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Training data for both dipole dipole and Alt3 wenner configurations included. Field dataset for both dipole dipole and Alt3 wenner configurations included (Test dataset). Python code for implementation of MLP-NN included. Results for inversion of both configurations included.
The impurity screening response of the high-field side (HFS) scrape-off layer (SOL) to localized nitrogen injection is investigated on Alcator C-Mod for magnetic equilibria spanning lower-single null, double-null and upper single null configurations under otherwise identical plasma conditions. L-mode, EDA H-mode and I-mode discharges are investigated. HFS impurity screening is found to depend on magnetic flux balance and the direction of B×?B relative to the most active divertor. Impurity ‘plume’ emission patterns indicate that both parallel and perpendicular (E×B) flows in the SOL contribute to the ‘flushing’ of impurities towards the active divertor, thereby affecting the overall impurity screening behavior. Despite the fact that the HFS scrape-off layer is extremely narrow in near-double-null configurations, this SOL is able to screen locally injected nitrogen at least as effectively as the low-field side (LFS) SOL – up to a factor of 10 more effective, depending on specific plasma conditions and whether the magnetic geometry produces parallel flows that work with or against E×B flows. For situations in which the E×B drift of the impurity ions opposes parallel flow toward the primary divertor, HFS impurity screening is found to be least effective. When E×B drifts assist parallel flow toward the primary divertor, HFS impurity screening is found to be very effective. These data support the idea of placing RF actuators and close-fitting wall components on the high-field side of the tokamak. With this configuration, near-double-null magnetic topologies may be used for active control of plasma parameters at the antenna/plasma interface for optimal RF coupling, mitigate the generation of local impurities from plasma-material interactions and, taking advantage of favorable plasma flows and good screening properties of the HFS SOL, further minimize the impact of wall-born impurity sources on the plasma core.
Over the years aerodynamicists have learned to rely heavily on wind-tunnel-model results in predicting the aerodynamic characteristics of conventional aircraft configurations. With the development of V/STOL configurations which have high slipstream deflection angles. such as the four-propeller tilt-wing XC-142A V/STOL aircraft, the reliability of small-scale wind-tunnel-model results in predicting full-scale airplane characteristics needs to be reexamined. Extensive tests have been made by NASA on several sizes of wind-tunnel models of the XC-142A V/STOL configuration and by Ling-Temco-Vought, Inc., on the airplane. These results show that models predict the slow-speed level-flight characteristics very well but that small models underpredict the descent capability of the airplane. Larger scale wind-tunnel models of approximately half size show better agreement with the airplane descent characteristics. It was found from smoke flow studies that small models can also predict the region in which self-generated disturbances will be encountered by tilt-wing configurations in ground proximity.
An investigation has been conducted in the Langley Unitary Plan wind tunnel to determine the effects of spherical nose bluntness on the aerodynamic characteristics of a typical cruciform-finned short-range missile configuration for roll angles of 0 deg and 45 deg and Mach numbers from 1.50 to 2.86. Also investigated was the effectiveness of a forward protruding nose spike in reducing the axial-force coefficient of a blunt-nose configuration. The results indicated that the effects of nose bluntness and nose spike on the longitudinal stability were small. For roll angles of both 0 deg and 45 deg, the greatest effects of nose bluntness on the aerodynamic-center location near zero angle of attack occurred at a Mach number of 2.86, where for each roll position, a rearward shift of approximately 3.5 percent of the body length occurred as bluntness was increased. Although large magnitudes of nose bluntness were accompanied by high axial-force coefficients, significant reductions in the axial-force coefficient of a blunt-nose configuration were achieved with the use of the nose spike. There were no noteworthy effects of nose bluntness on either the directional stability or side-force parameters. However, for high angles of attack at a Mach number of 1.50 increased nose bluntness generally resulted in decreased effective dihedral.
Aerodynamic data obtained in a 14-inch trisonic wind tunnel test on static stability and drag of 0.003366-scale models of the Grumman H-33 drop tank orbiter and the Titan T 3 (1207-4) expendable booster are presented. The orbiter was mounted to the booster both piggyback and tandem. Data were obtained over a Mach number range of 0.6 to 5.0. Angle of attack data were obtained from -16 deg to +4 deg at both 0 deg and -6 deg sideslip for the piggyback configuration, but were limited to -10 deg to +10 deg at 0 deg sideslip for the tandem configuration. Primary configuration variations, other than piggyback and tandem mount, were the number and location of solid propellant strap-on boosters.
For several proposed configurations of the modular space station, the gravity gradient torque distribution is examined for both Y-POP and inertial hold orbits. The necessary equations are derived to allow for solar panel gimbaling, and a procedure is established which will allow any desired space station configuration to be built up. Results indicated that solar panel gimbaling can be a significant factor in the torque distribution in Y-POP orbits. In addition, for the two configurations that were plotted, bias torque magnitudes varied widely in inertial hold modes.
Experimental aerodynamic investigations were conducted in the NASA/MSFC 14-inch trisonic wind tunnel during March 1972 on a .003366 scale model of a solid rocket motor version of the space shuttle ascent configuration. The configuration consisted of a parallel burn solid rocket motor booster on an external H-O centerline tank orbiter. Six component aerodynamic force and moment date were recorded over an angle of attack range from -10 to 10 deg at zero degrees sideslip and over a sideslip range from -10 to 10 deg at 0, +6, and -6 deg angle of attack. Mach number ranged from 0.6 to 4.96. The performance and stability characteristics of the complete ascent configuration and build-up, and the effects of variations in tank diameter, orbiter incidence, fairings and positioning of the solid rocket motors and tank fins were determined.
Experimental aerodynamic investigations were conducted in the NASA/MSFC 14-inch trisonic wind tunnel during April 1972 on a 0.004-scale model of a solid rocket motor version of the space shuttle ascent configuration. The configuration consisted of a parallel burn solid rocket motor booster on an external HO centerline tank orbiter. Six component aerodynamic force and moment data were recorded over an angle of attack range from -10 deg to +10 deg at zero degrees sideslip and over a sideslip range from -10 deg to +10 deg at zero degrees angle of attack. Mach numbers ranged from 0.6 to 4.96. The purpose of the test was to determine the performance and stability characteristics of the complete ascent configuration and buildup, and to determine the effects of variations in HO tank and SRM nose shaping, orbiter incidence and position, and position of the solid rocket motors.
A series of comparisons made on the longitudinal and lateral directional stability and control characteristics of the baseline 040A space shuttle orbiter configuration is documented. The data used for the comparisons were obtained from five wind tunnel tests conducted in four different wind tunnel facilities. The media for comparison were the aerodynamic slope values plotted as a function of Mach number for the different facilities. The slopes were computed using the SADSAC system and have been reduced relative to a common reference system. Data used for the comparison were selected based on a common baseline configuration. Identical configurations were available in all cases except MSFC 551 where an off-baseline vertical tail was tested. Some of the factors considered in comparing the various data were balance operating ranges and sensitivities, Reynolds number differences, scaling effects and control deflection accuracies.
Experimental aerodynamic investigations were carried out in the Boeing transonic and supersonic wind tunnels on a 0.008899-scale model of a proposed pressure-fed ballistic recoverable booster (BRB) configuration. The purpose of the test program was to determine the stability and control effectiveness of the basic configuration at high and low angles of attack, and to conduct parametric studies of various engine shroud, fin, and drag petal configurations. Six-component force data and base pressure data were obtained over a Mach number range of 0.35 to 4.0 at angles of attack of -5 to 25 and 55 to 85 at zero degrees sideslip and over a sideslip range of -10 to +10 at angles of attack ranging from -10 to 72.5. Two-component force data were also obtained with a fin balance on selected runs.
A parametric study of low-boom supersonic transport airplanes with conventional configurations was made to identify the features of specific configurations that promise relatively low sonic boom overpressures (less than 47.9 N/sq m). The range of values considered was gross weight from 28,300 to 170,000 kg; cruise Mach numbers of 2 to 3.2; and wing loadings of 1436, 2870, and 4309 N/sq m. Fuselage length was varied from 49.1 to 102.4 m and fuselage diameter from 2.75 to 3.98 m. A nominal Mach 2 configuration weighing 56,700 kg and having a wing loading of 2870 N/sq m was selected; and its gross geometric, aerodynamic, and structural features were estimated. At a cruise altitude of 18,300 m, lift-drag ratio was estimated to be 7.35, while sonic boom overpressure was 41.7 N/sq m. Takeoff thrust loading using four afterburning turbojet engines at maximum dry thrust was 0.32. Payload for a 4440-km range was 16.7 percent of gross weight, giving a direct operating cost of 0.82 cent per seat statute mile.
The 049 orbiter and launch configurations were tested in a trisonic wind tunnel to obtain preliminary loads information on the orbiter alone, orbiter in presence of the H/O tank and orbiter in the full launch configuration. The orbiter consisted of the baseline 049 double-delta wing, twin vertical stabilizers, seven degrees of dihedral and included abort rockets. The orbiter was mounted at minus 1.50 degree of incidence (fuselage centerline relative to H/O tank centerline) in the launch configuration. The solid rocket motors were mounted at a radial location of 21 degrees from the horizontal centerline of the H/O tank. The test was conducted over a Mach number range of 0.6 to 4.96. Nominal angle of attack and angle of sideslip ranges of minus 6 to plus 6 degrees were tested. In addition, the orbiter alone was tested over an angle of attack range of minus 6 to plus 26 degrees.
Configuration data and design information for the Atlas Centaur launched configuration are presented. Overall system definition, operations and control, and telecommunication service system, including link budgets, are discussed. A brief description of the user telecommunications equipment and ground station is presented. A summary description of the TDR spacecraft and all the subsystems is included. The data presented are largely in tabular form. A brief treatment of an optional configuration with enhanced telecommunications service is described.
Noise data were obtained with a small scale model stationary STOL configuration that used an eight lobe mixer nozzle with deflector mounted above a 32-cm-chord wing section. The factors varied to determine their effect upon the noise were wing flap angle, nozzle shape, nozzle location, deflector configuration, and jet velocity. The noise from the mixer nozzle model was compared to the noise from a model using a circular nozzle of the same area. The mixer nozzle model was quieter at the low to middle frequencies, while the circular nozzle was quieter at high frequencies. The perceived noise level (PNL) was calculated for an aircraft 10 times larger than the model. The PNL at 500 feet for the mixer nozzle turned out to be within 1 db of the PNL for the circular nozzle. For some configurations at highly directional broadband noise, which could be eliminated by changes in nozzle and/or deflector location, occurred below the wing.
Concurrent simulations of powered-lift STOL transport aircraft having either an externally blown flap configuration or an augmentor wing configuration were conducted. The following types of simulators of varying sophistication were used: (1) a simple fixed-base simulation with a simple visual display, (2) a more complex fixed-base simulation using a realistic transport cockpit and a high-quality visual display, and (3) a six-degree-of-freedom motion simulator that had a realistic transport cockpit and a sophisticated visual display. The unaugmented flying qualities determined from these simulations were rated as unacceptable for both the externally blown flap and augmentor wing configurations. The longitudinal, lateral-directional, and single-engine-failure characteristics were rated satisfactory with extensive augmentation, including pitch and roll command systems, flight-path (or speed) augmentation, turn coordination, and effective yaw damping. However, the flare and landing characteristics from any approach glide-path angle in excess of 4 deg were rated as unsatisfactory but acceptable.
A range network is defined to consist of ground stations and targets where only distances between the two sets of points are observed. Such a network is said to be fundamental when only those six constraints are used that are needed to define the coordinate system for an adjustment. When ground stations or targets have certain configurations, a unique adjustment in terms of coordinates may be impossible, even when the number of observations is sufficient and the coordinate system is uniquely defined. Such configurations, resulting in the singular solutions, are said to be critical. Attention is therefore focused on singular solutions rather than on numerical problems in some ill-conditioned systems of fundamental range networks. The critical configurations are presented in two separate groups. The first deals with ground stations lying all in a plane and the second deals with ground stations generally distributed. The two kinds of problems require different mathematical treatments and lead to quite different conclusions.
In the application of Monte Carlo computer techniques to the study of problems in the free molecular flow regime, the analogy between radiation heat transfer with diffuse emissions and the free molecular flow field makes it possible to use the radiation configuration factors as a test on the operation of the Monte Carlo program. A source plane which was an annular ring of arbitrary radius around a sphere with the center of the disk coincident with the center of the sphere was examined. Since various indicators seemed to verify the correct operation of the program, the program was used to calculate the radiation configuration factors. The study adds a new geometry to the general literature on radiation configuration factors.
Ground-wind load studies were conducted on three model configurations to assess the importance of aeroelastic instabilities of erected space shuttle vehicles. Roll damping was measured on a fuselage-alone model, which had a D cross section, and a fuselage and tail surfaces in combination with either a clipped-delta wing or a low-sweep tapered wing as the primary lifting surface. The largest negative roll-damping coefficients were measured with the fuselage-alone configuration and were a function of wind azimuth. At the wind azimuths at which the wing-fuselage configuration was unstable, the negative roll-damping coefficients were a function of reduced frequency.