Engineering topics

Tomsik, Thomas M.

Publications and source records attributed to Tomsik, Thomas M..

At least 19 records

NASA NTRS2018

Reconfiguration of NASA GRC's Vacuum Facility 6 for Testing of Advanced Electric Propulsion System (AEPS) Hardware

The NASA Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Hall thruster has been the subject of extensive technology maturation in preparation for development into a flight propulsion system. The HERMeS thruster is being developed and tested at NASA GRC and NASA JPL through support of the Space Technology Mission Directorate (STMD) and is intended to be used as the electric propulsion system on the Power and Propulsion Element (PPE) of the recently announced Deep Space Gateway (DSG). The Advanced Electric Propulsion System (AEPS) contract was awarded to Aerojet-Rocketdyne to develop the HERMeS system into a flight system for use by NASA. To address the hardware test needs of the AEPS project, NASA GRC launched an effort to reconfigure Vacuum Facility 6 (VF-6) for high-power electric propulsion testing including upgrades and reconfigurations necessary to conduct performance, plasma plume, and system level integration testing. Results of the verification and validation testing with HERMeS Technology Demonstration Unit (TDU)-1 and TDU-3 Hall thrusters are also included.

Peterson, Peter Y.↗

NASA NTRS2017

Large Scale Production of Densified Hydrogen Using Integrated Refrigeration and Storage

Recent demonstration of advanced liquid hydrogen storage techniques using Integrated Refrigeration and Storage (IRAS) technology at NASA Kennedy Space Center led to the production of large quantities of solid densified liquid and slush hydrogen in a 125,000 L tank. Production of densified hydrogen was performed at three different liquid levels and LH2 temperatures were measured by twenty silicon diode temperature sensors. System energy balances and solid mass fractions are calculated. Experimental data reveal hydrogen temperatures dropped well below the triple point during testing (up to 1 K), and were continuing to trend downward prior to system shutdown. Sub-triple point temperatures were seen to evolve in a time dependent manner along the length of the horizontal, cylindrical vessel. Twenty silicon diode temperature sensors were recorded over approximately one month for testing at two different fill levels (33 67). The phenomenon, observed at both two fill levels, is described and presented detailed and explained herein., and The implications of using IRAS for energy storage, propellant densification, and future cryofuel systems are discussed.

Notardonato, William U.↗

NASA NTRS2017

Modeling Xenon Tank Pressurization using One-Dimensional Thermodynamic and Heat Transfer Equations

As a first step in understanding what ground support equipment (GSE) is required to provide external cooling during the loading of 5,000 kg of xenon into 4 aluminum lined composite overwrapped pressure vessels (COPVs), a modeling analysis was performed using Microsoft Excel. The goals of the analysis were to predict xenon temperature and pressure throughout loading at the launch facility, estimate the time required to load one tank, and to get an early estimate of what provisions for cooling xenon might be needed while the tanks are being filled. The model uses the governing thermodynamic and heat transfer equations to achieve these goals. Results indicate that a single tank can be loaded in about 15 hours with reasonable external coolant requirements. The model developed in this study was successfully validated against flight and test data. The first data set is from the Dawn mission which also utilizes solar electric propulsion with xenon propellant, and the second is test data from the rapid loading of a hydrogen cylindrical COPV. The main benefit of this type of model is that the governing physical equations using bulk fluid solid temperatures can provide a quick and accurate estimate of the state of the propellant throughout loading which is much cheaper in terms of computational time and licensing costs than a Computation Fluid Dynamics (CFD) analysis while capturing the majority of the thermodynamics and heat transfer.

Thermodynamic tank modeling↗

NASA NTRS2011

An Active Broad Area Cooling Model of a Cryogenic Propellant Tank with a Single Stage Reverse Turbo-Brayton Cycle Cryocooler

As focus shifts towards long-duration space exploration missions, an increased interest in active thermal control of cryogenic propellants to achieve zero boil-off of cryogens has emerged. An active thermal control concept of considerable merit is the integration of a broad area cooling system for a cryogenic propellant tank with a combined cryocooler and circulator system that can be used to reduce or even eliminate liquid cryogen boil-off. One prospective cryocooler and circulator combination is the reverse turbo-Brayton cycle cryocooler. This system is unique in that it has the ability to both cool and circulate the coolant gas efficiently in the same loop as the broad area cooling lines, allowing for a single cooling gas loop, with the primary heat rejection occurring by way of a radiator and/or aftercooler. Currently few modeling tools exist that can size and characterize an integrated reverse turbo-Brayton cycle cryocooler in combination with a broad area cooling design. This paper addresses efforts to create such a tool to assist in gaining a broader understanding of these systems, and investigate their performance in potential space missions. The model uses conventional engineering and thermodynamic relationships to predict the preliminary design parameters, including input power requirements, pressure drops, flow rate, cycle performance, cooling lift, broad area cooler line sizing, and component operating temperatures and pressures given the cooling load operating temperature, heat rejection temperature, compressor inlet pressure, compressor rotational speed, and cryogenic tank geometry. In addition, the model allows for the preliminary design analysis of the broad area cooling tubing, to determine the effect of tube sizing on the reverse turbo-Brayton cycle system performance. At the time this paper was written, the model was verified to match existing theoretical documentation within a reasonable margin. While further experimental data is needed for full validation, this tool has already made significant steps towards giving a clearer understanding of the performance of a reverse turbo-Brayton cycle cryocooler integrated with broad area cooling technology for zero boil-off active thermal control.

Guzik, Monica C.↗

NASA NTRS2010

A Densified Liquid Methane Delivery System for the Altair Ascent Stage

The Altair Lunar Lander is currently carrying options for both cryogenic and hypergolic ascent stage propulsion modules. The cryogenic option uses liquid methane and liquid oxygen to propel Altair from the lunar surface back to rendezvous with the Orion command module. Recent studies have determined that the liquid methane should be densified by subcooling it to 93 K in order to prevent over-pressurization of the propellant tanks during the 210 day stay on the lunar surface. A trade study has been conducted to determine the preferred method of producing; loading, and maintaining the subcooled, densified liquid methane onboard Altair from a ground operations perspective. The trade study took into account the limitations in mass for the launch vehicle and the mobile launch platform as well as the historical reliability of various components and their thermal efficiencies. Several unique problems were encountered, namely delivering a small amount of a cryogenic propellant to a flight tank that is positioned over 350 ft above the launch pad as well as generating the desired delivery temperature of the methane at 93 K which is only 2.3 K above the methane triple point of 90.7 K. Over 20 methods of subcooled liquid methane production and delivery along with the associated system architectures were investigated to determine the best solutions to the problem. The top four cryogenic processing solutions were selected for further evaluation and detailed thermal modeling. This paper describes the results of the preliminary trade analysis of the 20 plus methane densification methods considered. The results of the detailed analysis will be briefed to the Altair Project Office and their propulsion team as well as the Ground Operations Project Office before the down-select is made between cryogenic and hypergolic ascent stages in August 2010.

Tomsik, Thomas M.↗

NASA NTRS2010

Liquid Oxygen Propellant Densification Production and Performance Test Results With a Large-Scale Flight-Weight Propellant Tank for the X33 RLV

This paper describes in-detail a test program that was initiated at the Glenn Research Center (GRC) involving the cryogenic densification of liquid oxygen (LO2). A large scale LO2 propellant densification system rated for 200 gpm and sized for the X-33 LO2 propellant tank, was designed, fabricated and tested at the GRC. Multiple objectives of the test program included validation of LO2 production unit hardware and characterization of densifier performance at design and transient conditions. First, performance data is presented for an initial series of LO2 densifier screening and check-out tests using densified liquid nitrogen. The second series of tests show performance data collected during LO2 densifier test operations with liquid oxygen as the densified product fluid. An overview of LO2 X-33 tanking operations and load tests with the 20,000 gallon Structural Test Article (STA) are described. Tank loading testing and the thermal stratification that occurs inside of a flight-weight launch vehicle propellant tank were investigated. These operations involved a closed-loop recirculation process of LO2 flow through the densifier and then back into the STA. Finally, in excess of 200,000 gallons of densified LO2 at 120 oR was produced with the propellant densification unit during the demonstration program, an achievement that s never been done before in the realm of large-scale cryogenic tests.

Tomsik, Thomas M.↗

NASA NTRS2009

Synthesis, Decomposition and Characterization of Fe and Ni Sulfides and Fe and CO Nanoparticles for Aerospace Applications

We describe several related studies where simple iron, nickel, and cobalt complexes were prepared, decomposed, and characterized for aeronautics (Fischer-Tropsch catalysts) and space (high-fidelity lunar regolith simulant additives) applications. We describe the synthesis and decomposition of several new nickel dithiocarbamate complexes. Decomposition resulted in a somewhat complicated product mix with NiS predominating. The thermogravimetric analysis of fifteen tris(diorganodithiocarbamato)iron(III) has been investigated. Each undergoes substantial mass loss upon pyrolysis in a nitrogen atmosphere between 195 and 370 C, with major mass losses occurring between 279 and 324 C. Steric repulsion between organic substituents generally decreased the decomposition temperature. The product of the pyrolysis was not well defined, but usually consistent with being either FeS or Fe2S3 or a combination of these. Iron nanoparticles were grown in a silica matrix with a long-term goal of introducing native iron into a commercial lunar dust simulant in order to more closely simulate actual lunar regolith. This was also one goal of the iron and nickel sulfide studies. Finally, cobalt nanoparticle synthesis is being studied in order to develop alternatives to crude processing of cobalt salts with ceramic supports for Fischer-Tropsch synthesis.

Cowen, Jonathan E.↗

NASA NTRS2009

Hydrogen Fuel System Design Trades for High-Altitude Long-Endurance Remotely- Operated Aircraft

Preliminary design trades are presented for liquid hydrogen fuel systems for remotely-operated, high-altitude aircraft that accommodate three different propulsion options: internal combustion engines, and electric motors powered by either polymer electrolyte membrane fuel cells or solid oxide fuel cells. Mission goal is sustained cruise at 60,000 ft altitude, with duration-aloft a key parameter. The subject aircraft specifies an engine power of 143 to 148 hp, gross liftoff weight of 9270 to 9450 lb, payload of 440 lb, and a hydrogen fuel capacity of 2650 to 2755 lb stored in two spherical tanks (8.5 ft inside diameter), each with a dry mass goal of 316 lb. Hydrogen schematics for all three propulsion options are provided. Each employs vacuum-jacketed tanks with multilayer insulation, augmented with a helium pressurant system, and using electric motor driven hydrogen pumps. The most significant schematic differences involve the heat exchangers and hydrogen reclamation equipment. Heat balances indicate that mission durations of 10 to 16 days appear achievable. The dry mass for the hydrogen system is estimated to be 1900 lb, including 645 lb for each tank. This tank mass is roughly twice that of the advanced tanks assumed in the initial conceptual vehicle. Control strategies are not addressed, nor are procedures for filling and draining the tanks.

Millis, Marc G.↗

NASA NTRS2006

Bench-Scale Monolith Autothermal Reformer Catalyst Screening Evaluations in a Micro-Reactor With Jet-A Fuel

Solid oxide fuel cell systems used in the aerospace or commercial aviation environment require a compact, light-weight and highly durable catalytic fuel processor. The fuel processing method considered here is an autothermal reforming (ATR) step. The ATR converts Jet-A fuel by a reaction with steam and air forming hydrogen (H2) and carbon monoxide (CO) to be used for production of electrical power in the fuel cell. This paper addresses the first phase of an experimental catalyst screening study, looking at the relative effectiveness of several monolith catalyst types when operating with untreated Jet-A fuel. Six monolith catalyst materials were selected for preliminary evaluation and experimental bench-scale screening in a small 0.05 kWe micro-reactor test apparatus. These tests were conducted to assess relative catalyst performance under atmospheric pressure ATR conditions and processing Jet-A fuel at a steam-to-carbon ratio of 3.5, a value higher than anticipated to be run in an optimized system. The average reformer efficiencies for the six catalysts tested ranged from 75 to 83 percent at a constant gas-hourly space velocity of 12,000 hr 1. The corresponding hydrocarbon conversion efficiency varied from 86 to 95 percent during experiments run at reaction temperatures between 750 to 830 C. Based on the results of the short-duration 100 hr tests reported herein, two of the highest performing catalysts were selected for further evaluation in a follow-on 1000 hr life durability study in Phase II.

Tomsik, Thomas M.↗

NASA NTRS2004

Auto-Thermal Reforming of Jet-A Fuel over Commercial Monolith Catalysts: MicroReactor Evaluation and Screening Test Results

This paper describes the results of a series of catalyst screening tests conducted with Jet-A fuel under auto-thermal reforming (ATR) process conditions at the research laboratories of SOFCo-EFS Holdings LLC under Glenn Research Center Contract. The primary objective is to identify best available catalysts for future testing at the NASA GRC 10-kW(sub e) reformer test facility. The new GRC reformer-injector test rig construction is due to complete by March 2004. Six commercially available monolithic catalyst materials were initially selected by the NASA/SOFCo team for evaluation and bench scale screening in an existing 0.05 kW(sub e) microreactor test apparatus. The catalyst screening tests performed lasted 70 to 100 hours in duration in order to allow comparison between the different samples over a defined range of ATR process conditions. Aging tests were subsequently performed with the top two ranked catalysts as a more representative evaluation of performance in a commercial aerospace application. The two catalyst aging tests conducted lasting for approximately 600 hours and 1000 hours, respectively.

Yen, Judy C. H.↗

NASA NTRS2002

Liquid Oxygen Propellant Densification Unit Ground Tested With a Large-Scale Flight-Weight Tank for the X-33 Reusable Launch Vehicle

Propellant densification has been identified as a critical technology in the development of single-stage-to-orbit reusable launch vehicles. Technology to create supercooled high-density liquid oxygen (LO2) and liquid hydrogen (LH2) is a key means to lowering launch vehicle costs. The densification of cryogenic propellants through subcooling allows 8 to 10 percent more propellant mass to be stored in a given unit volume, thereby improving the launch vehicle's overall performance. This allows for higher propellant mass fractions than would be possible with conventional normal boiling point cryogenic propellants, considering the normal boiling point of LO2 and LH2.

Tomsik, Thomas M.↗

NASA NTRS2001

Testing of Densified Liquid Hydrogen Stratification in a Scale Model Propellant Tank

This paper describes a test program that was conducted at NASA to demonstrate the ability to load densified LH2 into a subscale propellant tank. This work was done through a collaborative effort between NASA Glenn Research Center and the Lockheed Martin Michoud Space Systems (LMMSS). The Multilobe tank, which was made from composite materials similar to that to be used on X-33, was formed from two lobes with a center septum. Test results are shown for data that was collected on filling the subscale tank with densified liquid hydrogen (DLH2) propellant that was produced at the NASA Plum Brook Station. Data is compared to analytical predictions. Data collected for this test series agrees well with analytical predictions of the environmental heat leak into the tank and the thermal stratification characteristics of the hydrogen propellant in the tank as it was filled with DLH2.

Jurns, John M.↗

NASA NTRS2000

Recent Advances and Applications in Cryogenic Propellant Densification Technology

This purpose of this paper is to review several historical cryogenic test programs that were conducted at the NASA Glenn Research Center (GRC), Cleveland, Ohio over the past fifty years. More recently these technology programs were intended to study new and improved denser forms of liquid hydrogen (LH2) and liquid oxygen (LO2) cryogenic rocket fuels. Of particular interest are subcooled cryogenic propellants. This is due to the fact that they have a significantly higher density (eg. triple-point hydrogen, slush etc.), a lower vapor pressure and improved cooling capacity over the normal boiling point cryogen. This paper, which is intended to be a historical technology overview, will trace the past and recent development and testing of small and large-scale propellant densification production systems. Densifier units in the current GRC fuels program, were designed and are capable of processing subcooled LH2 and L02 propellant at the X33 Reusable Launch Vehicle (RLV) scale. One final objective of this technical briefing is to discuss some of the potential benefits and application which propellant densification technology may offer the industrial cryogenics production and end-user community. Density enhancements to cryogenic propellants (LH2, LO2, CH4) in rocket propulsion and aerospace application have provided the opportunity to either increase performance of existing launch vehicles or to reduce the overall size, mass and cost of a new vehicle system.

Tomsik, Thomas M.↗

NASA NTRS2000

Propellant Densification Ground Testing Conducted for Launch Vehicles

The NASA Glenn Research Center at Lewis Field has taken the lead in the development of practical densified cryogenic propellants for launch vehicle applications. The technology of subcooling cryogenic propellants below their normal boiling point to produce a denser fluid is one of the key process technologies necessary to meet the challenge of single-stage-to-orbit and reusable launch vehicles. Densified propellants are critical to lowering launch costs because they enable more propellant to be packed into a given unit volume, thus improving the performance by reducing the overall size and weight of the launch vehicle. This two-pronged research and test program has evolved into (1) conducting tank loading tests using densified liquid hydrogen and (2) developing two large-scale propellant densification systems that will be performance tested next year at Glenn. The propellant-loading test program was undertaken at Glenn in coordination with Lockheed Martin Michoud Space Systems. In this testing, the liquid hydrogen recirculation and densification process was simulated, and the thermal stratification of the densified propellant was recorded throughout the tank. The test article was a flight-weight tank constructed from composite materials similar to those to be used on the X-33 launch vehicle. The tank geometry as designed by Lockheed Martin had two cylindrical lobes with a center septum. Liquid hydrogen flow rate, pressure data, and temperature data plotted over time were collected while the subscale tank was filled with 27 R (15 K) densified liquid hydrogen propellant. This testing has validated mathematical models and demonstrated the readiness of densified propellant technology for near-term use. It marks the first time that such a process has been carried out with a multiple-lobe, flight-similar tank. Glenn researchers have also been working on providing a process and critical test data for the continuous production of densified liquid hydrogen (LH2) and densified liquid oxygen (LO2). Each densification production process uses a high-efficiency, subatmospheric boiling bath heat exchanger to cool the working fluid. A near triple-point hydrogen boiling bath is used to condition and subcool hydrogen to 27 R (15 K), and a nitrogen boiling bath is used to cool the liquid oxygen to 120 R (66.7 K). Multistage centrifugal compressors operating at cryogenic inlet conditions maintain the heat exchanger bath vapor pressure below 1 atm. The LO2 propellant densification unit shown in the photograph has a 30 lb/sec capacity, whereas the LH2 unit was designed to process 8 lb/sec of propellant. Each densification unit will be transported to Glenn's South Forty area after all fabrication work is completed sometime late next year. There the LO2 and LH2 densifier performance tests will be conducted with another larger Lockheed Martin tank designated the Structural Test Article (STA). This liquid oxygen tank is a full-scale, flight-weight, prototype aluminum tank designed for the X-33. It has a capacity of 20,000 gallons of LO2. The tank loading and recirculation testing planned for next year with STA will provide the data necessary for full-scale development of propellant densification technology.

Tomsik, Thomas M.↗

NASA NTRS1998

Recent Advancements in Propellant Densification

Next-generation launch vehicles demand several technological improvements to achieve lower cost and more reliable access to space. One technology area whose performance gains may far exceed others is densified propellants. The ideal rocket engine propellant is characterized by high specific impulse, high density, and low vapor pressure. A propellant combination of liquid hydrogen and liquid oxygen (LH2/LOX) is one of the highest performance propellants, but LH2 stored at standard conditions has a relatively low density and high vapor pressure. Propellant densification can significantly improve this propellant's properties relative to vehicle design and engine performance. Vehicle performance calculations based on an average of existing launch vehicles indicate that densified propellants may allow an increase in payload mass of up to 5 percent. Since the NASA Lewis Research Center became involved with the National Aerospace Plane program in the 1980's, it has been leading the way in making densified propellants a viable fuel for next-generation launch vehicles. Lewis researchers have been working to provide a method and critical data for continuous production of densified hydrogen and oxygen.

McNelis, Nancy B.↗

NASA NTRS1997

Performance Tests of a Liquid Hydrogen Propellant Densification Ground System for the X33/RLV

A concept for improving the performance of propulsion systems in expendable and single-stage-to-orbit (SSTO) launch vehicles much like the X33/RLV has been identified. The approach is to utilize densified cryogenic liquid hydrogen (LH2) and liquid oxygen (LOX) propellants to fuel the propulsion stage. The primary benefit for using this relatively high specific impulse densified propellant mixture is the subsequent reduction of the launch vehicle gross lift-off weight. Production of densified propellants however requires specialized equipment to actively subcool both the liquid oxygen and liquid hydrogen to temperatures below their normal boiling point. A propellant densification unit based on an external thermodynamic vent principle which operates at subatmospheric pressure and supercold temperatures provides a means for the LH2 and LOX densification process to occur. To demonstrate the production concept for the densification of the liquid hydrogen propellant, a system comprised of a multistage gaseous hydrogen compressor, LH2 recirculation pumps and a cryogenic LH2 heat exchanger was designed, built and tested at the NASA Lewis Research Center (LeRC). This paper presents the design configuration of the LH2 propellant densification production hardware, analytical details and results of performance testing conducted with the hydrogen densifier Ground Support Equipment (GSE).

Tomsik, Thomas M.↗

NASA NTRS1995

A Summary of the Slush Hydrogen Technology Program for the National Aero-Space Plane

Slush hydrogen, a mixture of solid and liquid hydrogen, offers advantages of higher density (16 percent) and higher heat capacity (18 percent) than normal boiling point hydrogen. The combination of increased density and heat capacity of slush hydrogen provided a potential to decrease the gross takeoff weight of the National Aero-Space Plane (NASP) and therefore slush hydrogen was selected as the propellant. However, no large-scale data was available on the production, transfer and tank pressure control characteristics required to use slush hydrogen as a fuel. Extensive testing has been performed at the NASA Lewis Research Center K-Site and Small Scale Hydrogen Test Facility between 1990 and the present to provide a database for the use of slush hydrogen. This paper summarizes the results of this testing.

Mcnelis, Nancy B.↗

NASA NTRS1994

A hydrogen-oxygen rocket engine coolant passage design program (RECOP) for fluid-cooled thrust chambers and nozzles

The design of coolant passages in regeneratively cooled thrust chambers is critical to the operation and safety of a rocket engine system. Designing a coolant passage is a complex thermal and hydraulic problem requiring an accurate understanding of the heat transfer between the combustion gas and the coolant. Every major rocket engine company has invested in the development of thrust chamber computer design and analysis tools; two examples are Rocketdyne's REGEN code and Aerojet's ELES program. In an effort to augment current design capabilities for government and industry, the NASA Lewis Research Center is developing a computer model to design coolant passages for advanced regeneratively cooled thrust chambers. The RECOP code incorporates state-of-the-art correlations, numerical techniques and design methods, certainly minimum requirements for generating optimum designs of future space chemical engines. A preliminary version of the RECOP model was recently completed and code validation work is in progress. This paper introduces major features of RECOP and compares the analysis to design points for the first test case engine; the Pratt & Whitney RL10A-3-3A thrust chamber.

Tomsik, Thomas M.↗