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At least 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Fluid property and operational effects on swirl injector atomization

Results of atomization tests of rocket injection elements are correlated in terms of the test fluid properties and the injection element operating conditions. The type of element modeled has a central, swirling liquid oxidizer flow; and a coaxial, axial gaseous fuel flow. Results of the correlations describe the effect of geometry, fluid properties, and operating conditions on liquid atomization. The ranges of test conditions allow reasonable extrapolation to actual high-pressure engine operating conditions. At such conditions, the swirl element should produce exceptionally finely-atomized and well-distributed spray for high combustion performance. United Technologies Research Center (UTRC) completed a test program to measure fluid spray atomization, in terms of droplet size and droplet size distributions, for swirling-flow injection elements. Element sizes and dimensions covered ranges of dimensions representative of typical cryogenic (O2/H2) liquid rocket engine practice. Relevant dimensions include the element diameter, the tangential-entry metering slot length, and the coastal fuel slot radial gap. Test fluids included water, simulated Jet-A, and Freon 113. These fluids cover ranges of properties which include or approach those of liquid oxygen. Gaseous nitrogen simulated the coaxial gaseous hydrogen fuel flow.

Cox, G. B., Jr.↗
NASA NTRS2017

Pressure-Fed LOX/LCH4 Reaction Control System for Spacecraft: Transient Modeling and Thermal Vacuum Hotfire Test Results

An integrated cryogenic liquid oxygen, liquid methane (LOX/LCH4) reaction control system (RCS) was tested at NASA Glenn Research Center's Plum Brook Station in the Spacecraft Propulsion Research Facility (B-2) under vacuum and thermal vacuum conditions. The RCS is a subsystem of the Integrated Cryogenic Propulsion Test Article (ICPTA), a pressure-fed LOX/LCH4 propulsion system composed of a single 2,800 lbf main engine, two 28 lbf RCS engines, and two 7 lbf RCS engines. Propellants are stored in four 48 inch diameter 5083 aluminum tanks that feed both the main engine and RCS engines in parallel. Helium stored cryogenically in a composite overwrapped pressure vessel (COPV) flows through a heat exchanger on the main engine before being used to pressurize the propellant tanks to a design operating pressure of 325 psi. The ICPTA is capable of simultaneous main engine and RCS operation. The RCS engines utilize a coil-on-plug (COP) ignition system designed for operation in a vacuum environment, eliminating corona discharge issues associated with a high voltage lead. There are two RCS pods on the ICPTA, with two engines on each pod. One of these two engines is a heritage flight engine from Project Morpheus. Its sea level nozzle was removed and replaced by an 85:1 nozzle machined using Inconel 718, resulting in a maximum thrust of 28 lbf under altitude conditions. The other engine is a scaled down version of the 28 lbf engine, designed to match the core and overall mixture ratios as well as other injector characteristics. This engine can produce a maximum thrust of 7 lbf with an 85:1 nozzle that was additively manufactured using Inconel 718. Both engines are film-cooled and capable of limited duration gas-gas and gas-liquid operation, as well as steady-state liquid-liquid operation. Each pod contains one of each version, such that two engines of the same thrust level can be fired as a couple on opposite pods. The RCS feed system is composed of symmetrical 3/8 inch lines that tap off of the main propellant manifold to send LOX and LCH4 outboard to the RCS pods. A Thermodynamic Vent System (TVS) is used to condition propellants at each pod by venting through an orifice and then routing the cold expansion products back through tubing that is welded along a large portion of the main RCS feed lines. Prior to final installation on the ICPTA, the RCS engines were tested in a small vacuum chamber at the Johnson Space Center (JSC) Energy Systems Test Area (ESTA) to verify functionality of the new COP ignition system and check out operation of the vacuum nozzles. After engine-level testing, the RCS engines were installed on the vehicle and a series of integrated hot-fire tests were performed at JSC consisting of various pulsing and steady-state firings as well as integrated main engine/RCS operation. The ICPTA was then integrated into the Plum Brook B-2 facility for vacuum and thermal/vacuum testing. Testing in the B-2 facility was composed of multiple thermal and pressure environments. The first set of tests were performed under ambient temperature and altitude pressure conditions. These tests consisted of a range of minimum impulse bit (MIB) pulsing sequences with low duty cycle, analogous to a coast phase in which the RCS is primarily used for station keeping. The primary goal of this sequence is to understand how propellant conditions were effected without an active TVS. In this scenario, consistent gas-gas operation is desirable since it results in a smaller MIB and more efficient propellant consumption. Multiple skin thermocouples are mounted on the feedlines, in addition to a submerged thermocouple on each commodity, in order to gather thermal data on the system. Higher duty cycle pulsing tests were then performed, analogous to an ascent or landing mission phase. The primary goal of this sequence was to examine how well the engines self-conditioned without active TVS when starting from a quiescent state. The TVS was then activated during some tests to demonstrate the capability to quickly condition the engines for higher pulsing demand scenarios. A thermocouple at the TVS outlet allows for the calculation of energy absorbed by the vented propellant. Lastly, tests with longer pulses and multiple engines firing either in sequence or simultaneously were run in order to gather transient system response data on waterhammer. Six total high-speed pressure transducers are installed on the RCS system, one sensor at the end of each propellant manifold line on the pods, and one at the tap-off location for each commodity. This will allow for the accurate characterization of waterhammer in the system under various propellant conditions and firing sequences. Other instrumentation for this test series includes nozzle throat thermocouples, chamber pressure measurement, heat soakback measurement, and tank wall plume impingement temperature measurement. The next set of tests were performed to demonstrate simultaneous main engine and RCS operation. Data from this test will be used to examine if there is any change to nominal operation of the RCS as a result of feed system interaction or other phenomenon. Some of these tests began under high vacuum conditions (target ambient pressure less than 1x10(exp -3) torr) and others began at altitude conditions. The last set of tests were performed with the B-2 cold wall active. Under these tests, many of the same low duty cycle MIB tests were repeated in order to characterize how propellant conditions changed with the lower heat leak. In this scenario the RCS manifold experiences much less heat leak, resulting in a change to how well the engines self-condition. As a result, an increase in maximum waterhammer pressures and a change in natural frequency of the system was expected due to higher density propellants. The lower heat leak should also result in a change to the MIB pulse profile, and data will be examined to understand how MIB repeatability is affected in the different operating environments. Parallel to the test efforts, a set of transient model development efforts were made to predict RCS performance. The primary effort was aimed at producing a SINDA/FLUINT model to predict propellant conditioning up to the engine inlet as a function of different environmental and operating parameters, with the goal of predicting chamber pressure, TVS performance, and propellant consumption over time. Preliminary results for this effort will be presented in comparison with test data. Additional modeling efforts were made using SINDA/FLUINT to predict waterhammer in the system since the software is capable of handling multiphase transient fluid dynamics. These results will be compared with the high-speed pressure transducer test data for validation purposes.

Atwell, Matthew J.↗
NASA NTRS2010

Northrop Grumman TR202 LOX/LH2 Deep Throttling Engine Project Status

NASA's Propulsion and Cryogenic Advanced Development (PCAD) project is currently developing enabling propulsion technologies in support of the Exploration Initiative, with a particular focus on the needs of the Altair Project. To meet Altair requirements, several technical challenges need to be overcome, one of which is the ability for the lunar descent engine(s) to operate over a deep throttle range with cryogenic propellants. To address this need, PCAD has enlisted Northrop Grumman Aerospace Systems (NGAS) in a technology development effort associated with the TR202, a LOX/LH2 expander cycle engine driven by independent turbopump assemblies and featuring a variable area pintle injector similar to the injector used on the TR200 Apollo Lunar Module Descent Engine (LMDE). Since the Apollo missions, NGAS has continued to mature deep throttling pintle injector technology. The TR202 program has completed two phases of pintle injector testing. The first phase of testing used ablative thrust chambers and demonstrated igniter operation as well as stable performance at several power levels across the designed 10:1 throttle range. The second phase of testing was performed on a calorimeter chamber and demonstrated injector performance at various power levels (75%, 50%, 25%, 10%, and 7.5%) across the throttle range as well as chamber heat flux to show that the engine can close an expander cycle design across the throttle range. This paper provides an overview of the TR202 program. It describes the different phases of the program with the key milestones of each phase. It then shows when those milestones were met. Next, it describes how the test data was used to update the conceptual design and how the test data has created a database for deep throttling cryogenic pintle technology that is readily scaleable and can be used to again update the design once the Altair program's requirements are firm. The final section of the paper describes the path forward, which includes demonstrating continuously throttling with an actuator and pursuing a path towards integrated engine sea-level test-bed testing.

Gromski, J.↗
DOE OSTI2025

Harnessing ferro-valleytricity in pentalayer rhombohedral graphene for memory and compute

Two-dimensional materials with multiple degrees of freedom, including spin, valleys, and orbitals, open up an exciting avenue for engineering multifunctional devices. Beyond spintronics, these degrees of freedom can lead to novel quantum effects such as valley-dependent Hall effects and orbital magnetism, which could revolutionize next-generation electronics. However, achieving independent control over valley polarization and orbital magnetism has been a challenge due to the need for large electric fields. A recent breakthrough involving pentalayer rhombohedral graphene has demonstrated the ability to individually manipulate anomalous Hall signals and orbital magnetic hysteresis, forming what is known as a valley-magnetic quartet. Here, we leverage the electrically tunable ferro-valleytricity of pentalayer rhombohedral graphene to develop nonvolatile memory and in-memory computation applications. We propose an architecture for a dense, scalable, and selector-less nonvolatile memory array that harnesses the electrically tunable ferro-valleytricity. In our designed array architecture, nondestructive read and write operations are conducted by sensing the valley state through two different pairs of terminals, allowing for independent optimization of read/write peripheral circuits. The power consumption of our PRG-based array is remarkably low, with only ∼6 nW required per write operation and ∼2.3 nW per read operation per cell. This consumption is orders of magnitude lower than that of the majority of state-of-the-art cryogenic memories. Additionally, we engineer in-memory computation by implementing majority logic operations within our proposed nonvolatile memory array without modifying the peripheral circuitry. In conclusion, our framework presents a promising pathway toward achieving ultra-dense cryogenic memory and in-memory computation capabilities.

2D materials↗
NASA NTRS1998

The Role of the Strutjet Engine in New Global and Space Markets

The Strutjet, discussed in previous IAF papers, was originally introduced as an enabling propulsion concept for single stage to orbit applications. Recent design considerations indicate that this systems also provides benefits supportive of other commercial non-space applications. This paper describes the technical progress of the Strutjet since 1997 together with a rationale why Rocket Based Combined Cycle Engines in general, and the Strutjet in particular, lend themselves uniquely to systems having the ability to expand current space and open new global 'rapid delivery' markets. During this decade, Strutjet technology has been evaluated in over 1000 tests. Its design maturity has been continuously improved and desired features, like simple variable geometry and low drag flowpath resulting in high performance, have been verified. In addition, data is now available which allows the designer, who is challenged to maximize system operability and economic feasibility, to choose between hydrogen or hydrocarbon fuels for a variety of application. The ability exists now to apply this propulsion system to various vehicles with a multitude of missions. In this paper, storable hydrocarbon and gaseous hydrogen Strutjet RBCC test data as accomplished to date and as planned for the future is presented, and the degree of required technology maturity achieved so far is assessed. Two vehicles, using cryogenic propane fuel Strutjet engines, and specifically designed for rapid point-to-point cargo delivery between Pacific rim locations are introduced, discussed, and compared.

Siebenhaar, A.↗
NASA NTRS2001

STS-102 Countdown Status Briefing

Joel Wells, NASA Public Affairs, introduces Pete Nickolenko, NASA Test Director, Glenn Chin, Leonardo Mission Manager, and Ed Priselac, Shuttle Weather Officer, in this STS-102 prelaunch press conference. Mr. Nickolenko gives an overview of the countdown and built-in hold times, the launch window, and prelaunch activities (such as activation and checkout of the onboard computer systems, closing the payload bay doors, servicing of the onboard cryogenic cell tanks, main engine tests, and power-up of the ground communications systems). Mr. Chin confirms that the payload is in the final flight configuration and is ready for launch. Mr. Priselac gives the weather forecast for the launch date. The men then answer questions from the press.

Source record↗
NASA NTRS2015

Use of Generalized Fluid System Simulation Program (GFSSP) for Teaching and Performing Senior Design Projects at the Educational Institutions

This paper describes the experience of the authors in using the Generalized Fluid System Simulation Program (GFSSP) in teaching Design of Thermal Systems class at University of Alabama in Huntsville. GFSSP is a finite volume based thermo-fluid system network analysis code, developed at NASA/Marshall Space Flight Center, and is extensively used in NASA, Department of Defense, and aerospace industries for propulsion system design, analysis, and performance evaluation. The educational version of GFSSP is freely available to all US higher education institutions. The main purpose of the paper is to illustrate the utilization of this user-friendly code for the thermal systems design and fluid engineering courses and to encourage the instructors to utilize the code for the class assignments as well as senior design projects. The need for a generalized computer program for thermofluid analysis in a flow network has been felt for a long time in aerospace industries. Designers of thermofluid systems often need to know pressures, temperatures, flow rates, concentrations, and heat transfer rates at different parts of a flow circuit for steady state or transient conditions. Such applications occur in propulsion systems for tank pressurization, internal flow analysis of rocket engine turbopumps, chilldown of cryogenic tanks and transfer lines, and many other applications of gas-liquid systems involving fluid transients and conjugate heat and mass transfer. Computer resource requirements to perform time-dependent, three-dimensional Navier-Stokes computational fluid dynamic (CFD) analysis of such systems are prohibitive and therefore are not practical. Available commercial codes are generally suitable for steady state, single-phase incompressible flow. Because of the proprietary nature of such codes, it is not possible to extend their capability to satisfy the above-mentioned needs. Therefore, the Generalized Fluid System Simulation Program (GFSSP1) has been developed at NASA Marshall Space Flight Center (MSFC) as a general fluid flow system solver capable of handling phase changes, compressibility, mixture thermodynamics and transient operations. It also includes the capability to model external body forces such as gravity and centrifugal effects in a complex flow network. The objectives of GFSSP development are: a) to develop a robust and efficient numerical algorithm to solve a system of equations describing a flow network containing phase changes, mixing, and rotation; and b) to implement the algorithm in a structured, easy-to-use computer program. The analysis of thermofluid dynamics in a complex network requires resolution of the system into fluid nodes and branches, and solid nodes and conductors as shown in Figure 1. Figure 1 shows a schematic and GFSSP flow circuit of a counter-flow heat exchanger. Hot nitrogen gas is flowing through a pipe, colder nitrogen is flowing counter to the hot stream in the annulus pipe and heat transfer occurs through metal tubes. The problem considered is to calculate flowrates and temperature distributions in both streams. GFSSP has a unique data structure, as shown in Figure 2, that allows constructing all possible arrangements of a flow network with no limit on the number of elements. The elements of a flow network are boundary nodes where pressure and temperature are specified, internal nodes where pressure and temperature are calculated, and branches where flowrates are calculated. For conjugate heat transfer problems, there are three additional elements: solid node, ambient node, and conductor. The solid and fluid nodes are connected with solid-fluid conductors. GFSSP solves the conservation equations of mass and energy, and equation of state in internal nodes to calculate pressure, temperature and resident mass. The momentum conservation equation is solved in branches to calculate flowrate. It also solves for energy conservation equations to calculate temperatures of solid nodes. The equations are coupled and nonlinear; therefore, they are solved by an iterative numerical scheme. GFSSP employs a unique numerical scheme known as simultaneous adjustment with successive substitution (SASS), which is a combination of Newton-Raphson and successive substitution methods. The mass and momentum conservation equations and the equation of state are solved by the Newton-Raphson method while the conservation of energy and species are solved by the successive substitution method. GFSSP is linked with two thermodynamic property programs, GASP2 and WASP3 and GASPAK4, that provide thermodynamic and thermophysical properties of selected fluids. Both programs cover a range of pressure and temperature that allows fluid properties to be evaluated for liquid, liquid-vapor (saturation), and vapor region. GASP and WASP provide properties of 12 fluids. GASPAK includes a library of 36 fluids. GFSSP has three major parts. The first part is the graphical user interface (GUI), visual thermofluid analyzer of systems and components (VTASC). VTASC allows users to create a flow circuit by a 'point and click' paradigm. It creates the GFSSP input file after the completion of the model building process. GFSSP's GUI provides the users a platform to build and run their models. It also allows post-processing of results. The network flow circuit is first built using three basic elements: boundary node, internal node, and branch.

Majumdar, A. K.↗
NASA NTRS2018

NASA's Space Launch System: Opportunities for Small Satellites to Deep Space Destinations

The first flight of NASA's new exploration-class launch vehicle, the Space Launch System (SLS), will test a myriad of systems designed to enable the next generation of deep space human spaceflight, while also providing the rare opportunity for 13 6U CubeSat-class payloads to be deployed in several locations along the flight path. The first mission of SLS and NASA's new Orion crew vehicle, Exploration Mission-1 (EM-1), will launch from upgraded facilities at Kennedy Space Center no earlier than fiscal year 2020. The initial Block 1 configuration for EM-1 will be capable of lofting at least 26 metric tons (t) of payload to the moon, with propulsion supplied by twin five-segment solid rocket boosters, four RS-25 engines and an Interim Cryogenic Propulsion Stage (ICPS). SLS will send Orion into a distant retrograde lunar orbit, paving the way for future missions to cislunar space and eventually Mars. The multidisciplinary small satellites for EM-1 derive from NASA research, as well as from international partners, industry and academia. Research subjects for the various smallsats include the moon, sun and an asteroid. Science objectives vary from characterizing the effects of radiation on living organisms (yeast) to landing the smallest spacecraft yet on the moon to supporting space weather research. Some of the payloads are technology demonstrations that will pave the way for more ambitious future missions that will be deployed by the more powerful SLS Block 1B configuration.

Robinson, Kimberly F.↗
NASA NTRS2023

Reliable Ignition of LOX Methane Propellants in Thermal Vacuum Environment

The goal of this project is to demonstrate reliable ignition of liquid oxygen-liquid methane at thermal vacuum conditions using a coil-on-plug (COP) sparking exciter. Testing will be performed using EP’s Frost-Mint cryogenic propellant test stand, using an engine updated based on data collected from previous engine testing at Plum Brook Station, to determine conditions under which the cryogenic propellants will consistently and reliably ignite. In FY22 the project successfully completed the following milestones. The cryogenic test stand Frost-Mint was upgraded to include an integrated vacuum tube and liquid nitrogen cooling system enabling cold vacuum ignition testing. The Integrated GN2/LN2 Ultra-Low-temperature Enclosure (IGLUE), a custom-built close-coupled heater exchanger shroud was designed, fabricated, tested and integrated into the test stand. The vacuum tube and IGLUE enabled the thruster body and propellants to be conditioned to temperatures and pressures desired. This work recreated the no-light conditions experienced during vehicle-level thermal vacuum testing and obtained thruster performance data and ignition imagery.

liquid oxygen↗
NASA NTRSDate not supplied

Green Run Modal Test of the NASA Space Launch System Core Stage

The Core Stage of the new NASA Space Launch System (SLS) is a 212-foot-tall rocket assembly—consisting primarily of two cryogenic propellant tanks, an engine section, and four RS-25 rocket engines—that will send crew and large payloads to the moon and beyond for NASA’s Artemis program. Prior to SLS assembly, the Core Stage completed a series of structural and functional tests in the B-2 Test Stand at Stennis Space Center, designated Green Run. The goal of Green Run was to verify analytical models, confirm proper subsystem operation, and test-fire all four RS-25 engines of the Core Stage. In January 2020, Green Run testing began with an experimental modal analysis test, performed by the Marshal Space Flight Center modal test team. A free-boundary test condition of the Core Stage was simulated as close as possible by suspending the massive launch vehicle from the B-2 Test Stand crane. Modal excitation was provided by a pair of 250-lb electro-dynamic shakers for multi-shaker random vibration testing, as well as a 12-pound instrumented hammer for impact testing. Modal response was measured with 550 accelerometer channels distributed on both the Core Stage and the B-2 Test Stand derrick crane. Following one very long day of testing, frequency response functions were calculated from the measured time histories in the target mode frequency range of 5 Hz to 15 Hz, and mode shapes, frequencies, and damping values were successfully estimated. The case-study presented in this paper will discuss Green Run, the SLS Core Stage, the modal test setup and execution, as well as a brief overview of the test results. Challenges associated with testing such a large, suspended structure in an outdoor environment will be discussed as well.

Space Launch System↗
NASA NTRSDate not supplied

Cryogenic Fuel Aviation – Challenges and Opportunities

NASA aeronautics goals include pioneering new technology to increase efficiency and reduce emissions from air travel. The feasibility of using of on-board cryogenic fuels systems in combination with fuel cells, high efficiency engines, and electric motors for high efficiency commercial transport aircraft is being examined at NASA in collaborations with external partners. Growing emphasis on next generation air transportation resulted in the development of strategic investigations into cryogenic fuels aviation, including roadmaps for maturing new technologies, needs to improve testing and manufacturing capabilities, and the requirements for establishing regulatory and certification standards. A NASA cross-organizational multidisciplinary project team is exploring new aircraft architectures, enabling materials, and the engineering challenges associated with cryogenic fuels for aircraft. The engineering challenges include high performance cryogenic, thermal management, and power/propulsion components and subsystems to tackle the challenging problem of development of commercially viable aircraft that would radically transform air transportation.

liquid natural gas↗
NASA NTRS2010

Northrop Grumman TR202 LOX/LH2 Deep Throttling Engine Technology Project Status

NASA's Propulsion and Cryogenic Advanced Development (PCAD) project is currently developing enabling propulsion technologies in support of future lander missions. To meet lander requirements, several technical challenges need to be overcome, one of which is the ability for the descent engine(s) to operate over a deep throttle range with cryogenic propellants. To address this need, PCAD has enlisted Northrop Grumman Aerospace Systems (NGAS) in a technology development effort associated with the TR202 engine. The TR202 is a LOX/LH2 expander cycle engine driven by independent turbopump assemblies and featuring a variable area pintle injector similar to the injector used on the TR200 Apollo Lunar Module Descent Engine (LMDE). Since the Apollo missions, NGAS has continued to mature deep throttling pintle injector technology. The TR202 program has completed two series of pintle injector testing. The first series of testing used ablative thrust chambers and demonstrated igniter operation as well as stable performance at discrete points throughout the designed 10:1 throttle range. The second series was conducted with calorimeter chambers and demonstrated injector performance at discrete points throughout the throttle range as well as chamber heat flow adequate to power an expander cycle design across the throttle range. This paper provides an overview of the TR202 program, describing the different phases and key milestones. It describes how test data was correlated to the engine conceptual design. The test data obtained has created a valuable database for deep throttling cryogenic pintle technology, a technology that is readily scalable in thrust level.

Gromski, Jason↗
NASA NTRS2008

Liquid Methane/Liquid Oxygen Propellant Conditioning Feed System (PCFS) Test Rigs

As part of their Propulsion and Cryogenic Advanced Development (PCAD) program, NASA has embarked upon an effort to develop chemical rocket engines which utilize non-toxic, cryogenic propellants such as liquid oxygen (LO2) and liquid methane (LCH4). This effort includes the development and testing of a 100 lbf Reaction Control Engine (RCE) that will be used to evaluate the performance of a LO2/LCH4 rocket engine over a broad range of propellant temperatures and pressures. This testing will take place at NASA-Glenn Research Center's (GRC) Research Combustion Laboratory (RCL) test facility in Cleveland, OH, and is currently scheduled to begin in late 2008. While the initial tests will be performed at sea level, follow-on testing will be performed at NASA-GRC's Altitude Combustion Stand (ACS) for altitude testing. In support of these tests, Sierra Lobo, Inc. (SLI) has designed, developed, and fabricated two separate portable propellant feed systems under the Propellant Conditioning and Feed System (PCFS) task: one system for LCH4, and one for LO2. These systems will be capable of supplying propellants over a large range of conditions from highly densified to several hundred pounds per square inch (psi) saturated. This paper presents the details of the PCFS design and explores the full capability of these propellant feed systems.

Skaff, A.↗