaluates pressurization system performance using the conventional method of supplying pressurant gas directly to the propellant tank ullage, and then supplying pressurant gas beneath the liquid surface. The pressurization tests were conducted on the Engineering Development Unit (EDU) located at Test Stand 300 at NASA Marshall Space Flight Center (MSFC). EDU is a ground based Cryogenic Fluid Management (CFM) test article supported by Glenn Research Center (GRC) and MSFC. A 150 ft3 propellant tank was filled with liquid hydrogen (LH2). The pressurization system used regulated ambient helium (GHe) as a pressurant, a variable position valve to maintain flow rate, and two identical independent pressurant diffusers. The ullage diffuser was located in the forward end of the tank and was completely exposed to the tank ullage. The submerged diffuser was located in the aft end of the tank and was completely submerged when the tank liquid level was 10% or greater. The ullage diffuser tests were conducted as a baseline to evaluate the performance of the pressurization system, and the submerged diffuser tests showed how the performance of the pressurization system was compromised when the diffuser was submerged in LH2. The test results are evaluated and compared, and included in this report for various propellant tank fill levels.

Stephens, J. R.↗
NASA NTRS1973

Theoretical analysis of rotating two phase detonation in a rocket motor

Tangential mode, non-linear wave motion in a liquid propellant rocket engine is studied, using a two phase detonation wave as the reaction model. Because the detonation wave is followed immediately by expansion waves, due to the side relief in the axial direction, it is a Chapman-Jouguet wave. The strength of this wave, which may be characterized by the pressure ratio across the wave, as well as the wave speed and the local wave Mach number, are related to design parameters such as the contraction ratio, chamber speed of sound, chamber diameter, propellant injection density and velocity, and the specific heat ratio of the burned gases. In addition, the distribution of flow properties along the injector face can be computed. Numerical calculations show favorable comparison with experimental findings. Finally, the effects of drop size are discussed and a simple criterion is found to set the lower limit of validity of this strong wave analysis.

Shen, I.↗
NASA NTRS1976

Evaluation of ice and frost accumulation on the Space Shuttle external tank

Ice/frost formation on the Space Shuttle cryogenic propellant tanks presents a different problem from that of past launch vehicles. Lift off weight addition has been the primary concern on past launch vehicles. The primary ice/frost concern on the Shuttle vehicle is damage to the Orbiter Thermal Protection System due to ice/frost impact. The approach used to arrive at a solution to this unique Shuttle problem is presented. The launch vehicle configuration selected and its limitations are described, along with contingency ground support equipment.

Rhodes, R. E.↗
NASA NTRS1978

Rocket effluent - Its ice nucleation activity and related properties

To investigate the possibility of inadvertent weather modification from rocket effluent, aerosol samples were collected from an instrumented aircraft subsequent to the Voyager I and II launches. The aerosol's morphology, concentration and size distribution were examined with an electron microscope. The elemental compositions of individual particles were analyzed with an X-ray energy spectrometer. Ice nucleus concentration was measured with a subfreezing thermal diffusion chamber. The particles' physical and chemical properties were related to their ice nucleation activity. A laboratory experiment on rocket propellant exhaust was conducted under controlled conditions. Both laboratory and field experimental results indicated that rocket propellant exhaust can produce active ice nuclei. Their consequences for potential inadvertant weather modification demand additional study.

Parungo, F. P.↗
NASA NTRS1984

Investigation of fuselage acoustic treatment for a twin-engine turboprop aircraft in flight and laboratory tests

A flight and laboratory study of sidewall acoustic treatment for cabin noise control is described. In flight, cabin noise levels were measured at six locations with three treatment configurations. Noise levels from narrow-band analysis are reduced to one-third octave format and used to calculate insertion loss, IL, defined as the reduction of interior noise associated with the addition of a treatment. Laboratory tests used a specially constructed structural panel modeled after the propeller plane section of the aircraft sidewall, and acoustic treatments representing those used in flight. Lab measured transmission loss and absorption values were combined using classical acoustic procedures to obtain a prediction of IL. Comparison with IL values measured in flight for the boundary layer component of the noise indicated general agreement.

Mixson, J. S.↗