An investigation of ejectors having short mixing sections.
Ejectors with short mixing sections used to simulate altitudes for rocket engine testing in vertical position
ENGINE TESTING↗
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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
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NASA Plum Brook Station In-Space Propulsion Facility Test Stand Characterization Hot Fire Test
A test facility modification to enable small scale altitude propulsion testing at the NASA Glenn Research Center's In-Space Propulsion (ISP) Facility was verified with a hot fire test campaign. As the facility's primary steam supply system undergoes refurbishment, the alternate facility configuration, known as the "vacuum accumulator" mode, would enable rocket engine testing up to 10,000 lbf thrust. The NASA Johnson Space Center developed the vehicle for the verification test campaign: the Integrated Cryogenic Propulsion Test Article (ICPTA). Constructed primarily from assets of the former Morpheus Project, the ICPTA provided an integrated liquid oxygen (LOX) / liquid methane (LCH4) propulsion system including a 2,800 lbf thrust main engine. The ISP Facility's vacuum accumulator configuration leveraged the large test volume of the facility and a diffuser insert to maintain altitude conditions. During hot fire, the ICPTA main engine "started" the diffuser insert constructed for the test campaign. As a result, the test chamber upstream of the diffuser insert remained at altitude conditions throughout the hot fire. Upon engine shut down, a backflow deflector mitigated blow back into the test chamber by restricting the mass flow and redirecting it away from the test article. The test campaign successfully characterized the performance of the vacuum accumulator configuration. In addition, it provided an opportunity to collect data for an integrated LOX / LCH4 propulsion system in an altitude and thermal vacuum environment.
Rocket Engine Testing↗
Liquid Oxygen/Liquid Methane Test Results of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility
Tests were conducted with the RS-18 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 propulsion 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 ~122,000 ft (~37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. 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. Propellant flow rate was measured using a coriolis mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup due to two-phase flow effects. Subsequent cold-flow testing demonstrated that the propellant manifolds must be adequately flushed in order for the coriolis flow meters to give accurate data. The coriolis flow meters were later shown to provide accurate steady-state data, but the turbine flow meter data should be used in transient phases of operation. Thrust was measured using three load cells in parallel, which also provides the capability to calculate thrust vector alignment. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. 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. All of these objectives were met with the RS-18 data and additional testing data from subsequent LO2/methane test programs in 2009 which included the first simulated-altitude pyrotechnic ignition demonstration of LO2/methane.
Melcher, John C., IV↗
GT-2 flight test data. Book XII - Timing, pressure and altitude comparisons
Pressure and altitude data for GT-2 flight test
ALTITUDE↗
Liquid Oxygen/Liquid Methane Test Summary of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility
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.
Melcher, John C., IV↗
Second Generation Prototype Design and Testing for a High Altitude Venus Balloon
This paper describes the development of a second generation prototype balloon intended for flight in the upper atmosphere of Venus. The design of this new prototype incorporates lessons learned from the construction and testing of the first generation prototype, including finite element analyses of the balloon stresses and deformations, measured leak performance after handling and packaging, permeability and optical property measurements on material samples, and sulfuric acid testing. An improved design for the second generation prototype was formulated based on these results, although the spherical shape and 5.5 m diameter size were retained. The resulting balloon has a volume of 87 cubic meters and is capable of carrying a 45 kg payload at a 55 km altitude at Venus. The design and fabrication of the new prototype is described, along with test data for inflation and leakage performance.
mobility↗
Space Suit Electrocardiographic Electrode Selection: Are commercial electrodes better than the old Apollo technology?
The NASA Manned Space Program uses an electrocardiograph (ECG) system to monitor astronauts during extravehicular activity (EVA). This ECG system, called the Operational Bioinstrumentation System (OBS), was developed during the Apollo era. Throughout the Shuttle program these electrodes experienced failures during several EVAs performed from the Space Shuttle and International Space Station (ISS) airlocks. An attempt during Shuttle Flight STS-109 to replace the old electrodes with new commercial off-the-shelf (COTS) disposable electrodes proved unsuccessful. One assumption for failure of the STS-109 COTS electrodes was the expansion of trapped gases under the foam electrode pad, causing the electrode to be displaced from the skin. Given that our current electrodes provide insufficient reliability, a number of COTS ECG electrodes were tested at the NASA Altitude Manned Chamber Test Facility. Methods: OBS disposable electrodes were tested on human test subjects in an altitude chamber simulating an Extravehicular Mobility Unit (EMU) operating pressure of 4.3 psia with the following goals: (1) to confirm the root cause of the flight certified, disposable electrode failure during flight STS-109. (2) to identify an adequate COTS replacement electrode and determine if further modifications to the electrodes are required. (3) to evaluate the adhesion of each disposable electrode without preparation of the skin with isopropyl alcohol. Results: There were several electrodes that failed the pressure testing at 4.3psia, including the electrodes used during flight STS-109. Two electrodes functioned well throughout all testing and were selected for further testing in an EMU at altitude. A vent hole placed in all electrodes was also tested as a possible solution to prevent gas expansion from causing electrode failures. Conclusions: Two failure modes were identified: (1) foam-based porous electrodes entrapped air bubbles under the pad (2) poor adhesion caused some electrodes to fail
Redmond, M.↗
Ignition Delay Experiments with Small-scale Rocket Engine at Simulated Altitude Conditions Using Various Fuels with Nitric Acid Oxidants / Dezso J. Ladanyi
Ignition delay determinations of several fuels with nitric oxidants were made at simulated altitude conditions utilizing a small-scale rocket engine of approximately 50 pounds thrust. Included in the fuels were aniline, hydrazine hydrate, furfuryl alcohol, furfuryl mercaptan, turpentine, and mixtures of triethylamine with mixed xylidines and diallyaniline. Red fuming, white fuming, and anhydrous nitric acids were used with and without additives. A diallylaniline - triethylamine mixture and a red fuming nitric acid analyzing 3.5 percent water and 16 percent NO2 by weight was found to have a wide temperature-pressure ignition range, yielding average delays from 13 milliseconds at 110 degrees F to 55 milliseconds at -95 degrees F regardless of the initial ambient pressure that ranged from sea-level pressure altitude of 94,000 feet.
XYLIDINE↗
Effects of Altitude on Turbojet Engine Performance
Component and over-all performance characteristics of several turbojet engines investigated in the altitude facilities of the NACA Lewis Laboratory during the last several years are summarized to indicate the effects of altitude on turbojet engine performance. Data presented show that failure of turbojet engine performance to generalize for all altitudes can be traced to reductions in compressor efficiency, corrected air flow, and combustion efficiency at altitude. In addition, it is shown that although engines of different design may have equal thrusts at sea level, the thrusts at altitude may vary widely because of differences in compressor performance characteristics from one engine to another.
COMPRESSORS - EFFICIENCY↗
Improved Exhaust Diffuser for Jet-Engine Testing
High-altitude simulator reduced power requirements. Test cell uses its exhaust-capture duct only to remove gases from engine; cooling air evacuated through separate path by auxiliary suction system. This way, capture duct cross-sectional area kept close to exhaust jet area, leading to greatly improved recovery performance.
Parikh, P. G.↗
Performance measurement of low concentration ratio solar array for space application
The measured performance of a silicon and a gallium arsenide low concentration ratio solar array (LCRSA) element is presented. The element characteristics measured in natural sunlight are off pointing performance and response to mechanical distortions. Laboratory measurements of individual silicon and gallium-arsenide solar cell assemblies are also made. The characteristics measured in the laboratory involved responses to temperature and intensity variations as well as to the application of reverse bias potentials. The element design details covered include the materials, the solar cells, and the rationale for selecting these specific characteristics. The measured performance characteristics are contrasted with the predicted values for both laboratory testing and high altitude natural sunlight testing. Excellent agreement between analytical predictions and measured performance is observed.
Mills, M. W.↗
Atmospheric analysis for airdata calibration on research aircraft
In-flight airdata calibrations are used to determine the aerodynamic influence of an airplane on pitot-static pressure measurements of altitude and speed. Conventional flight-test calibration techniques are briefly reviewed and meteorological analysis methods for estimating calibration reference values of atmospheric conditions are described. There are cases where some conventional in-flight techniques are not entirely satisfactory for research aircraft because of added equipment requirements or flight envelope and location limitations. In these cases, atmospheric wind and pressure information can be used to complement conventional techniques. Accuracy of the atmospheric measurements and the variability of upper-air winds and pressure values are discussed. Results from several flight research aircraft show that wind reference calibration is generally less accurate than calibration accuracy standards for civil and research aircraft. Examples of pressure reference altimetry derived from meteorological analyses are also presented for a variety of flight research programs. These flight data show that the reference pressure accuracy provided by meteorological analyses is usually within civil aircraft and flight research airdata calibration accuracy standards. Meteorological analyses altimetry is particularly useful when it is not feasible to restrict the test airplane altitude, location, or maneuver envelope.
Ehernberger, L. J.↗
Effect of altitude on power of aviation engines
These notes are intended to furnish practical and general data on the effect of altitude on engine power. The effective horsepower of an engine is a function of the mean pressure of the fluid acting on the pistons, of the R.P.M. of the engine and of the mechanical efficiency.
ENGINES - PERFORMANCE↗
Collaborative Systems Engineering in the Ascent Abort-2 Crew Module/Separation Ring Project
Generally speaking, systems engineering (SE) tool-sets face a dilemma balancing power and accessibility. High-powered SE tools (MagicDraw, Cradle, Core, etc.) tend to be specialized and are available only to highly trained Systems Engineers, and/or through the use of a 'back room' developer team making the output products available to the broader team. On the other hand, highly accessible tools (MS Word, Excel, etc.) do not have the power to implement SE in a rigorous manner. NASA has to test all aspects of the new human-rated Orion Multi-Purpose Crew Vehicle spacecraft prior to its first crewed mission. The test program includes uncrewed launch abort flight tests to demonstrate the capability to save the crew in the event that a launch failure occurs. Orion's second abort flight test will be a low-altitude flight test known as "Ascent Abort 2 (AA-2)." This test is currently scheduled to be carried out at Cape Canaveral Air Force Station's Space Launch Complex 46 (SLC-46) in Florida in 2019. NASA's in-house AA-2 Crew Module and Separation Ring (CSR) Team is producing the crew module and separation ring. Operating jointly as both an Advanced Exploration Systems (AES) Project and an Orion Project, the CSR project charter includes development of innovative, streamlined and generally more efficient practices for creation of flight hardware and software. One result of this tasking has been development of a collaborative and data-centric systems engineering environment within the team's shared web environment (Microsoft SharePoint). Through the use of built-in, 'out of the box capabilities' present in MS SharePoint, the CSR Systems Engineering team has created (with some limited developer support) a data-centric architecture for the project's SE implementation, including functional and interface analysis, requirements development and management, risk management, verification planning and management, test results, and end item management. Data elements are linked between data structures so as to define and control relationships between item types, link requirements to parents and children, and link tests to the requirements that they verify. The overall project team integration is increased by also linking SE content to project management content over the project life cycle, including team communication, action items, configuration management, decisional and meeting materials, and life cycle reviews. This presentation will provide an overview of the collaborative SE environment, showing how it provides the power for a number of SE tasks while still providing the accessibility and transparency to allow the full project team to collaborate and succeed. Given the project phase, we'll be able to present a nearly full lifecycle discussion, from concept through verification and approaching delivery.
Systems Engineering environments↗
Maraia Capsule Flight Testing and Results for Entry, Descent, and Landing
The Maraia concept is a modest size (150 lb., 30" diameter) capsule that has been proposed as an ISS based, mostly autonomous earth return capability to function either as an Entry, Descent, and Landing (EDL) technology test platform or as a small on-demand sample return vehicle. A flight test program has been completed including high altitude balloon testing of the proposed capsule shape, with the purpose of investigating aerodynamics and stability during the latter portion of the entry flight regime, along with demonstrating a potential recovery system. This paper includes description, objectives, and results from the test program.
Sostaric, Ronald R.↗
SCALE MODEL TESTING OF 90 DEG TURN EJECTOR SYSTEMS FOR ROCKET ENGINE ALTITUDE SIMULATION
Scale model testing of 90 degree turn ejector systems for rocket engine altitude simulation
NUCLEAR ROCKET↗
Development flight tests of the Viking decelerator system.
Significant aspects of a low altitude flight test phase of the overall Viking decelerator system development are given. This test series included nine aircraft drop tests that were conducted at the Joint Parachute Test Facility, El Centro, California, between September 1971 and May 1972. The test technique and analytical planning method utilized to best simulate loading conditions in a low density environment are presented and some test results are shown to assess their adequacy. Performance effects relating to suspension line lengths of 1.7 D sub o with different canopy loadings are noted. System hardware developments are described, in particular the utilization of a fabric deployment mortar cover which remained attached to the parachute canopy. Finally, the contribution of this test series to the overall program is assessed.
Murrow, H. N.↗
Investigation of the turbulent wind field below 150-meter altitude at the eastern test range
Statistical effects of wind and atmospheric turbulence on launch vehicles
Blackadar, A. K.↗