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Generator, comprised largely of integrated circuits, has operational amplifier connected as integrator to provide linear voltage ramp, pair of logic gates and one-shot multivibrator function as sweep reverse circuit feeding the integrator. Solid state switch effects generator hold capability.
A power factor type motor controller in which the conventional power factor constant voltage command signal is replaced during a starting interval with a graduated control voltage. The present invention adds to the three-phase system of pending application Ser. No. 199,765, filed Oct. 23, 1980, means for modifying the operation of the system for a motor start-up interval of 5 to 30 seconds. The modification is that of providing via ramp generator 174 an initial ramp-like signal which replaces a constant power factor signal supplied by potentiometer 70. The ramp-like signal is applied to terminal 40 where it is summed with an operating power factor signal from phase detectors 32, 34, and 36 to thereby obtain a control signal for ultimately controlling SCR devices 12, 14, and 16 to effect a gradual turn-on of motor 10. The significant difference of the present invention over prior art is that the SCR devices are turned on at an advancing rate with time responsive to the combination signal described rather than simply a function of a ramp-like signal alone. The added signal, the operating power factor signal, enables the production of a control signal which effectively eliminates a prior problem with many motor starting circuits, which is that of accompanying motor instabilities.
Relative performance of several passive and active methods for controlling two-dimensional turbulent separated flow associated with a curved backward-facing ramp were investigated at low speeds. Surface static pressure measurement and oil flow visualization results indicate that submerged vortex generators, vortex generator jets, elongated arches at +-alpha, and large-eddy breakup devices at +-alpha placed near the baseline separation location reduce flow separation and increase pressure recovery. Spanwise cylinders reduce flow separation but decrease pressure recovery downstream. Arches with alpha = 0 deg, Helmholtz resonators, and Viets' fluidic flappers examined so far have no significant effect in reducing separation. Wall cooling computation indicates that separation delay on a partially cooled ramp is nearly the same as on a fully-cooled ramp while minimizing the frictional drag increase associated with the wall cooling process.
The external combustion of hydrogen to reduce transonic drag was investigated. A control volume analysis is developed and indicates that the specific impulse performance of external burning is competitive with other forms of airbreathing propulsion and depends on the fuel-air ratio, freestream Mach number, and the severity of the base drag. A method is presented for sizing fuel injectors for a desired fuel-air ratio in the unconfined stream. A two-dimensional Euler analysis is also presented which indicates that the total axial force generated by external burning depends on the total amount of energy input and is independent of the transverse and streamwise distribution of heat addition. Good agreement between the Euler and control volume analysis is demonstrated. Features of the inviscid external burning flowfield are discussed. Most notably, a strong compression forms at the sonic line within the burning stream which may induce separation of the plume and prevent realization of the full performance potential. An experimental program was conducted in a Mach 1.26 free-jet to demonstrate drag reduction on a simple expansion ramp geometry, and verify hydrogen-air stability limits at external burning conditions. Stable combustion appears feasible to Mach number of between 1.4 and 2 depending on the vehicle flight trajectory. Drag reduction is demonstrated on the expansion ramp at Mach 1.26; however, force levels showed little dependence on fuel pressure or altitude in contrast to control volume analysis predictions. Various facility interference mechanisms and scaling issues were studied and are discussed.
As one of the first technology development programs awarded by NASA under the Vision for Space Exploration, the Pratt & Whitney Rocketdyne (PWR) Deep Throttling, Common Extensible Cryogenic Engine (CECE) program was selected by NASA in November 2004 to begin technology development and demonstration toward a deep throttling, cryogenic engine supporting ongoing trade studies for NASA s Lunar Lander descent stage. The CECE program leverages the maturity and previous investment of a flight-proven hydrogen/oxygen expander cycle engine, the PWR RL10, to develop and demonstrate an unprecedented combination of reliability, safety, durability, throttlability, and restart capabilities in high-energy, cryogenic, in-space propulsion. The testbed selected for the deep throttling demonstration phases of this program was a minimally modified RL10 engine, allowing for maximum current production engine commonality and extensibility with minimum program cost. Four series of demonstrator engine tests have been successfully completed between April 2006 and April 2010, accumulating 7,436 seconds of hot fire time over 47 separate tests. While the first two test series explored low power combustion (chug) and system instabilities, the third test series investigated and was ultimately successful in demonstrating several mitigating technologies for these instabilities and achieved a stable throttling ratio of 13:1. The fourth test series significantly expanded the engine s operability envelope by successfully demonstrating a closed-loop control system and extensive transient modeling to enable lower power engine starting, faster throttle ramp rates, and mission-specific ignition testing. The final hot fire test demonstrated a chug-free, minimum power level of 5.9%, corresponding to an overall 17.6:1 throttling ratio achieved. In total, these tests have provided an early technology demonstration of an enabling cryogenic propulsion concept with invaluable system-level technology data acquisition toward design and development risk mitigation for future lander descent main engines.
This presentation describes current work under UEET Active Flow Control CFD Research Tool Development. The goal of this work is to develop computational tools for inlet active flow control design. This year s objectives were to perform CFD simulations of fully gridded vane vortex generators, micro-vortex genera- tors, and synthetic jets, and to compare flowfield results with wind tunnel tests of simple geometries with flow control devices. Comparisons are shown for a single micro-vortex generator on a flat plate, and for flow over an expansion ramp with sidewall effects. Vortex core location, pressure gradient and oil flow patterns are compared between experiment and computation. This work lays the groundwork for evaluating simplified modeling of arrays of devices, and provides the opportunity to test simple flow control device/sensor/ control loop interaction.
Flow separation in internal and external flows generally results in a significant degradation in aircraft performance. For internal flows, such as inlets and transmission ducts in aircraft propulsion systems, separation is undesirable as it reduces the overall system performance. The aim of this research has been to understand the nature of separation and more importantly, to explore techniques to actively control it. In this research, we extended our investigation of active separation control (under a previous NASA grant) where we explored the use of microjets for the control of boundary layer separation. The geometry used for the initial study was a simple diverging Stratford ramp, equipped with arrays of microjets. These early results clearly show that the activation of microjets eliminated flow separation. Furthermore, the velocity-field measurements, using PIV, also demonstrate that the gain in momentum due to the elimination of separation is at least an order of magnitude larger (two orders of magnitude larger in most cases) than the momentum injected by the microjets and is accomplished with very little mass flow through the microjets. Based on our initial promising results this research was continued under the present grant, using a more flexible model. This model allows for the magnitude and extent of separation as well as the microjet parameters to be independently varied. The results, using this model were even more encouraging and demonstrated that microjet control completely eliminated significant regions of flow separation over a wide range of conditions with almost negligible mass flow. Detailed studies of the flowfield and its response to microjets were further examined using 3-component PIV and unsteady pressure measurements, among others. As the results presented this report will show, microjets were successfully used to control the separation of a much larger extent and magnitude than demonstrated in our earlier experiments. In fact, using the appropriate combination of control parameters (microjet, location, angle and pressure) separation was completely eliminated for the largest separated flowfield we could generate with the present model. Separation control also resulted in a significant reduction in the unsteady pressures in the flow where the unsteady pressure field was found to be directly responsive to the state of the flow above the surface. Hence, our study indicates that the unsteady pressure signature is a strong candidate for a flow state sensor , which can be used to estimate the location, magnitude and other properties of the separated flowfield. Once better understood and properly utilized, this behavior can be of significant practical importance for developing and implementing online control.
A scalable, compact oven testbed system for simultaneously evaluating a multitude of high temperature integrated circuits (ICs) for prolonged operating times of up to 600 °C has been prototyped. The new testbed system enables long-duration high temperature testing in sufficient statistical quantities consistent with standard aerospace electronics engineering standards. This setup is comprised of multiple compact ovens housing chips or packages mounted to ceramic circuit boards. Each oven is a compact 15.2 cm length by 15.2 cm width by 12.7 cm depth with a maximum 400 Watts of heating power. The custom-made silicon oxide ceramic heating block inside each oven is based on a 3D printed design adapted for the easy insertion of the IC device under test (DUT). This innovative design provides the quick insertion of ICs with or without a ceramic package into a 600 °C environment by utilizing a movable 11.43 cm long ceramic substrate with electrical traces extending from the oven hot zone to external standard plastic-based board connectors. Another key oven design feature is the minimization of the DUT exposure to electromagnetic interference (EMI) by utilizing a filtered DC power source to reduce heating element noise. Additionally, the ovens can be configured in a parallel arrangement allowing global data monitoring over a single industrial RS422 serial port. This feature is important for scaling up to test multiple ICs semi-simultaneously. A USB serial port is provided to independently control the operating parameters of each oven such as the oven target temperature and oven ramp rate. The oven temperature can reach up to 600 °C with a confirmed +/- 4 °C maximum deviation across the test zone region. The ramp rate can be programmed from 1 °C/minute up to 10 °C/minute. Furthermore, a programmable switchboard is used to interface with the DUT. This switchboard comprises a National Instruments Peripheral Component Interconnect eXtension for Instrumentation (NI PXI) system and a breakout board to send and receive power, analog or digital test signals. By using this unique oven testbed system, a variety of ICs can now be tested in parallel using the same test components configured for a diverse set of requirements.
The Common Extensible Cryogenic Engine (CECE) is a deep throttling cryogenic Lunar Module Descent Engine (LMDE) technology development demonstrator based on the Pratt & Whitney Rocketdyne (PWR) RL10 engine. Previous testing on this engine occurred during 2006 as Demo 1.0, then Demo 1.5 in 2007, Demo 1.6 in 2008 and finally Demo 1.7 in 2010. A review of Demo 1.0, 1.5, and preliminary results of 1.6 were reported in previous JANNAF papers. Demo 1.6 was tested at the PWR E-6 test facility in November 2008. The primary goal of this series was to mitigate low frequency combustion instability observed at low power levels. To mitigate the chug, the Demo 1.6 injector was modified from the previous configurations to include an approximately 0.050 inch thick teflon-type spray-on insulation to reduce heat transfer to the LOX manifold, which was believed to be a significant contributor to the low power instability. In addition, gaseous helium injection into the LOX manifold was used as a means to stabilize the system. Also explored in this test series was mitigation for a low power 1 Hz fuel system oscillation caused by sub-critical hydrogen boiling in the chamber cooling jacket. Reduced area gas venturis were utilized to avoid the 1 Hz fuel-size oscillation by keeping the cooling jacket supercritical down to lower engine power levels. Demo 1.7 began testing in March 2010. Its primary objectives were to demonstrate closed loop control capability on mixture ratio and chamber pressure, start to lower power levels and increased throttling ramp rates. Secondary test objectives that are discussed include multiple engine starts, higher mixture ratio excursions and additional time at low power level. While the complete test series and data reduction is not yet complete for Demo 1.7, an overview and up-to-date status is provided.
The German Aerospace Center (DLR) and the National Aeronautics and Space Administration (NASA) have been collaborating to conduct joint research addressing future surface traffic management challenges. The surface management tool from DLR, called Taxi Routing for Aircraft: Creation and Controlling (TRACC), was adapted to be integrated in NASA's fast-time simulation environment called Surface Operations Simulator and Scheduler (SOSS). The research described in this paper 1) applied TRACC to trajectory-based ramp traffic management, where TRACC generates conflict-free aircraft trajectories in a congested ramp area, 2) investigated the feasibility of the concept through the integrated TRACC-SOSS fast-time simulation, and 3) evaluated the performance of the integrated system. For this activity, TRACC was adapted for ramp operations at Charlotte Douglas International Airport, called TRACC_PB (TRACC for pushback optimization). TRACC_PB provides four-dimensional taxi trajectories with a command speed profile for each aircraft following standard taxi routes within the ramp area. In this study, departures are given the Target Movement Area entry Times (TMATs) provided by the baseline surface metering scheduler based on NASA's Spot and Runway Departure Advisor (SARDA). TRACC_PB also calculates optimal pushback times for departures, as well as the times when arrivals shall enter the ramp, the Target Movement area Exit Times (TMETs). The initial results showed that the TRACC_PB successfully generated conflict-free trajectories for the ramp area taxi operations and improved taxiing efficiency compared to the baseline results. TRACC_PB aimed to provide conflict-free taxi routes avoiding any stops while taxiing. This resulted in longer gate hold times for departures and postponed throughput values compared to the baseline simulation without trajectory optimization. Having conflict-free routes without stoppage also created shorter taxi times but required renegotiation of the given TMATs. TRACC_PB also achieved reductions in both fuel consumption and engine emissions (17% for departures and 10% for arrivals), which correlate with the ramp taxi time reduction.
Devices for generating streamwise vorticity in a boundary includes various forms of vortex generators. One form of a split-ramp vortex generator includes a first ramp element and a second ramp element with front ends and back ends, ramp surfaces extending between the front ends and the back ends, and vertical surfaces extending between the front ends and the back ends adjacent the ramp surfaces. A flow channel is between the first ramp element and the second ramp element. The back ends of the ramp elements have a height greater than a height of the front ends, and the front ends of the ramp elements have a width greater than a width of the back ends.
Within this paper, control-relevant vehicle design concepts are examined using a widely used 3 DOF (plus flexibility) nonlinear model for the longitudinal dynamics of a generic carrot-shaped scramjet powered hypersonic vehicle. Trade studies associated with vehicle/engine parameters are examined. The impact of parameters on control-relevant static properties (e.g. level-flight trimmable region, trim controls, AOA, thrust margin) and dynamic properties (e.g. instability and right half plane zero associated with flight path angle) are examined. Specific parameters considered include: inlet height, diffuser area ratio, lower forebody compression ramp inclination angle, engine location, center of gravity, and mass. Vehicle optimizations is also examined. Both static and dynamic considerations are addressed. The gap-metric optimized vehicle is obtained to illustrate how this control-centric concept can be used to "reduce" scheduling requirements for the final control system. A classic inner-outer loop control architecture and methodology is used to shed light on how specific vehicle/engine design parameter selections impact control system design. In short, the work represents an important first step toward revealing fundamental tradeoffs and systematically treating control-relevant vehicle design.
The requirement for greater tactical aircraft operational capabilities has led to increasing research emphasis on the refinement of engine/airframe integration methods and exhaust nozzle flexibility. A major prospective advancement in the development of these capabilities takes the form of multifunctional exhaust nozzle systems with thrust reversal and thrust vectoring features, whose operation will be shared by both airframe and powerplant control systems. Attention is presently given to the two-dimensional convergent-divergent and single expansion ramp nozzle designs, with emphasis on the variable geometry mechanical systems by which they assume cruising flight, vectoring, and thrust reversal operations. The nozzles have been wind tunnel model-tested for the cases of the F-18 fighter and a supersonic cruise configuration concept.
The design of a joint-space adaptive control scheme is presented for controlling the slave arm motion of a dual-arm telerobot system developed at Goddard Space Flight Center (GSFC) to study telerobotic operations in space. Each slave arm of the dual-arm system is a kinematically redundant manipulator with 7 degrees of freedom (DOF). Using the concept of model reference adaptive control (MRAC) and Lyapunov direct method, an adatation algorithm is derived which adjusts the PD controller gains of the control scheme. The development of the adaptive control scheme assumes that the slave arm motion is non-compliant and slowly-varying. The implementation of the derived control scheme does not need the computation of the manipulator dynamics, which makes the control scheme sufficiently fast for real-time applications. Computer simulation study performed for the 7-DOF slave arm shows that the developed control scheme can efficiently adapt to sudden change in payloads while tracking various test trajectories such as ramp or sinusoids with negligible position errors.
Examination of the literature shows that the comparison between experiment and computation for highly separated unswept compression ramp flows is generally poor, irrespective of the turbulence model used. In general, the upstream influence is not correct, the wall pressure rise through separation is too steep, and the pressures under the separated shear layer are too high. In the current study, the objective is to determine if these discrepancies might be attributed more to other factors such as flowfield unsteadiness or three-dimensionality, rather than to inadequate turbulence modeling. To examine this possibility, multichannel wall pressure fluctuations were measured under the unsteady separation shock wave in a 28-deg unswept compression ramp flow at Mach 5. The results show that the large scale, low frequency separation shock unsteadiness controls the distribution of time-averaged surface properties and that neglect of the unsteadiness is probably the primary cause of the discrepancy between experiment and computation.
This paper presents the design of a joint-space adaptive control scheme for controlling the slave arm motion of a dual-arm telerobot system developed at Goddard Space Flight Center (GSFC) to study telerobotic operations in space. Each slave arm of the dual-arm system is a kinematically redundant manipulator with seven degrees of freedom (DOF). Using the concept of model reference adaptive control (MRAC) and Liupunov direct method, we derive an adaptation algorithm that adjusts the PD controller gains of the control scheme. The development of the adaptive control scheme assumes that the slave arm motion is non-compliant and slowly varying. The implementation of the derived control scheme does not require the computation of manipulator dynamics which makes the control scheme sufficiently fast for real-time applications. Computer simulation study performed for the 7-DOF slave arm shows that the developed control scheme can efficiently adapt to sudden change in payload while tracking various test trajectories such as ramp or sinusoids with negligible position errors.
This paper presents a hierarchical adaptive nonlinear iteration method (HANIM) implemented in the NASA computational fluid dynamics code, FUN3D, to improve robustness and computational efficiency. In contrast to the legacy FUN3D iterative solver that relies on an approximate Jacobian, a simple multicolor Gauss-Seidel point-implicit iteration scheme, and linear Courant-Friedrichs-Lewy number (CFL) ramping, HANIM is based upon a hierarchy of modules including preconditioner, generalized conjugate residual, realizability check, nonlinear control, and CFL adaption modules. HANIM performance is systematically compared with the performance of the legacy solver of FUN3D and a baseline solver based on a preconditioner alone. Iterative solutions are compared for three benchmark cases: a subsonic separated flow around a hemisphere cylinder, a supersonic flow through a long duct, and a subsonic flow over the NASA wing-fuselage juncture model. Two Reynolds-averaged Navier-Stokes turbulence models are used in these computations, namely, the negative variant of the linear one-equation Spalart-Allmaras model and its nonlinear extension based on quadratic constitutive relations.
Some of the technical requirements for an expedient application of temperature control to advanced protein crystal growth activities are defined. Lysozome was used to study the effects of temperature ramping and temperature gradients for nucleation/dissolution and consecutive growth of sizable crystals and, to determine a prototype temperature program. The solubility study was conducted using equine serum albumin (ESA) which is an extremely stable, clinically important protein due to its capability to bind and transport many different small ions and molecules.