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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 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS2009

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↗
NASA NTRS1997

Evolutionary Design of Controlled Structures

Basic physical concepts of structural delay and transmissibility are provided for simple rod and beam structures. Investigations show the sensitivity of these concepts to differing controlled-structures variables, and to rational system modeling effects. An evolutionary controls/structures design method is developed. The basis of the method is an accurate model formulation for dynamic compensator optimization and Genetic Algorithm based updating of sensor/actuator placement and structural attributes. One and three dimensional examples from the literature are used to validate the method. Frequency domain interpretation of these controlled structure systems provide physical insight as to how the objective is optimized and consequently what is important in the objective. Several disturbance rejection type controls-structures systems are optimized for a stellar interferometer spacecraft application. The interferometric designs include closed loop tracking optics. Designs are generated for differing structural aspect ratios, differing disturbance attributes, and differing sensor selections. Physical limitations in achieving performance are given in terms of average system transfer function gains and system phase loss. A spacecraft-like optical interferometry system is investigated experimentally over several different optimized controlled structures configurations. Configurations represent common and not-so-common approaches to mitigating pathlength errors induced by disturbances of two different spectra. Results show that an optimized controlled structure for low frequency broadband disturbances achieves modest performance gains over a mass equivalent regular structure, while an optimized structure for high frequency narrow band disturbances is four times better in terms of root-mean-square pathlength. These results are predictable given the nature of the physical system and the optimization design variables. Fundamental limits on controlled performance are discussed based on the measured and fit average system transfer function gains and system phase loss.

Masters, Brett P.↗
NASA NTRS2020

Pterodactyl: Thermal Protection System for Integrated Control Design of a Mechanically Deployed Entry Vehicle

The need for precision landing of high mass payloads on Mars and the return of sensitive samples from other planetary bodies to specific locations on Earth is driving the development of an innovative NASA technology referred to as the Deployable Entry Vehicle (DEV). A DEV has the potential to deliver an equivalent science payload with a stowed diameter 3 to 4 times smaller than a traditional rigid capsule configuration. However, the DEV design does not easily lend itself to traditional methods of directional control. The NASA Space Technology Mission Directorate (STMD)’s Pterodactyl project is currently investigating the effectiveness of three different Guidance and Control (G&C) systems – actuated flaps, Center of Gravity (CG) or mass movement, and Reaction Control System (RCS) – for use with a DEV using the Adaptable, Deployable, Entry, and Placement Technology (ADEPT) design. This paper details the Thermal Protection System (TPS) design and associated mass estimation efforts for each of the G&C systems. TPS is needed for the nose cap of the DEV and the flaps of the actuated flap control system. The development of a TPS selection, sizing, and mass estimation method designed to deal with the varying requirements for the G&C options throughout the trajectory is presented. The paper discusses the methods used to i) obtain heating environments throughout the trajectory with respect to the chosen control system and resulting geometry; ii) determine a suitable TPS material; iii) produce TPS thickness estimations; and, iv) determine the final TPS mass estimation based on TPS thickness, vehicle control system, vehicle structure, and vehicle payload.

Hays, Zane B.↗
NASA NTRS2011

Propulsion Physics Using the Chameleon Density Model

To grow as a space faring race, future spaceflight systems will require a new theory of propulsion. Specifically one that does not require mass ejection without limiting the high thrust necessary to accelerate within or beyond our solar system and return within a normal work period or lifetime. The Chameleon Density Model (CDM) is one such model that could provide new paths in propulsion toward this end. The CDM is based on Chameleon Cosmology a dark matter theory; introduced by Khrouy and Weltman in 2004. Chameleon as it is hidden within known physics, where the Chameleon field represents a scalar field within and about an object; even in the vacuum. The CDM relates to density changes in the Chameleon field, where the density changes are related to matter accelerations within and about an object. These density changes in turn change how an object couples to its environment. Whereby, thrust is achieved by causing a differential in the environmental coupling about an object. As a demonstration to show that the CDM fits within known propulsion physics, this paper uses the model to estimate the thrust from a solid rocket motor. Under the CDM, a solid rocket constitutes a two body system, i.e., the changing density of the rocket and the changing density in the nozzle arising from the accelerated mass. Whereby, the interactions between these systems cause a differential coupling to the local gravity environment of the earth. It is shown that the resulting differential in coupling produces a calculated value for the thrust near equivalent to the conventional thrust model used in Sutton and Ross, Rocket Propulsion Elements. Even though imbedded in the equations are the Universe energy scale factor, the reduced Planck mass and the Planck length, which relates the large Universe scale to the subatomic scale.

Robertson, Glen A.↗
NASA NTRS2019

Development of Pericyclic Gearbox for Roving Application

Many aerospace mechanisms require high gear reduction ratios to reduce the high output speed of a motor to much lower speeds for precision rotation. Conventionally this actuation is performed with multiple gear stages, typically including a harmonic gearbox, and is required to be lightweight for spaceflight. One application of multiple gear stages is in rover wheel actuators intended for use on the Moon and Mars. These actuators require low total mass, high reliability, and precision actuation due to their remote operation. The pericyclic gearbox has the capability in a single stage to provide high reduction ratios equivalent to multiple conventional stages. Some benefits of utilizing pericyclic gear technology over multiple conventional stages are a compact form, fewer total parts, no reliance on flexible members (such as in harmonic drives), and lower total system mass. The goal of this work was to modify and build on existing modeling tools to determine the feasibility of utilizing the pericyclic gearbox for such an application. Critical aspects to consider were gear tooth loads, gear mesh efficiency, bearing loads, bearing efficiency, and total system mass. Developed tools were then used to examine a single stage pericyclic gearbox's trends in mass and efficiency for a Mars 2020 rover wheel actuator mechanism design case which was then compared to multistage gearbox designs.

Krantz, Timothy↗
NASA NTRS1977

Alternate approaches to vibration and shock analysis using NASTRAN

A method that derives an approximate equivalent static load to a base excitation shock analysis is described. The transient analysis in the current level of NASTRAN, level 16, does not directly provide for either input acceleration forcing functions or enforced boundary displacement. In the suggested alternate analysis format, equivalent force input functions are applied to the constrained locations by using the artifice of placing a large mass, with respect to the total system mass, at the desired acceleration input points. This shortcut static analysis approach is presented to approximate the expensive and time consuming dynamics analysis approach to the base excitation shock analysis.

Denver, R. E.↗
NASA NTRS1995

A study of the radiation environment on board the space shuttle flight STS-57

A joint NASA-Russian study of the radiation environment inside a SPACEHAB 2 locker on space shuttle flight STS-57 was conducted. The shuttle flew in a nearly circular orbit of 28.5 deg inclination and 462 km altitude. The locker carried a charged particle spectrometer, a tissue equivalent proportional counter (TEPC), and two area passive detectors consisting of combined NASA plastic nuclear track detectors (PNTD's) and thermoluminescent detectors (TLD's), and Russian nuclear emulsions, PNTD's, and TLD's. All the detector systems were shielded by the same shuttle mass distribution. This makes possible a direct comparison of the various dose measurement techniques. In addition, measurements of the neutron energy spectrum were made using the proton recoil technique. The results show good agreement between the integral LET spectrum of the combined galactic and trapped particles using the tissue equivalent proportional counter and track detectors between about 15 keV/micron and 200 keV/micron. The LET spectrum determined from nuclear emulsions was systematically lower by about 50%, possibly due to emulsion fading. The results show that the TEPC measured an absorbed dose 20% higher than TLD's, due primarily to an increased TEPC response to neutrons and a low sensitivity of TLD's to high LET particles under normal processing techniques. There is a significant flux of high energy neutrons that is currently not taken into consideration in dose equivalent calculations. The results of the analysis of the spectrometer data will be reported separately.

Badhwar, G. D.↗
NASA NTRS1979

Determination of muscle mass changes in legs from K-40 measurements

The K-40 content of the upper legs was periodically measured in several subjects whose injured legs had been in a cast for 6 weeks or more. As the subjects began using the leg again, the K-40 content increased as the muscle tissue was replaced. A 25% increase in K-40 content in 6 months is typical for a normal leg use and recovery. This is equivalent to an original muscle mass loss of 20%. By measuring specific body regions, such as arms or legs, with a high-efficiency detector system, muscle mass changes which exceed a few percent can be measured. These methods could be used in space flight and bedrest studies, and in studying nutritional deficiencies due to disease or diet.

Palmer, H. E.↗
NASA NTRS2010

A Study of Parallels Between Antarctica South Pole Traverse Equipment and Lunar/Mars Surface Systems

The parallels between an actual Antarctica South Pole re-supply traverse conducted by the National Science Foundation (NSF) Office of Polar Programs in 2009 have been studied with respect to the latest mission architecture concepts being generated by the United States National Aeronautics and Space Administration (NASA) for lunar and Mars surface systems scenarios. The challenges faced by both endeavors are similar since they must both deliver equipment and supplies to support operations in an extreme environment with little margin for error in order to be successful. By carefully and closely monitoring the manifesting and operational support equipment lists which will enable this South Pole traverse, functional areas have been identified. The equipment required to support these functions will be listed with relevant properties such as mass, volume, spare parts and maintenance schedules. This equipment will be compared to space systems currently in use and projected to be required to support equivalent and parallel functions in Lunar and Mars missions in order to provide a level of realistic benchmarking. Space operations have historically required significant amounts of support equipment and tools to operate and maintain the space systems that are the primary focus of the mission. By gaining insight and expertise in Antarctic South Pole traverses, space missions can use the experience gained over the last half century of Antarctic operations in order to design for operations, maintenance, dual use, robustness and safety which will result in a more cost effective, user friendly, and lower risk surface system on the Moon and Mars. It is anticipated that the U.S Antarctic Program (USAP) will also realize benefits for this interaction with NASA in at least two areas: an understanding of how NASA plans and carries out its missions and possible improved efficiency through factors such as weight savings, alternative technologies, or modifications in training and operations.

Mueller, Robert P.↗
NASA NTRSDate not supplied

Enabling New and Innovative Low Cost Mars Science Missions with the Adaptable, Deployable, Entry and Placement Technology (ADEPT)

The Adaptive Deployable Entry and Placement Technology (ADEPT) offers a delivery capability for Small Sat or CubeSat orbiter(s), in-situ elements, or landers. ADEPT can deliver the same science payload to a destination with a stowed diameter a factor of 3-4 times smaller than an equivalent rigid aeroshell, alleviating volumetric constraints on the secondary payload accommodation or primary carrier spacecraft bus. The ADEPT system can support payload over a range of mass and geometries to offer a delivery capability for a single or constellation of spacecraft supporting the goals and objectives of “Low Cost Mars Science Missions.”

ADEPT↗
NASA NTRS1988

Nonlinear modeling and estimation of slew induced structural deformations

A model of the nonlinear dynamics of a deformable maneuvering multibody system is described, whereby elastic deformation are modeled by restoring forces and dissipative forces at point mass appendages. This model is brought into bilinear form. Estimation of deformations occasioned by rapid slewing maneuvers is carried out by a filter based on a globally equivalent linear model of the bilinear dynamics, and is shown to be an improvement over the extended Kalman filter. To further alleviate the computational burden, the estimated deformation state is propagated between observations by a low dimensional operator spline interpolator of bilinear system Volterra series, which is easily implemented.

Dwyer, T. A. W., III↗
NASA NTRS1991

Bilinear modeling and state estimation of slew-induced deformations of a flexible multibody system

A model of the nonlinear dynamics of a deformable maneuvering multibody system is described, whereby elastic deformations are modeled by restoring forces and dissipative forces at point mass appendages. This model is brought into bilinear form. Estimation of deformations occasioned by rapid slewing maneuvers is carried out by a filter based on a globally equivalent linear model of the bilinear dynamics, and is shown to be an improvement over the extended Kalman filter. To further alleviate the computational burden, the estimated deformation state is propagated between observations by a low dimensional operator spline interpolator of bilinear system Volterra series, which is easily implemented.

Dwyer, Thomas A. W., III↗
NASA NTRS2006

Semiautonomous Avionics-and-Sensors System for a UAV

Unmanned Aerial Vehicles (UAVs) autonomous or remotely controlled pilotless aircraft have been recently thrust into the spotlight for military applications, for homeland security, and as test beds for research. In addition to these functions, there are many space applications in which lightweight, inexpensive, small UAVS can be used e.g., to determine the chemical composition and other qualities of the atmospheres of remote planets. Moreover, on Earth, such UAVs can be used to obtain information about weather in various regions; in particular, they can be used to analyze wide-band acoustic signals to aid in determining the complex dynamics of movement of hurricanes. The Advanced Sensors and Electronics group at Langley Research Center has developed an inexpensive, small, integrated avionics-and-sensors system to be installed in a UAV that serves two purposes. The first purpose is to provide flight data to an AI (Artificial Intelligence) controller as part of an autonomous flight-control system. The second purpose is to store data from a subsystem of distributed MEMS (microelectromechanical systems) sensors. Examples of these MEMS sensors include humidity, temperature, and acoustic sensors, plus chemical sensors for detecting various vapors and other gases in the environment. The critical sensors used for flight control are a differential- pressure sensor that is part of an apparatus for determining airspeed, an absolute-pressure sensor for determining altitude, three orthogonal accelerometers for determining tilt and acceleration, and three orthogonal angular-rate detectors (gyroscopes). By using these eight sensors, it is possible to determine the orientation, height, speed, and rates of roll, pitch, and yaw of the UAV. This avionics-and-sensors system is shown in the figure. During the last few years, there has been rapid growth and advancement in the technological disciplines of MEMS, of onboard artificial-intelligence systems, and of smaller, faster, and smarter wireless telemetry systems. The major attraction of MEMS lies in orders-of-magnitude reductions of power requirements relative to traditional electronic components that perform equivalent functions. In addition, the compactness of MEMS, relative to functionally equivalent traditional electronics systems, makes MEMS attractive for UAV applications. Recent advances in MEMS have made it possible to produce pressure, acceleration, humidity, and temperature sensors having masses in subgram range and possessing sensitivities and accuracies comparable to those of larger devices.

Shams, Qamar↗
NASA NTRS1994

Comparison of Mixing Calculations for Reacting and Non-Reacting Flows in a Cylindrical Duct

A production 3-D elliptic flow code has been used to calculate non-reacting and reacting flow fields in an experimental mixing section relevant to a rich burn/quick mix/lean burn (RQL) combustion system. A number of test cases have been run to assess the effects of the variation in the number of orifices, mass flow ratio, and rich-zone equivalence ratio on the flow field and mixing rates. The calculated normalized temperature profiles for the non-reacting flow field agree qualitatively well with the normalized conserved variable isopleths for the reacting flow field indicating that non-reacting mixing experiments are appropriate for screening and ranking potential rapid mixing concepts. For a given set of jet momentum-flux ratio, mass flow ratio, and density ratio (J, MR, and DR), the reacting flow calculations show a reduced level of mixing compared to the non-reacting cases. In addition, the rich-zone equivalence ratio has noticeable effect on the mixing flow characteristics for reacting flows.

Oechsle, V. L.↗
NASA NTRS1994

Discovery of a new white dwarf in a binary system (EUVE 0720-317) in the Extreme Ultraviolet Explorer survey and implications for the late stages of stellar evolution

A new precataclysmic binary is identified in the Extreme Ultraviolet Explorer (EUVE) all-sky survey. The bright source EUVE 0720-317 shows a hot hydrogen-rich white dwarf optical continuum with overlying narrow Balmer-line emission. Using high signal-to-noise ratio spectroscopy in the 4100-6700 A range, we identify a late-type companion and find a 1.3d periodic modulation in the emission-line velocities and strengths. This is the signature of Feige 24-type close binary systems. We determine the components' spectral types (DAO and dM0-2), orbital velocities (K(sub DAO) = 104 +/- 12 km/s, K(sub dM) = 96 +/- 7 km/s), and systemic velocity (gamma = 15 +/- 12 km/s). A first estimate of the white dwarf gravitational redshift, gamma(sub g) = 45 +/- 20 km/s, and theoretical mass-radius relationships imply R(sub DAO) = 0.010-0.016 solar radius and M(sub DAO) = 0.55-0.90 solar mass. The orbital inclination is therefore i greater than or equal to 52 deg, consistent with the large amplitude variations found in H-alpha equivalent widths that imply i greater than or equal to 42 deg. We show that the discovery of new close WD + MS binary systems in extreme ultraviolet (EUV) sky surveys has important implications for theory of common-envelope evolution, in particular for the predicted close binary birthrate and orbital and stellar parameters.

Vennes, Stephane↗
NASA NTRS2008

Fuel Cell Thermal Management Through Conductive Cooling Plates

An analysis was performed to evaluate the concept of utilizing conductive cooling plates to remove heat from a fuel cell stack, as opposed to a conventional internal cooling loop. The potential advantages of this type of cooling system are reduced stack complexity and weight and increased reliability through the reduction of the number of internal fluid seals. The conductive cooling plates would extract heat from the stack transferring it to an external coolant loop. The analysis was performed to determine the required thickness of these plates. The analysis was based on an energy balance between the thermal energy produced within the stack and the heat removal from the cooling plates. To accomplish the energy balance, the heat flow into and along the plates to the cooling fluid was modeled. Results were generated for various numbers of cells being cooled by a single cooling plate. The results provided cooling plate thickness, mass, and operating temperature of the plates. It was determined that utilizing high-conductivity pyrolitic graphite cooling plates can provide a specific cooling capacity (W/kg) equivalent to or potentially greater than a conventional internal cooling loop system.

Colozza, Anthony J.↗
NASA NTRS1980

Thermal-structural design study of an airframe-integrated Scramjet

The development and evaluation of a design concept for the cooled structures assembly for the Scramjet engine is discussed. Development concepts for engine subsystems and design concepts for the aircraft/engine interface are included. A thermal protection system was defined which makes it possible to attain a life of 100 hr and 1000 cycles, the specified goal. The coolant equivalence ratio at the Mach 10 maximum thermal loading condition is 0.6, indicating a capacity for airframe cooling. The mechanical design is feasible for manufacture using conventional materials. For the cooled structures in a six module engine, the mass per unit capture area is 1256 kg/sq m. The total mass of a six module engine assembly including the fuel system is 1502 kg.

Killackey, J. J.↗