SEARCH · Search NASA
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Program cost and planning data based on 1980 technology and shown in 1979 dollars for a 20K lb Thrust Staged Combustion Cycle Engine are presented. These data were compared with those for the Advanced Expander Cycle Engine at 10K lb and 20K lb thrust levels.
Operations conducted from a space station in low Earth orbit which have different constraints and opportunities than those conducted from direct Earth launch were examined. While a space station relieves many size and performance constraints on the space shuttle, the space station's inertial orbit has different launch window constraints from those associated with customary Earth launches which reflect upon upper stage capability. A performance requirements analysis was developed to provide a reference source of parametric data, and specific case solutions and upper stage sizing trade to assist potential space station users and space station and upper stage developers assess the impacts of a space station on missions of interest.
The purpose of our experiments was to estimate basic sensitivity to motion gradients and to evaluate the evidence for second-order integration and differentiation of motion signals. We measured sensitivity to spatially sinusoidal contrast modulation between two oppositely moving bandpass-filtered noise images. The motion-contrast sensitivity function, defined as the inverse of threshold modulation amplitude as a function of modulation spatial frequency, was bandpass in shape with declines at both highest and lowest frequencies. The functions for three noise spatial frequencies had approximately the same shape when modulation frequency was expressed as a fraction of noise frequency. We compared the data with a model in which linear motion filters, whose outputs are squared or rectified, are followed by a second stage of excitatory or inhibitory pooling. The data are consistent with a model in which (1) all excitatory pooling occurs at the linear stage and (2) the second stage contains a large inhibitory pooling area, with a radius approximately eight times that of the linear receptive field.
The space transportation system (STS) is discussed which will include a propulsive stage that is carried into low earth orbit by the space shuttle. Data were accumulated from the analyses of various stage concepts, operating modes, and projected missions (space tug systems studies, growth stage studies, engine studies, and critical area studies). The foundation formulated by these studies aided in establishing a tentative two-phase approach for the extension of the STS operating regime beyond the space shuttle including plane changes, higher orbits, geosynchronous orbits and beyond. The orbital operations study, was conducted to provide the generation and analysis of operational plans, requirement, and concepts for the utilization of these vehicles. Primary emphasis was placed on methods and techniques to provide a sound technical approach to operations at reduced cost for the STS mission period.
This report is the documentation of the work performed under the Hypersonic Project of the NASA's Fundamental Aeronautics Program. It was funded through Task Number NNC10E444T under GESS-2 Contract NNC06BA07B. The objective of the task is to develop advanced computational tools for the simulation of multi-stage turbomachinery in support of aeropropulsion. This includes work elements in extending the TURBO code and validating the multi-stage URANS (Unsteady Reynolds Averaged Navier Stokes) simulation results with the experimental data. The unsteady CFD (Computation Fluid Dynamics) calculations were performed in full wheel mode with and without screen generated total pressure distortion at the computational inflow boundary, as well as in single passage phase lag mode for uniform inflow. The experimental data were provided by NASA from the single stage RTA (Revolutionary Turbine Accelerator) fan test program.Significant non-uniform flow condition at the fan-face of the aeropropulsion system is frequentlyencountered in many of the advanced aerospace vehicles. These propulsion systems can be eithera podded or an embedded design employed in HWB (Hybrid Wing Body) airframe concept. It isalso a topic of interest in military applications, in which advanced air vehicles have already deployedsome form of embedded propulsion systems in their design because of the requirementsof compact and low observable inlets. Even in the conventional airframe/engine design, the fancould operate under such condition when the air vehicle is undergoing rapid maneuvering action.It is believed that a better understanding of the fan’s aerodynamic and aeromechanical responseto this type of operating condition or off design operation would be beneficial to designing distortiontolerant blades for improved engine operability.The objective for this research is to assess the capability of turbomachinery code as an analysistool in understanding the effects and evaluating the impact of flow distortion on the aerodynamicand aeromechanical performance of the fan in advanced propulsion systems. Results from thetesting of an advanced fan stage released by NASA are available and will be used here for CFDcode validation. The experiment was performed at NASA’s high speed compressor facility aspart of the RTA (Revolutionary Turbine Accelerator) demonstration project, a joint effort ofNASA Glenn Research Center and GE Aircraft Engines in developing an advanced Mach 4TBCC (Turbine Based Combined Cycle) turbofan/ramjet engine for access to space. Part of thetest was to assess the aerodynamic performance and operability of the fan stage under nonuniforminflow condition. Various flow distortion patterns were created at the fan-face by manipulatingsets of screens placed upstream of the wind tunnel. Measurements at the fan-face willprovide the necessary distortion flow information as the inflow boundary condition for the CFDin a full wheel simulation. Therefore the purpose of this work is to demonstrate the NASA supportedmulti-stage turbomachinery code, TURBO [1-5], in the aerodynamic performance analysisof a modern fan design operating under off design condition, and in particular to validate theCFD results with the RTA fan test data.A brief description of the RTA fan rig configuration is given in the next section, explaining onhow flow distortion were measured in the test and constructed for the CFD at the fan-face. It isfollowed by a section summarizing previous CFD work performed at NASA relevant to the currentfan configuration. A short description of the TURBO code is given next, followed by detailsin the computational model of the fan rig, the required computing resources, and the numericalprocedure for the simulations. The CFD results are presented in the discussion section and finallyconcluding remarks are summarized.
Considerable data on various parameters of a satellite and its launching vehicle can be obtained from the received-signal-strength records at the various satellite receiver stations. The doppler shifts of the satellite can be used to determine the orbital characteristics quite accurately. (Ref. 1 discusses the rapid determination of these parameters.) Information concerning the velocity increment and direction of each stage can be obtained if doppler data are received by several stations during the high-speed staging. The variations in signal strength during launch can be used to find the spin rate of the satellite before, during, and after staging, and to find variations in attitude, including precession, during launch. With spin and precession frequencies available, any changes in the ratios of moments of inertia can be determined. In this paper the launchings of the Explorer satellites will be used as a basis for the generalized conclusions concerning utilization of radio-signal-strength records. The discussion includes a description of the Explorer launching system.
This lecture note discusses the role of computer modelling in the design, development, and validation of a performance monitoring system. The basic requirements of an engine health monitoring system are discussed in the context of the user environment. A form of model based on stage characteristics provides the basis for describing engine measurements. Faults are modelled in accordance with empirical data obtained from tests conducted at the stage level. The model is used to investigate various parameters that provide unambiguous identification of the fault in question. Fault libraries have been developed for field use.
The advent of the new Ares I launch vehicle has highlighted the need for advanced dynamic analysis tools for variable mass, flexible structures. This system is composed of interconnected flexible stages or components undergoing rapid mass depletion through the consumption of solid or liquid propellant. In addition to large rigid body configuration changes, the system simultaneously experiences elastic deformations. In most applications, the elastic deformations are compatible with linear strain-displacement relationships and are typically modeled using the assumed modes technique. The deformation of the system is approximated through the linear combination of the products of spatial shape functions and generalized time coordinates. Spatial shape functions are traditionally composed of normal mode shapes of the system or even constraint modes and static deformations derived from finite element models of the system. Equations of motion for systems undergoing coupled large rigid body motion and elastic deformation have previously been derived through a number of techniques [1]. However, in these derivations, the mode shapes or spatial shape functions of the system components were considered constant. But with the Ares I vehicle, the structural characteristics of the system are changing with the mass of the system. Previous approaches to solving this problem involve periodic updates to the spatial shape functions or interpolation between shape functions based on system mass or elapsed mission time. These solutions often introduce misleading or even unstable numerical transients into the system. Plus, interpolation on a shape function is not intuitive. This paper presents an approach in which the shape functions are held constant and operate on the changing mass and stiffness matrices of the vehicle components. Each vehicle stage or component finite element model is broken into dry structure and propellant models. A library of propellant models is used to describe the distribution of mass in the fuel tank or Solid Rocket Booster (SRB) case for various propellant levels. Based on the mass consumed by the liquid engine or SRB, the appropriate propellant model is coupled with the dry structure model for the stage. Then using vehicle configuration data, the integrated vehicle model is assembled and operated on by the constant system shape functions. The system mode shapes and frequencies can then be computed from the resulting generalized mass and stiffness matrices for that mass configuration. The rigid body mass properties of the vehicle are derived from the integrated vehicle model. The coupling terms between the vehicle rigid body motion and elastic deformation are also updated from the constant system shape functions and the integrated vehicle model. This approach was first used to analyze variable mass spinning beams and then prototyped into a generic dynamics simulation engine. The resulting code was tested against Crew Launch Vehicle (CLV-)class problems worked in the TREETOPS simulation package and by Wilson [2]. The Ares I System Integration Laboratory (SIL) is currently being developed at the Marshall Space Flight Center (MSFC) to test vehicle avionics hardware and software in a hardware-in-the-loop (HWIL) environment and certify that the integrated system is prepared for flight. The Ares I SIL utilizes the Ares Real-Time Environment for Modeling, Integration, and Simulation (ARTEMIS) tool to simulate the launch vehicle and stimulate avionics hardware. Due to the presence of vehicle control system filters and the thrust oscillation suppression system, which are tuned to the structural characteristics of the vehicle, ARTEMIS must incorporate accurate structural models of the Ares I launch vehicle. The ARTEMIS core dynamics simulation models the highly coupled nature of the vehicle flexible body dynamics, propellant slosh, and vehicle nozzle inertia effects combined with mass and flexible body properties that vary significant with time during the flight. All forces that act on the vehicle during flight must be simulated, including deflected engine thrust force, spatially distributed aerodynamic forces, gravity, and reaction control jet thrust forces. These forces are used to excite an integrated flexible vehicle, slosh, and nozzle dynamics model for the vehicle stack that simulates large rigid body translations and rotations along with small elastic deformations. Highly effective matrix math operations on a distributed, threaded high-performance simulation node allow ARTEMIS to retain up to 30 modes of flex for real-time simulation. Stage elements that separate from the stack during flight are propagated as independent rigid six degrees of freedom (6DOF) bodies. This paper will present the formulation of the resulting equations of motion, solutions to example problems, and describe the resulting dynamics simulation engine within ARTEMIS.
The original design analysis and shop test data are presented for a three stage (poster) air compressor with impellers mounted on the extensions of a twin pinion gear, and driven by an 8000 hp synchronous motor. Also included are field test data, subsequent rotor dynamics analysis, modifications, and final rotor behavior. A subsynchronous instability existed on a geared, overhung rotor. State-of-the-art rotor dynamics analysis techniques provided a reasonable analytical model of the rotor. A bearing modification arrived at analytically eliminated the instability.
The National Bureau of Standards (NBS) Data Management Technology Program is discussed. The NBS Data Management Technology Program addresses major problems encountered during the following stages of an application's lifetime: requirements analysis and data base design, system selection and implementation, operations management and conversion. Products developed include standard software specifications, guides to best practice, standard data elements and representations, and reports documenting the experiences of other organizations as they attempt to improve the management of their computing resources. Data base Laboratory facilities are maintained for the investigation and analysis of state of the art data base technology. These facilities support collaborative testing with researchers, vendors, users, and standards developers.
Data processor with plural register stages for selectively interconnecting with each other to effect multiplicity of operations
Multispectral data processing procedures are outlined beginning with the data display process used to accomplish data editing and proceeding through clustering, feature selection criterion for error probability estimation, and sample clustering and sample classification. The effective utilization of large quantities of remote sensing data by formulating a three stage sampling model for evaluation of crop acreage estimates represents an improvement in determining the cost benefit relationship associated with remote sensing technology.
Tradeoffs between aerodynamic and acoustic properties of various fan configurations led to the selection of a variety of fans that produce noise levels in the range of 100 to 120 PNdb. Fan configuration and design pressure ratio required for specific mission operations depend in part on the type of mission to be performed. Noise data obtained for single-stage low speed fans, single-stage high speed fans, and two-stage fan engines are presented in table form.
On Apr. 15, 1970, the Apollo 13 S-IVB stage impacted the nighttime lunar surface. Beginning 20 sec after impact, the Suprathermal Ion Detector Experiment and the Solar Wind Spectrometer observed a large flux of positive ions (maximum flux of about 3 x 10 to the 8th ions/sq cm/sec/ster) and electrons. Two separate streams of ions were observed: a horizontal flux that appeared to be deflected solar wind ions and a smaller vertical flux of predominantly heavy ions (greater than 10 amu), which probably were material vaporized from the S-IVB stage. An examination of the data shows that collisions between neutral molecules and hot electrons (50 eV) were probably an important ionization mechanism in the impact-produced neutral gas cloud. These electrons, which were detected by the Solar Wind Spectrometer, are thought to have been energized in a shock front or some form of intense interaction region between the cloud and the solar wind. Thus strong ionization and acceleration are seen under conditions approaching a collisionless state.
Available experimental data comparing the in-flight and static behavior of fan noise are reviewed. These results are then compared with recent data obtained for a fan stage tested with forward velocity in a low speed wind tunnel. Tentative conclusions are presented about the significance and nature of the changes in noise observed when a forward velocity is imposed. Finally, the implications of the emerging picture of in-flight fan source noise for suppressor design are discussed.
The available experimental data comparing the in-flight and static behavior of fan noise are reviewed. These results are then compared with recent data obtained for a fan stage tested with forward velocity in the NASA Lewis 9x15 low speed wind tunnel. Tentative conclusions are presented, based on the author's judgments, about the significance and nature of the changes in noise observed when a forward velocity is imposed. Finally, the implications of the emerging picture of in-flight fan source noise for suppressor design are discussed.
The author has identified the following significant results. Recognition of rice varieties at the flowering stage by using airborne scanner data at low altitude (1500 m) seems to be feasible. The accuracies obtained on a reduced test area (3 sq km) range from 65% to 83%. Variations of a single cultural factor, such as nitrogen fertilization, induce variations of the total rice biomass at harvest, which can be correlated closely to the values of the reflectance ratio at earing. When grain production is correlated to total biomass, prediction of yield can be achieved based on reflectance data measured two months before harvest.