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The tests described in this paper afford a direct comparison of the efficiency and smoothness of flow obtained with propeller fan and multiblade centrifugal fan drives in the same wind tunnel. The propeller fan was found to be superior to the centrifugal fan in that the efficiency was about twice as great, and the flow much smoother.
NASA and its commercial industry team achieved significant progress in 2016 in manufacturing and testing of the Block 1 vehicle for the first launch of the Space Launch System (SLS). Test and flight article hardware for the liquid hydrogen fuel tank as well as the engine section for the core stage were completed at Michoud Assembly Facility (MAF) in New Orleans. Test stands neared completion at Marshall Space Flight Center for the propellant tanks, engine section, intertank and payload section. Stennis Space Center completed major structural renovations on the B2 test stand, where the core stage "green run" test program will be conducted. The SLS team completed a hotfire test series at Stennis to successfully demonstrate the ability of the RS-25 engine to operate under SLS environments and performance conditions. The team also test fired the second qualification five-segment solid rocket motor and cast the first six motor segments for the first SLS mission. The Interim Cryogenic Propulsion Stage (ICPS) test article was delivered to Marshall for structural tests, and work is nearly finished on the flight stage. Flight software testing completed at Marshall included power quality and command and data handling. In 2017, that work continues. SLS completed Preliminary Design Review (PDR) on the Exploration Upper Stage (EUS), a powerful, human-rated spacecraft that will propel explorers to cis-lunar space. In 2017, hardware will continue to be integrated at MAF for core stage structural test articles and the first two operational flights. RS-25 hotfire testing will continue to explore engine performance, as well as test flight-like software and four new Engine Controller Units (ECUs) for the first mission. Production of development components for a more affordable RS-25 design is underway. Core stage structural test articles have begun arriving at Marshall. While engineering challenges typical of a new development are possible, SLS is working toward launch readiness in late 2018. This paper will discuss these and other technical and programmatic successes and challenges over the past year and provide a preview of work ahead before first flight
Unsteady aerodynamics of a pusher propeller operating in the wake of a pylon is investigated through analysis of flight test data. Twenty-two surface-mounted miniature pressure transducers were installed at the 75 and 90 percent radius locations on a propeller blade on a test-bed aircraft. Twenty-six different flight conditions were flown to cover the range of speeds, rpm, altitudes, and flap positions encountered by an advanced turboprop general aviation aircraft. Preliminary analysis of the flight test data indicate a strong three-dimensionality to the perturbed flow at the 75 and 90 percent radii due to pylon wake encounter. Time-history of the pressure transducer waveforms recorded for 700 revolutions exhibit a one-per-cycle wave oscillation on the wake signature. No explanation is found for this behavior yet.
The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation ion propulsion system with significant enhancements beyond the state-of-the-art in ion propulsion to provide future NASA science missions with enhanced mission capabilities at a low total development cost. As part of a comprehensive thruster service life assessment utilizing both testing and analyses, a Long-Duration Test (LDT) was initiated to validate and qualify the NEXT propellant throughput capability to a qualification-level of 450 kg, 1.5 times the mission-derived throughput requirement of 300 kg. This wear test is being conducted with a modified, flight-representative NEXT engineering model ion thruster, designated EM3. As of June 21, 2007, the thruster has accumulated 10,100 hr of operation at the thruster full-input-power of 6.9 kW with 3.52 A beam current and 1800 V beam power supply voltage. The thruster has processed 207 kg of xenon and demonstrated a total impulse of 8.5 106 N-s; the highest total impulse ever demonstrated by an ion thruster in the history of space propulsion. Thruster performance tests are conducted periodically over the entire NEXT throttle table with input power ranging 0.5 to 6.9 kW. Overall ion thruster performance parameters including thrust, input power, specific impulse, and thruster efficiency have been nominal with little variation to date. Lifetime-limiting component erosion rates have been consistent with the NEXT service life assessment, which predicts the earliest failure sometime after 750 kg of xenon propellant throughput; well beyond the mission-derived lifetime requirement. The NEXT wear test data confirm that the erosion of the discharge keeper orifice, enlarging of nominal-current-density accelerator grid aperture cusps, and the decrease in cold grid-gap observed during the NSTAR Extended Life Test have been mitigated. This paper presents the status of the NEXT LDT to date.
The lack of slipstream static pressure distribution seriously affected the results but recommendations for removing the deficiency are discussed. The wake survey rake is shown to be a valuable tool in aircraft flight testing. Flow characteristics in the wake of the propeller were examined.
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The results of the Resonant Infrasonic Gauging System (RIGS) development program are reported. The RIGS is a gauging system that is capable of measuring propellant quantity under zero-G as well as under accelerated (one-G) conditions. With the exception of liquid hydrogen, it can be used to gauge virtually any propellant in liquid form including cryogenics. The gauge consists of a sensor unit which is attached to the propellant tank and electronic control unit which may be positioned separately from the sensor. The control unit receives the signals from the sensor as well as the ullage gas pressure and propellant temperature measurements, and computes the propellant quantity remaining in the tank. During the course of this program two prototype RIGS sensors were designed and constructed. The sensors were tested first in the laboratory using water as the simulated propellant and, later, using LN2 in a 100-gallon tank. The system tests proved that the gauge operates virtually as predicted by theory and yielded an accuracy better than 1%.
Tests were conducted with the RS18 rocket engine using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA s Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propellant systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to approx.120,000 ft (approx.37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. Propellant flow rate was measured using a coriolis-style mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup. LO2 flow ranged from 5.9-9.5 lbm/sec (2.7-4.3 kg/sec), and LCH4 flow varied from 3.0-4.4 lbm/sec (1.4-2.0 kg/sec) during the RS-18 hot-fire test series. Thrust was measured using three load cells in parallel. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. Data was obtained at multiple chamber pressures, and calculations were performed for specific impulse, C* combustion efficiency, and thrust vector alignment. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes.
The NASA White Sands Test Facility (WSTF) routinely operates hypergolic propulsion systems. Some of the onsite activities include performing long duration studies on the operational life of these systems. A few of them have been in use for over twenty years. During this span of time contamination has built up in the propellant and some of the distribution infrastructure. This study investigated the nature of this contamination, the pathology of its generation, and developed a process for removal of the contamination that was cost efficient with minimal waste generation.
The throttling capabilities of a 30 kW class ammonia arcjet and its compatibility with a breadboard power conditioning unit (PCU) were tested in two series of tests. The first series was performed to determine the performance and operating characteristics of the arcjet and the PCU over a range of power levels and propellant flow rates. The power levels for the tests were nominally between 10 and 30 kW, with some operation below 10 kW at the lower flow rates. The ammonia flow rates varied between 0.16 and 0.35 g/s. The second series of tests was an extensive investigation of operation below 12 kW using three cathode spacings. The ammonia flow rates were between 0.115 and 0.335 g/s. Operation of the arcjet from 1.5 kW up to the 30 kW design point was demonstrated with the PCU.
Apollo lunar module cryogenically stored supercritical He propellant tank pressurization system in descent propulsion subsystem, discussing design and performance tests
The model was tested at two different elevations with the wing pivot at 1.008 and 2.425 propeller diameters above the ground. The slipstream of the propellers was deflected by tilting the wing and propellers, by deflections of large-chord trailing-edge flaps, and by combinations of flap deflection and wing tilt. Tests were conducted over a range of propeller disk loadings from 7.41 to 29.70 pounds per square foot. Force data for the complete model and pressure distributions for the wing and flaps behind one propeller were recorded and are presented in tabular form without analysis.
The principal objective of this two phase program is to conduct the development research required to make the low power arcjet a flight ready technology. Many important results were obtained during Phase 1 to move closer to this objective. Fundamental analyses were performed of the arcjet nozzle, the gas kinetic reaction effects, the thermal environment, and the arc stabilizing vortex. These aided the conceptual understanding of the arcjet and guided design work. A hydrazine (N2H4) arcjet was designed that combined a flight qualified catalyst bed with a modular arcjet. Extensive testing was performed which demonstrated the feasibility of using this propellant in an arcjet for the first time. Startup techniques were developed, stability maintained, material compatibility tests conducted, and performance mapping tests performed. Specific impulse values from 400 to 730 seconds were produced with a non-optimized design. These levels are higher than were originally thought possible and proved that extremely high enthalpy values can be obtained with constricted arc technology. Erosion rate data are promising for lifetime extensions to meet flight application requirements. Power control unit (PCU) development was started with the design and fabrication of a laboratory high switching frequency supply. Valuable data were obtained on PCU operation and on the interaction with the dynamic arc. Phase 2 efforts presently underway are resolving key issues for multi-hundred hour lifetimes, are continuing to investigate arcjet/PCU interactions, and will demonstrate duty cycle N2H4 arcjet/PCU operation in a simulated flight mode for lifetimes consistent with initial applications.
This investigation was made for the purpose of determining the characteristics of five full-scale propellers in flight. The equipment consisted of five propellers in conjunction with a VE-7 airplane and a Wright E-2 engine. The propellers were of the same diameter and aspect ratio. Four of them differed uniformly in thickness and pitch and the fifth propeller was identical with one of the other four with exception of a change of the airfoil section. The propeller efficiencies measured in flight are found to be consistently lower than those obtained in model tests. It is probable that this is mainly a result of the higher tip speeds used in the full-scale tests. The results show also that because of differences in propeller deflections it is difficult to obtain accurate comparisons of propeller characteristics. From this it is concluded that for accurate comparisons it is necessary to know the propeller pitch angles under actual operating conditions. (author)
NASA is embarked on a new era of space exploration that will lead to new capabilities, new destinations, and new discoveries by both human and robotic explorers. Today, the International Space Station (ISS), supported by NASA's commercial partners, and robotic probes, are yielding knowledge that will help make this exploration possible. NASA is developing both the Orion crew vehicle and the Space Launch System (SLS) that will carry out a series of increasingly challenging missions that will eventually lead to human exploration of Mars. This paper will discuss the development and progress on the SLS. The SLS architecture was designed to be safe, affordable, and sustainable. The current configuration is the result of literally thousands of trade studies involving cost, performance, mission requirements, and other metrics. The initial configuration of SLS, designated Block 1, will launch a minimum of 70 metric tons (t) into low Earth orbit - significantly greater capability than any current launch vehicle. It is designed to evolve to a capability of 130 t through the use of upgraded main engines, advanced boosters, and a new upper stage. With more payload mass and volume capability than any rocket in history, SLS offers mission planners larger payloads, faster trip times, simpler design, shorter design cycles, and greater opportunity for mission success. Since the program was officially created in fall 2011, it has made significant progress toward first launch readiness of the Block 1 vehicle in 2018. Every major element of SLS continued to make significant progress in 2015. The Boosters element fired Qualification Motor 1 (QM-1) in March 2015, to test the 5-segment motor, including new insulation, joint, and propellant grain designs. The Stages element marked the completion of more than 70 major components of test article and flight core stage tanks. The Liquid Engines element conducted seven test firings of an RS-25 engine under SLS conditions. The Spacecraft/Payload Integration and Evolution element marked completion of the upper stage test article. Major work continues in 2016 as the program continues both flight and development RS-25 engine testing, begins welding test article and flight core stage tanks, completes stage adapter manufacturing, and test fires the second booster qualification motor. This paper will discuss the program's key accomplishments to date and the challenging work ahead for what will be the world's most capable launch vehicle.
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.
ABSTRACT An environmentally friendly, cost effective technology has been developed and demonstrated by a team of Naval Air Warfare Center and Lockheed Martin personnel to dispose of Shillelagh solid rocket motor propellants. The Shillelagh is a surface to surface anti-tank weapon approaching the end of its service life. The current demilitarization process employs open detonation, but the presence of lead stearate in the N5 propellant grain motivated the need for the development of an environmentally friendly disposal method. Contained burning of the propellant followed by propellant exhaust processing was chosen as the disposal methodology. The developmental test bed, completed in February 1998, is inexpensive and transportable. Contained burning of Shillelagh propellants posed two technical hurdles: 1) removal of the sub micron lead and cadmium particulate generated during combustion, and 2) secondary combustion of the significant quantifies of carbon monoxide and hydrogen. A firing chamber with a stepped nozzle, air injection, and active ignition was developed to combust the carbon monoxide and hydrogen in real time. The hot gases and particulates from the combustion process are completely contained within a gas holder. The gases are subsequently cooled and routed through a treatment facility to remove the heavy metal particulate. Results indicate that the lead and cadmium particulates are removed below their respective detection limits (2 micro-g/cu m & 0.2 micro-g/cu m) of the analytical procedures employed and that the carbon monoxide and hydrogen levels have been reduced well below the lower flammability limits. Organic concentrations, principally benzene, are I ppm or less. A semi-automated machine has been developed which can rapidly prepare Shillelagh missiles for the contained burn facility. This machine allows the contained burn technology to be more competitive with current open bum open detonation disposal rates.
While evaluating various options for liquid methane and liquid oxygen propellant management for lunar missions, Innovative Engineering Solutions (IES) conceived the flexible screen device as a potential simple alternative to conventional propellant management devices (PMD). An apparatus was designed and fabricated to test flexible screen devices in liquid nitrogen. After resolution of a number of issues (discussed in detail in the paper), a fine mesh screen (325 by 2300 wires per inch) spring return assembly was successfully tested. No significant degradation in the screen bubble point was observed either due to the screen stretching process or due to cyclic fatigue during testing. An estimated 30 to 50 deflection cycles, and approximately 3 to 5 thermal cycles, were performed on the final screen specimen, prior to and between formally recorded testing. These cycles included some "abusive" pressure cycling, where gas or liquid was driven through the screen at rates that produced differential pressures across the screen of several times the bubble point pressure. No obvious performance degradation or other changes were observed over the duration of testing. In summary, it is felt by the author that these simple tests validated the feasibility of the flexible screen PMD concept for use with cryogenic propellants.