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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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At least 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1947

Aerodynamic Characteristics at High Speeds of Full-Scale Propellers having Different Shank Designs

Tests of two 10-foot-diameter two-blade propellers which differed only in shank design have been made in the Langley 16-foot high-speed tunnel. The propellers are designated by their blade design numbers, NACA 10-(5)(08)-03, which had aerodynamically efficient airfoil shank sections, and NACA l0-(5)(08)-03R which had thick cylindrical shank sections typical of conventiona1 blades, The propellers mere tested on a 2000-horsepower dynamometer through a range of blade-angles from 20deg to 55deg at various rotational speeds and at airspeeds up to 496 miles per hour. The resultant tip speeds obtained simulate actual flight conditions, and the variation of air-stream Mach number with advance ratio is within the range of full-scale constant-speed propeller operation. Both propellers were very efficient, the maximum envelope efficiency being approximately 0,95 for the NACA 10-(5)(08)-03 propeller and about 5 percent less for the NACA 10-(5)(08)-03R propeller. Based on constant power and rotational speed, the efficiency of the NACA 10-(05)(08)-03 propeller was from 2.8 to 12 percent higher than that of the NACA 10-(5)(08)-03R propeller over a range of airspeeds from 225 to 450 miles per hour. The loss in maximum efficiency at the design blade angle for the NACA 10-(5)(08)-03 and 10-(5)(08)-03R propellers vas about 22 and 25 percent, respectively, for an increase in helical tip Mach number from 0.70 to 1.14.

Maynard, Julian D.↗
NASA NTRS2013

Thermal Design and Analysis of the Supersonic Flight Dynamics Test Vehicle for the Low Density Supersonic Decelerator Project

The thermal design and analysis of the experimental Supersonic Flight Dynamics Test (SFDT) vehicle is presented. The SFDT vehicle is currently being designed as a platform to help demonstrate key technologies for NASA's Low Density Supersonic Decelerator (LDSD) project. The LDSD project is charged by NASA's Office of the Chief Technologist (OCT) with the task of advancing the state of the art in Mars Entry, Descent, and Landing (EDL) systems by developing and testing three new technologies required for landing heavier payloads on Mars. The enabling technologies under development consist of a large 33.5 meter diameter Supersonic Ringsail (SSRS) parachute and two different types of Supersonic Inflatable Aerodynamic Decelerator (SIAD) devices - a robotic class, SIAD-R, that inflates to a 6 meter diameter torus, and an exploration class, SIAD-E, that inflates to an 8 meter diameter isotensoid. As part of the technology development effort, the various elements of the new supersonic decelerator system must be tested in a Mars-like environment. This is currently planned to be accomplished by sending a series of SFDT vehicles into Earth's stratosphere. Each SFDT vehicle will be lifted to a stable float altitude by a large helium carrier balloon. Once at altitude, the SFDT vehicles will be released from their carrier balloon and spun up via spin motors to provide trajectory stability. An onboard third stage solid rocket motor will propel each test vehicle to supersonic flight in the upper atmosphere. After main engine burnout, each vehicle will be despun and testing of the deceleration system will begin: first an inflatable decelerator will be deployed around the aeroshell to increase the drag surface area, and then the large parachute will be deployed to continue the deceleration and return the vehicle back to the Earth's surface. The SFDT vehicle thermal system must passively protect the vehicle structure and its components from cold temperatures experienced during the ascent phase of the mission as well as from the extreme heat fluxes produced during the supersonic test phase by the main motor plume and aeroheating. The passive thermal design approach for the SFDT vehicle relies upon careful and complex bounding analysis of all three modes of heat transfer - conduction, convection, and radiation - coupled with a tightly managed transient power dissipation timeline for onboard electronics components throughout all mission phases.

thermal desktop↗
NASA NTRS1992

Advanced small rocket chambers: Option 1, 14 lbf Ir-Re rocket

A high performance Ir-Re 14 lbf (62 N) chamber and nozzle which can be a direct replacement for a production engine was designed, built, hot fired and vibration acceptance tested. It passed all acceptance tests satisfactorily and demonstrated a 20 sec increase in specific impulse (Is) over the conventional 14 lbf silicide coated Cb chamber. The high performance engine uses the production valve and injector without modification. Incorporation of a secondary mixing device or Boundary Layer Trip within the combustion chamber results in elimination of the fuel film coolant, improvement in flow uniformity, the 20 sec performance increase, and reduction of a potential source of spacecraft contamination. Measured Is was 305 sec at 75:1 area ratio, with monomenthylhydrazine and nitrogen tetroxide propellants. Qualification tests remain to be done.

Jassowski, Donald M.↗
NASA NTRS1989

Low-speed wind tunnel performance of high-speed counterrotation propellers at angle-of-attack

The low-speed aerodynamic performance characteristics of two advanced counterrotation pusher-propeller configurations with cruise design Mach numbers of 0.72 were investigated in the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel. The tests were conducted at Mach number 0.20, which is representative of the aircraft take/off/landing flight regime. The investigation determined the effect of nonuniform inflow on the propeller performance characteristics for several blade angle settings and a range of rotational speeds. The inflow was varied by yawing the propeller mode to angle-of-attack by as much as plus or minus 16 degrees and by installing on the counterrotation propeller test rig near the propeller rotors a model simulator of an aircraft engine support pylon and fuselage. The results of the investigation indicated that the low-speed performance of the counterrotation propeller configurations near the take-off target operating points were reasonable and were fairly insensitive to changes in model angle-of-attack without the aircraft pylon/fuselage simulators installed on the propeller test rig. When the aircraft pylon/fuselage simulators were installed, small changes in propeller performance were seen at zero angle-of-attack, but fairly large changes in total power coefficient and very large changes of aft-to-forward-rotor torque ratio were produced when the propeller model was taken to angle-of-attack. The propeller net efficiency, though, was fairly insensitive to any changes in the propeller flowfield conditions near the take-off target operating points.

Hughes, Christopher E.↗
NASA NTRS1987

Compatability of dispersion-strengthened platinum with resistojet propellants

Resistojets for the Space Station require long life and multipropellant capability. The choice of available materials to meet these requirements is limited. Dispersion-strengthened platinum was selected. Past results indicated that it should be suffieiently inert in candidate propellant environments and should be capable of operating at moderate temperatures for extended periods. A series of propellant compatibility tests was done with platinum strengthened with either yttria or zirconia. Data presented included the results of 1000-hr tests in CO2, H2, ammonia (NH3), N2, steam, hydrazine (N2H4), and methane (CH4); and 2000-hr tests in H2 and NH3. The platinum samples were tested at 1400 C in CO2, H2, NH3, N2, steam, and N2H4; at 500 C in CH4; and at 800 C in N2H4. The mass-loss results indicated material life, exptrapolated from experimental mass-loss data, in excess of 100 000 hr in all environments except steam and N2H4, where it was greater than or =45000 hr. Generally, on the basis of mass loss, there were no compatibility concerns in any of the environments considered. Optical and scanning electron microscopy were used to determine the effect of propellants on the material surface and to evaluate material stability.

Whalen, Margaret V.↗
NASA NTRSDate not supplied

NGHT-1X Pole Cover Erosion Measurements on Xenon and Krypton

Northrop Grumman’s NGHT-1X Hall thruster has completed the Engineering Model development phase and is progressing towards qualification and flight unit build and test. First flight of the NGHT-1X will occur in 2025 on Northrop Grumman’s Mission Extension Pod spacecraft. The NGHT-1X is a state of the art 1 kW-class Hall thruster which implements magnetic shielding to increase thruster lifetime by reducing discharge channel wall erosion. As a result, like other magnetically shielded Hall thrusters, the NGHT-1X exhibits erosion of the front-facing pole covers, which is the primary life-limiting failure mode of the thruster. During the development of the NGHT-1X, several Short Duration Wear Tests were performed to characterize erosion rates and forecast compliance to lifetime requirements in advance of the ongoing full-life Long Duration Wear Test. Erosion rate measurements from four tests at different operating conditions are presented and comparisons made between tests. Two tests were performed in the same background pressure environment at 700 W 300 V and 900 W 350 V throttle conditions, both using xenon propellant. A third test was performed again at 900 W 350 V with xenon propellant, but at a lower background pressure. Lastly, to assess the viability of potential future applications with alternate propellants, a wear test was conducted using krypton propellant at 900 W 250 V, the results of which are presented here. The erosion measurements from each test are compared with one another, and insights are discussed. All xenon wear tests support a thruster lifetime capability in excess of 16,000 hours, or 3.3 MNs at 900 W 350 V, which supports the Mission Extension Pod mission with margin as well as many other potential applications on small- to medium-sized spacecraft.

Hall↗
NASA NTRS1989

Low-speed wind tunnel performance of high-speed counterrotation propellers at angle-of-attack

The low-speed aerodynamic performance characteristics of two advanced counterrotation pusher-propeller configurations with cruise design Mach numbers of 0.72 were investigated in the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel. The tests were conducted at Mach number 0.20, which is representative of the aircraft take-off/landing flight regime. The investigation determined the effect of nonuniform inflow on the propeller performance characteristics for several blade angle settings and a range of rotational speeds. The inflow was varied by yawing the propeller model to angle-of-attack by as much as plus or minus 16 degrees and by installing on the counterrotation propeller test rig near the propeller rotors a model simulator of an aircraft engine support pylon and fuselage. The results of the investigation indicated that the low-speed performance of the counterrotation propeller configurations near the take-off target operating points were reasonable and were fairly insensitive to changes in model angle-of-attack without the aircraft pylon/fuselage simulators installed on the propeller test rig. When the aircraft pylon/fuselage simulators were installed, small changes in propeller performance were seen at zero angle-of-attack, but fairly large changes in total power coefficient and very large changes of aft-to-forward-rotor torque ratio were produced when the propeller model was taken to angle-of-attack. The propeller net efficiency, though, was fairly insensitive to any changes in the propeller flowfield conditions near the take-off target operating points.

Hughes, Christopher E.↗
NASA NTRS1961

Large-Scale Wind-Tunnel Tests of an Airplane Model with an Unswept, Tilt Wing of Aspect Ratio 5.5, and with Four Propellers and Blowing Flaps

Tests were made of a large-scale tilt-wing deflected-slipstream VTOL airplane with blowing-type BLC trailing-edge flaps. The model was tested with flap deflections of 0 deg. without BLC, 50 deg. with and without BLC, and 80 deg. with BLC for wing-tilt angles of 0, 30, and 50 deg. Included are results of tests of the model equipped with a leading-edge flap and the results of tests of the model in the presence of a ground plane.

Weiberg, James A.↗
NASA NTRS2016

Green Monopropellant Status at Marshall Space Flight Center

NASA Marshall Space Flight Center is continuing investigations into the use of green monopropellants as a replacement for hydrazine in spacecraft propulsion systems. Work to date has been to push technology development through multiple activities designed to understand the capabilities of these technologies. Future work will begin to transition to mission pull as these technologies are mature while still keeping a solid goal of pushing technology development as opportunities become available. The AF-M315E activities began with hot-fire demonstration testing of a 1N monopropellant thruster in FY 14 and FY15. Following successful completion of the preliminary campaign, changes to the test stand to accommodate propellant conditioning capability and better control of propellant operations was incorporated to make testing more streamlined. The goal is to conduct hot-fire testing with warm and cold propellants using the existing feed system and original thruster design. Following the 1N testing, a NASA owned 100 mN thruster will be hot-fire tested in the same facility to show feasibility of scaling to smaller thrusters for cubesat applications. The end goal is to conduct a hot-fire test of an integrated cubesat propulsion system using an SLM printed propellant tank, an MSFC designed propulsion system electronic controller and the 100 mN thruster. In addition to the AF-M315E testing, MSFC is pursuing hot-fire testing with LMP-103S. Following our successful hot-fire testing of the 22N thruster in April 2015, a test campaign was proposed for a 440N LMP-103S thruster with Orbital ATK and Plasma Processes. This activity was funded through the Space Technology Mission Directorate (STMD) ACO funding call in the last quarter of CY15. Under the same funding source a test activity with Busek and Glenn Research Center for testing of 5N AF-M315E thrusters was proposed and awarded. Both activities are in-work with expected completion of hot-fire testing by the end of FY17. MSFC is continuing to coordinate with the AF and academia on understanding the chemical reactions that occur in AF-M315E. An on-going investigation of the catalyst bed species using Raman Spectroscopy through the NASA Technology Research Fellowship Program (NSTRF) is looking for ways to minimize the amount of computation required by understanding the intermediate species created in the catalyst bed. The MSFC team is also working with commercial partners through Cooperative Agreement Notices (CAN's). Partnerships with commercial and academia include work in non-catalytic ignition of AF-M315, spark ignition of hybrid cubesat systems, printed SLM tanks, and dual-mode (electric and chemical) propulsion systems is continuing.

Burnside, Christopher G.↗
NASA NTRS1973

Slosh testing of a spherical mercury propellant tank with positive-expulsion diaphragm

A preliminary investigation was conducted to evaluate the lateral slosh characteristics of a 23-cm diameter mercury propellant tank with a positive-expulsion diaphragm and 17.5% ullage. Data are presented for tank sinusoidal acceleration levels between 0.05 and 0.5 g at frequencies ranging from 5 to 50 Hz. Results indicate that slosh characteristics are highly nonlinear at acceleration levels approaching those in flight, and depend heavily on the shape and stiffness of the diaphragm. Nyquist plots of driving force over acceleration are shown to be a useful tool for determining natural frequency, damping, and modal mass characteristics. A computerized nonlinear least squares method for extracting the modal parameters from the Nyquist plots is described and results of applying the method are presented.

Ross, R. G., Jr.↗