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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 127 records · Page 7

An evaluation of the role of eddy diffusion in stratospheric interactive two-dimensional models

An interactive two-dimensional model of the stratosphere, consisting of a primitive equation dynamics module, a simplified HO(x) ozone model, and a full radiative transfer scheme, is used to study the effect of eddy diffusion in the model. Consideration is given to the effects of nonlocal forcing from dissipation in the model troposphere and frictional drag at mesospheric levels, mechanical damping in the stratosphere itself, and potential vorticity flux due to large scale waves. It is found that the ozone distributions generated with the model are very sensitive to the choice of values for the friction and the eddy diffusion coefficients. It is shown that reasonable latitudinal gradients of ozone may be obtained by using small values for the mechanical damping for the mid- and high-latitude stratopsphere.

Schneider, Hans R.↗

Orientation of Space Station Freedom electrical power system in environmental effects assessment

The orientation effects of six Space Station Freedom Electrical Power System (EPS) components are evaluated for three environmental interactions: aerodynamic drag, atomic oxygen erosion, and orbital debris impact. Designers can directly apply these orientation factors to estimate the magnitude of the examined environment and the environmental effects for the EPS component of interest. The six EPS components are the solar array, photovoltaic module radiator, integrated equipment assembly, solar dynamic concentrator, solar dynamic radiator, and beta gimbal.

Lu, Cheng-Yi↗

A Computational Study of Plume Modeling For Space Launch System Abort Scenarios

This extended abstract details work on Space Launch System (SLS) core stage drag during abort scenarios that is currently being conducted. In the final paper, viscous computational fluid dynamic simulations will be used to study the axial force coefficient on the SLS core stage as a function of its distance to the upstream crew module and launch abort system, the freestream conditions, and the chemical model used to represent the launch abort system abort motor and attitude control motor plumes. In addition, the computational expense of each method will be studied in order to help determine the appropriate trade-off between fidelity and simulation turn around time.

CFD↗

Orion Entry Performance-Based Center-of-Gravity Box

The Orion capsule is designed both for Low Earth Orbit missions to the ISS and for missions to the moon. For ISS class missions, the capsule will use an Apollo-style direct entry. For lunar return missions, depending on the timing of the mission, the capsule could perform a direct entry or a skip entry of up to 4800 n.mi. in order to land in the coastal waters of California. The physics of atmospheric re-entry determine the capability of the Orion vehicle. For a given vehicle mass and shape, physics tells us that the driving parameters for an entry vehicle are the hypersonic lift-to-drag ratio (L/D) and the flight path angle at entry interface (gamma(sub EI)). The design of the Orion atmospheric re-entry must meet constraints during both nominal and dispersed flight conditions on landing accuracy, heating rate, total heat load, sensed acceleration, and proper disposal of the Service Module. These constraints define an entry corridor in the space of L/D-gamma(sub EI); if the vehicle falls within this corridor, then all constraints are met. The gamma(sub EI) dimension of the corridor can be further constrained by the gloads experienced during emergency entries. Thus, the entry performance for the Orion vehicle can be described completely by the L/D. Bounds on the hypersonic L/D necessary to achieve all the mission requirements can be defined for the given entry corridor. Landing accuracy performance drives the lower limit on L/D. In order to achieve the desired landing accuracy, a minimum L/D must be ensured. The design of the Thermal Protection System (TPS) drives the upper limit on L/D. A higher L/D can drive mass into the design of the TPS. Conversely, once the TPS is designed, the L/D must be ensured to stay below a certain limit in order for the TPS to stay within its design envelop. The L/D must stay within its upper and lower bounds during dispersed flight conditions. L/D is a function of both the aerodynamics and the center-of-gravity (CG) of the vehicle. The aerodynamics of the vehicle are determined by Computational Fluid Mechanics (CFD) and wind tunnel tests. However, the aerodynamics are not known precisely. Instead, an aerodynamic database has been developed where the aerodynamic coefficients are known to fall within a probabilistic band defined by upper and lower bounds. It is expected that the probabilistic band will shrink after the first missions are flown and real-world data is collected. Until that time, the Orion must be designed to the current aerodynamic database. Thus, for a given aerodynamic database with given uncertainties, the allowable range in L/D can be mapped to an allowable box for the CG location. The CG box is used to set requirements on the dispersions allowed for vehicle packaging and cargo storage. As the aerodynamic uncertainties decrease, the size of the CG box can increase. This paper discusses the technique used to map the minimum and maximum L/D bounds set by the entry performance requirements to the allowable dispersions in CG while accounting for aerodynamic uncertainties. The L/D is defined as the ratio of the lift force to the drag force. It is equivalent to the ratio of lift coefficient (C(sub L)) over drag coefficient (C(sub D)). C(sub L) and C(sub D) are functions of Mach number (M) and angle of attack (alpha). A Mach number of 25 is used as a measuring point of the hypersonic L/D. Variations in C(sub L), C(sub D) and alpha cause variations in L/D. Equation (1) shows the three contributions to the variation in L/D.

Rea, Jeremy R.↗

X-29 initial flight test results

It is announced that the X-29 forward-swept-wing (FSW) aircraft has been built, with flight testing under way and proceeding smoothly. The X-29 is a single-seat, single-engine supersonic aircraft that blends an optimized FSW, a close-coupled near-coplar canard, an F-5A forward fuselage module employing two side-mounted engine inlets, and a new aft fuselage. An F404-GE-400 engine with afterburner provides about 16,000 lb of thrust. The X-29 was designed to be near neutrally stable in the supersonic region to minimize drag; it becomes highly unstable transonically and subsonically, which dictates the use of a computerized fly-by-wire flight control system capable of stabilizing the aircraft. The X-29 advanced technology demonstrator began flight testing on Dec. 14, 1984; by Sept. 26, 1985, its envelope had been expanded to 0.75 Mach, 350 knots estimated airspeed, and 30,000 feet altitude. The aircraft has flown with very few problems and aerodynamic, structural, and control system results have correlated well with predictions. Areas where prediction and performance are not so well correlated are noted.

Ishmael, S. D.↗

Summary analysis of the Gemini entry aerodynamics

The aerodynamic data that were derived in 1967 from the analysis of flight-generated data for the Gemini entry module are presented. These data represent the aerodynamic characteristics exhibited by the vehicle during the entry portion of Gemini 2, 3, 5, 8, 10, 11, and 12 missions. For the Gemini, 5, 8, 10, 11, and 12 missions, the flight-generated lift-to-drag ratios and corresponding angles of attack are compared with the wind tunnel data. These comparisons show that the flight generated lift-to-drag ratios are consistently lower than were anticipated from the tunnel data. Numerous data uncertainties are cited that provide an insight into the problems that are related to an analysis of flight data developed from instrumentation systems, the primary functions of which are other than the evaluation of flight aerodynamic performance.

Whitnah, A. M.↗

Mapping the Martian Meteorology

The Mars-adapted version of the NASA/GISS general circulation model (GCM) has been applied to the hourly/daily simulation of the planet's meteorology over several seasonal orbits. The current running version of the model includes a diurnal solar cycle, CO2 sublimation, and a mature parameterization of upper level wave drag with a vertical domain extending from the surface up to the 6microb level. The benchmark simulations provide a four-dimensional archive for the comparative evaluation of various schemes for the retrieval of winds from anticipated polar orbiter measurements of temperatures by the Pressure Modulator Infrared Radiometer. Additional information is contained in the original extended abstract.

Allison, M.↗

Mapping the Martian Meteorology

The Mars-adapted version of the NASA/GISS general circulation model (GCM) has been applied to the hourly/daily simulation of the planet's meteorology over several seasonal orbits. The current running version of the model includes a diurnal solar cycle, CO2 sublimation, and a mature parameterization of upper level wave drag with a vertical domain extending from the surface up to the 6 micro b level. The benchmark simulations provide a four-dimensional archive for the comparative evaluation of various schemes for the retrieval of winds from anticipated polar orbiter measurements of temperatures by the Pressure Modulator Infrared Radiometer.

Allison, Michael↗

Progress Toward Efficient Laminar Flow Analysis and Design

A multi-fidelity system of computer codes for the analysis and design of vehicles having extensive areas of laminar flow is under development at the NASA Langley Research Center. The overall approach consists of the loose coupling of a flow solver, a transition prediction method and a design module using shell scripts, along with interface modules to prepare the input for each method. This approach allows the user to select the flow solver and transition prediction module, as well as run mode for each code, based on the fidelity most compatible with the problem and available resources. The design module can be any method that designs to a specified target pressure distribution. In addition to the interface modules, two new components have been developed: 1) an efficient, empirical transition prediction module (MATTC) that provides n-factor growth distributions without requiring boundary layer information; and 2) an automated target pressure generation code (ATPG) that develops a target pressure distribution that meets a variety of flow and geometry constraints. The ATPG code also includes empirical estimates of several drag components to allow the optimization of the target pressure distribution. The current system has been developed for the design of subsonic and transonic airfoils and wings, but may be extendable to other speed ranges and components. Several analysis and design examples are included to demonstrate the current capabilities of the system.

Campbell, Richard L.↗

Observations of disk-shaped bodies in free flight at terminal velocity

Ten disk-shaped models of a proposed nuclear heat source module were released from an aircraft and observed by radar. The initial launch attitude, spin rate, and mass of the models were varied. Significant differences were observed in the mode of flight and terminal velocity among models of different mass and launch attitudes. The data were analyzed to yield lift and drag coefficients as a function of Reynolds number. The total sea-level velocity of the models was found to be well correlated as a function of mass per unit frontal area. The demonstrated terminal velocity of the modular disk heat source, about 27 m/sec for this specific design, is only 33% of that of existing heat source designs.

Vorreiter, J. W.↗

Conjunction Assessment Past, Present, and Future

Since 1957, humankind's reliance on the space domain for military, humanitarian, and commercial applications has continued to increase. 1960 first successful use of a meteorological satellite,1963 first use of a geosynchronous communications satellite, 1985 Block I of GPS fielded, 1998 first module of ISS, 2001 first satellite radio broadcast over North America. What you take into space, stays in space: launch vehicle, rocket-bodies, mission-related debris. Debris can also be generated on-orbit: fuel/battery explosions, collisions. Only naturally-occurring retarding effect is orbital decay due to atmospheric drag: some remediation measures available, active debris removal not yet viable option. Because of our reliance on space and the fact that space really isn't limitless, the Big Sky theory is no longer an acceptable risk posture. There have been eight (8) on-orbit collisions reported to date, half of which occurred in the last 10 years.

Assessment↗

A feasibility study of a microgravity enhancement system for Space Station Freedom

The current low frequency microgravity requirements for Space Station Freedom (SSF) call for a level of less than 1 micro-g over 50 percent of all the laboratory racks for continuous periods of 30 days for at least 180 days per year. While this requirement is attainable for some of the laboratory modules for the Permanently Manned Configuration (PMC), it can not be met for the Man-Tended Configuration (MTC). In addition, many experiments would prefer even lower acceleration levels. To improve the microgravity environment, the Microgravity Enhancement System (MESYS) will apply a continuous thrust to SSF, to negate the disturbing gravity gradient and drag forces. The MESYS consists of a sensor, throttle-able thrusters and a control system. Both a proof mass system and accelerometer were evaluated for use as the sensor. The net result of the MESYS will be to shift the microgravity contours from the center of mass to a chosen location. Results indicate the MESYS is not feasible for MTC since it will require 5,073 kg of hydrazine fuel and 7,660 watts of power for 30 days of operation during average atmospheric conditions. For PMC, the MESYS is much more practical since only 4,008 kg of fuel and 5,640 watts of power are required.

Diamond, Preston S.↗

Does an orbiting star cause periodic modulation of X-rays from NGC6814?

An obvious candidate for the phenomenon underlying the periodicity in the X-ray emission from the Seyfert galaxy NGC6814 is the orbital motion of a star or low-mass compact object around the central black hole. It is shown here that the presence of an orbiting star could be easily verified by looking for the effects of Lense-Thirring precession of the orbital plane caused by the dragging of inertial frames around a rotating black hole. Precession-induced variations in the waveform and in the phase of the observed periodicity should have a period of between a month and a year. Such variations could account for the different waveforms present in the Ginga and Exosat data set from observations of NGC6814 and may be detectable in existing Ginga and future Rosat, OSSE/BRO, and Astro-D data.

Sikora, Marek↗

Pointing and tracking control for freedom's Solar Dynamic modules and vibration control of freedom

A control strategy is presented for pointing particular modules of flexible multibody space structures while simultaneously attenuating structural vibrations. The application that is addressed is the planned Space Station Freedom in a growth configuration with Solar Dynamic (SD) module. A NASTRAN model of Freedom is used to demonstrate the control strategy. Two cases of SD concentrator fine-pointing controller bandwidths are studied with examples. The effect of limiting the controller motor torques to realistic baseline values is examined. SD pointing and station vibration control is accomplished during realistic disturbances due to aerodynamic drag, Shuttle docking, and Shuttle reaction control system plume impingement on SD. Gravity gradient induced torques on SD are relatively small and pseudo-steady.

Quinn, Roger D.↗

Graphical Language for Data Processing

A graphical language for processing data allows processing elements to be connected with virtual wires that represent data flows between processing modules. The processing of complex data, such as lidar data, requires many different algorithms to be applied. The purpose of this innovation is to automate the processing of complex data, such as LIDAR, without the need for complex scripting and programming languages. The system consists of a set of user-interface components that allow the user to drag and drop various algorithmic and processing components onto a process graph. By working graphically, the user can completely visualize the process flow and create complex diagrams. This innovation supports the nesting of graphs, such that a graph can be included in another graph as a single step for processing. In addition to the user interface components, the system includes a set of .NET classes that represent the graph internally. These classes provide the internal system representation of the graphical user interface. The system includes a graph execution component that reads the internal representation of the graph (as described above) and executes that graph. The execution of the graph follows the interpreted model of execution in that each node is traversed and executed from the original internal representation. In addition, there are components that allow external code elements, such as algorithms, to be easily integrated into the system, thus making the system infinitely expandable.

Alphonso, Keith↗

An Approximate Analytical Method for Studying Atmosphere Entry of Vehicles with Modulated Aerodynamic Forces

The dimensionless, transformed, nonlinear differential equation developed in NASA TR R-11 for describing the approximate motion and heating during entry into planetary atmospheres for constant aerodynamic coefficients and vehicle shape has been modified to include entries during which the aerodynamic coefficients and the vehicle shape are varied. The generality of the application of the original equation to vehicles of arbitrary weight, size, and shape and to arbitrary atmospheres is retained. A closed-form solution for the motion, heating, and the variation of drag loading parameter m/C(D)A has been obtained for the case of constant maximum resultant deceleration during nonlifting entries. This solution requires certain simplifying assumptions which do not compromise the accuracy of the results. The closed-form solution has been used to determine the variation of m/C(D)A required to reduce peak decelerations and to broaden the corridor for nonlifting entry into the earth's atmosphere at escape velocity. The attendant heating penalty is also studied.

Levy, Lionel L., Jr.↗

CubeSub

This presentation introduces and discusses the development of the CubeSub submersible concept, an Autonomous Underwater Vehicle (AUV) designed around the CubeSat satellite form factor. The presented work is part of the author's MSc thesis in Aerospace Engineering at the Royal Institute of Technology, Stockholm, Sweden, and was performed during an internship at the Mission Design Division of the NASA Ames Research Center, Moffett Field, CA. Still in the early stages of its development, the CubeSub is to become a submersible test-bed for technology qualified for underwater and space environments. With the long-term goal of exploring the underwater environments in outer space, such as the alleged subsurface ocean of Jupiter's moon Europa, a number of technology and operational procedures must be developed and matured. To assist in this, the CubeSub platform is introduced as a tool to allow engineers and scientists to easily test qualified technology underwater. A CubeSat is a class of miniaturized satellite built to a standardized size. The base size is 1U (U for unit), corresponding to a 100 x 100 x 113.5 cu mm cube. A 1U CubeSat can in other words easily be held in one hand. Stacking units give larger satellite sizes such as the also commonly used 1.5U, 2U and 3U. The CubeSat standard is in itself already well established and hundreds of CubeSats have to date been launched into space. Compatible technology is readily available and the know-how exists in the space industry, all of which makes it a firm ground to stand on for the CubeSub. The rationale behind using the CubeSat form factor is to make use of this pre-existing foundation, making the CubeSub easy to develop, modular and readily available. It will thereby aid in the process of maturing the concept of a fully space qualified submersible headed for outer space. As a further clarification, the CubeSub is itself not meant for outer space, but to facilitate development of such a vessel. Along with its uses as a testbed, the CubeSub also holds the potential to become a useful tool for exploration and experimentation here on Earth. A highly standardized system utilizing well-known hardware can reduce the cost and required work load for researchers wishing to perform experiments and exploration. Users could design sensors and experiments to comply with the already well established CubeSat standard, which are then carried by the CubeSub to the region of interest. This in turn means that the end users can focus more on formulating the experiment itself and less about how to get it where they want it. The CubeSub is designed to be built up by modules, which can be assembled in different configurations to fulfill different needs. Each module will be powered individually and intermodular communication will be wireless, removing the need for wiring. The inside of the cylindrical hull will be flooded with ambient water to enhance the interaction between payloads and surrounding environment. The overall torpedo-like shape is similar to that of a conventional AUV, slender and smooth. This is to make for a low drag, reduce the risk of snagging on surrounding objects and make it possible to deploy through an ice sheet via a narrow borehole or navigate in tight areas. To keep costs low and further accelerate development, rapid prototyping is utilized wherever possible. Full-scale prototypes are being constructed through 3D-printing and using COTS (Commercial Off-The-Shelf) components. 3D-printing is used both for the largest hull components and the relatively small and delicate propellers. Arduino boards are used for control and internal communication

AUV↗

Saturn's E, G, and F rings - Modulated by the plasma sheet?

Saturn's broad E ring, the narrow G ring, and the structured and apparently time-variable F ring(s) contain many micron and submicron-sized particles, which make up the 'visible' component. These rings (or ring systems) are in direct contact with magnetospheric plasma. Fluctuations in the plasma density and/or mean energy, due to magnetospheric and solar wind processes, may induce stochastic charge variations on the dust particles, which in turn lead to an orbit perturbation and spatial diffusion. In addition, Coulomb drag forces may be important, in particular for the E ring. The possibility that electromagnetic effects may play a role in determining the F ring structure and its possible time variations is critically examined. Sputtering of micron-sized dust particles in the E ring by magnetospheric ions yields lifetimes of 100 to 10,000 years. This effect as well as the plasma induced transport processes require an active source for the E ring, probably Enceladus.

Morfill, G. E.↗