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Results for “PROPELLERS, CONTROLLABLE - PITCH”

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 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Design of power electronics for TVC and EMA systems

The Component Development Division of the Propulsion Laboratory at Marshall Space Flight Center (MSFC) is currently developing a class of electromechanical actuators (EMA's) for use in space transportation applications such as thrust vector control (TVC) and propellant control valves (PCV). These high power servomechanisms will require rugged, reliable, and compact power electronic modules capable of modulating several hundred amperes of current at up to 270 volts. MSFC has selected the brushless dc motor for implementation in EMA's. A previous project performed by Auburn University examined the use of the resonant dc link (RDCL) inverter, pulse density modulation (PDM), and mos-controlled thyristors (MCT's) for speed control of a brushless dc motor. The speed of the brushless dc motor is proportional to the applied stator voltage. In a PDM system, the control system determines the number of resonant voltage pulses which must be applied to the stator to achieve a desired speed. The addition of a waveshaping circuit to the front end of a standard three-phase inverter yields a RDCL inverter; the resonant voltage pulses are produced through the action of this wave shaping circuit and the inverter. This project has focused on the implementation of a system which permits zero-voltage switching with the bus voltage clamped at the input voltage level. In the same manner as the RDCL inverter, the inverter selected for this implementation is a combination of waveshaping circuit and a standard three-phase inverter. In addition, this inverter allows a pulse-width modulated (PWM)-like control scheme instead of a PDM scheme. The operation of waveshaping circuit will be described through analysis and waveforms. Design relationships will also be presented.

Nelms, R. Mark↗
NASA NTRSDate not supplied

The Zero-Boil-Off Tank Experiment: the Effects of Non-Condensable Gases on Pressurization and Pressure Control of Propellant Tanks in Microgravity

Introduction Integral to all phases of NASA’s projected planetary expeditions is affordable and reliable cryogenic fluid storage for use in propellant or life support systems. It is greatly advantageous to develop innovative vent-less pressure control designs based on cooling/mixing of the bulk tank fluid to allow storage of the cryogenic fluid with zero or reduced boil-off. The presence of noncondensable gases, can interfere with the condensation at the interface impacting tank pressure control during subcooled jet mixing especially in microgravity. The Zero-Boil-Off Tank (ZBOT) Experiments are a series of small-scale experiments aboard the International Space Station (ISS) that use a transparent volatile simulant fluid in a transparent sealed tank to delineate various fundamental fluid flow, heat and mass transport, and phase change phenomena associated with storage tank pressurization and pressure control in microgravity. The ZBOT-1 experiment was performed on the ISS 2017-2018 timeframe and collected data to validate a state-of-the-art CFD model for tank pressurization and pressure control for a pure system. The ZBOT-NC Experiment is the second experiment in the series to be performed on the International Space Station (ISS) in 2025. Its goal is to investigate the effects of noncondensable gases on interfacial evaporation and condensation during self -pressurization and jet-mixing pressure control in microgravity for a two-component system. Materials & Methods In this work, we will describe the detailed features of the ZBOT-NC experimental hardware and Diagnostics that includes nonintrusive Quantum Dot Thermometry (QDT) for whole field temperature measurement. All microgravity pressurization and pressure control tests will be performed for both the pure and the two-components systems with Xenon and Neon as the two noncondensable gases spanning small and large molecular weights and sizes. The two-phase CFD model that is developed as part of the project will be also presented and discussed. Results Ground-based pressurization and jet mixing experiments and CFD simulation results are compared to each other to validate both the fidelity of the CFD model predictions and the accuracy of the QDT measurement. Model simulations for noncondensable gas effects will also be compared against large Cryogenic LH2-GHe experiments to indicate the noncondensable gas effects on tank pressure control during jet mixing in 1G. Finally, CFD results will be presented to predict the effects of the noncondensable gas during subcooled jet mixing during the ZBOT-NC in advance of the microgravity experiment in 2015.

Evaporation Condensation↗
NASA NTRS2018

Auralization of an Unmanned Aerial Vehicle Under Propeller Phase Control

Community noise has been identified as a major barrier to entry into service of unmanned aerial vehicle platforms. Common among such vehicle architectures is a distributed electric propulsion system, in which many propellers and/or rotors are used for thrust and/or lift. A promising noise reduction technique was recently developed to favorably modify the vehicle source noise directivity, by adjusting the relative azimuthal position of the prop/rotor blades between propulsors. Un- der idealized conditions, such phase control can yield noise reductions on the order of 20 dB at a ground observer relative to the average case having a random phase distribution. However, human perception of the resulting noise may not be fully captured by such a simple acoustic measure. To this end, this paper under- takes the auralization of the tonal vehicle noise. Specifically, total propeller tonal noise is optimized over a set of observer zones, e.g., 45 deg. wedges, around the vehicle. Each zone represents a unique set of relative phases. The forward flight of the vehicle adjacent to a noise-sensitive area is simulated by prescribing a series of zones over the flight path having minimum noise in the direction of the ob- server. Auralizations and their associated metrics are compared with a reference condition.

Pascioni, Kyle A.↗
NASA NTRS1993

Cryogenic propellant thermal control system design considerations, analyses, and concepts applied to a Mars human exploration mission

This paper analyzes, defines, and sizes cryogenic storage thermal control systems that meet the requirements of future NASA Mars human exploration missions. The design issues of this system include the projection of the existing Multilayer Insulation data base for cryogenic storage to much thicker (10 cm or more) insulation systems, the unknown heat leak from mechanical interfaces, and the thermal and structural performance effects of the large tank sizes required for a Mars mission. Acknowledging these unknown effects, heat loss projections are made based on extrapolation of the existing data base. The results indicate that hydrogen, methane, and oxygen are feasible propellants, and that the best suited thermal control sytems are 'thick' MLI, thermodynamic vent sytems, cryocoolers, and vacuum jackets.

Plachta, David W.↗
NASA NTRS2010

Ares I Reaction Control System Propellant Feedline Decontamination Modeling

The objective of the work presented here is to quantify the effects of purge gas temperature, pressure, and mass flow rate on Hydrazine (Hz) decontamination rates of the Ares I Roll Control System and Reaction Control System. A survey of experts in this field revealed the absence of any decontamination rate prediction models. Three basic decontamination methods were identified for analysis and modeling. These include low pressure eduction, high flow rate purge, and pulse purge. For each method, an approach to predict the Hz mass transfer rate, as a function of system pressure, temperature, and purge gas mass flow rate, is developed based on the applicable physics. The models show that low pressure eduction is two orders of magnitude more effective than the high velocity purge, which in turn is two orders of magnitude more effective than the pure diffusion component of pulse purging of deadheads. Eduction subjects the system to low pressure conditions that promote the extraction of Hz vapors. At 120 F, Hz is saturated at approximately 1 psia. At lower pressures and 120 F, Hz will boil, which is an extremely efficient means to remove liquid Hz. The Hz boiling rate is predicted by equating the rate at which energy is added to the saturated liquid Hz through heaters at the tube outer wall with the energy removed from the liquid through evaporation. Boil-off fluxes were predicted by iterating through the range of local pressures with limits set by the minimum allowed pressure of 0.2 psia and maximum allowed wall temperature of 120 F established by the heaters, which gives a saturation pressure of approximately 1.0 psia. Figure 1 shows the resulting boil-off fluxes as a function of local eduction pressure. As depicted in figure 1, the flux is a strong inverse function of eduction pressure, and that minimizing the eduction pressure maximizes the boil-off flux. Also, higher outer wall temperatures lead to higher boil-off fluxes and allow for boil-off over a greater range of eduction pressures.

Pasch, James J.↗
NASA NTRS1991

Adaptive control of propellant slosh for a launch vehicle with multiple tanks

This paper discusses the use of an adaptive system to stabilize a multi-tank launch vehicle with sloshable fuel. The adaptive system utilizes a sensor to measure the fluid slosh angle, a recursive-least-squares estimator to estimate the natural frequency of the sloshing fuel, and a linear filter which stabilizes the unstable fuel slosh using the frequency estimate. Single-input and multiinput estimators were investigated. The time-varying complexity of the slosh control loop keeps the system persistently excited through flight, even though the attitude is controlled by a conventional PID controller. For the first portion of flight, the adaptive system maintains stability. A modification is proposed for latter portions of flight. The approach requires no tank baffles or partitions.

Saugen, John D.↗
NASA NTRSDate not supplied

Overview and Assessment of the ESM Pressure Control Performance on Artemis I

The European Service Module propulsion system is a bipropellant hypergolic serial system used to provide translational thrust and attitude control for Orion. To control propellant tank pressure, a bang-bang pressure control system is employed. Each propellant commodity is regulated by a pressure control assembly consisting of two pressurization branches (a primary and redundant pressurization path) where each branch includes 3 valves in series. Regulation is accomplished via flight software control of two downstream solenoid valves triggered off propellant tank ullage pressure. This paper presents an overview of system level challenges which have been overcome to enable a successful Artemis I flight. Principle among the challenges was valve-to-valve pneumatic interactions which drove changes to the control scheme. During the Artemis I mission, the pressure control assembly was able to control tank pressure within allowable tolerances. Comparison between flight data and mathematical models are presented showing excellent agreement. Finally, during flight, a pressure surge was observed during the first regulation cycle when there was propellant in the upstream propellant tank. This was attributed to a gas hammer effect within the pressurization system and was not observable in a 1g environment. This paper also discusses the conclusion that this gas hammer effect is a nominal feature of the system during operations. Assessment of the in-flight performance of the electronic pressure regulation scheme on the European Service Module propulsion system shows the system behaved nominally during the Artemis I mission.

propulsion system↗
NASA NTRSDate not supplied

Overview and Assessment of the ESM Pressure Control Performance on Artemis I

The European Service Module propulsion system is a bipropellant hypergolic serial system used to provide translational thrust and attitude control for Orion. To control propellant tank pressure, a bang-bang pressure control system is employed. Each propellant commodity is regulated by a pressure control assembly consisting of two pressurization branches (a primary and redundant pressurization path) where each branch includes 3 valves in series. Regulation is accomplished via flight software control of two downstream solenoid valves triggered off propellant tank ullage pressure. This paper presents an overview of system level challenges which have been overcome to enable a successful Artemis I flight. Principle among the challenges was valve-to-valve pneumatic interactions which drove changes to the control scheme. During the Artemis I mission, the pressure control assembly was able to control tank pressure within allowable tolerances. Comparison between flight data and mathematical models are presented showing excellent agreement. Finally, during flight, a pressure surge was observed during the first regulation cycle when there was propellant in the upstream propellant tank. This was attributed to a gas hammer effect within the pressurization system and was not observable in a 1g environment. This paper also discusses the conclusion that this gas hammer effect is a nominal feature of the system during operations. Assessment of the in-flight performance of the electronic pressure regulation scheme on the European Service Module propulsion system shows the system behaved nominally during the Artemis I mission.

propulsion system↗
NASA NTRS2016

MLI Blanket Effective Emittance Variance and Its Effect on Spacecraft Propellant Line Thermal Control

Multi-Layer Insulation (MLI) blankets commonly employed for temperature control of spacecraft propellant lines consistently show large variances in their effective emittances. These variances are primarily due to workmanship, hence are hard to control or predict. Unfortunately, the variances are also the primary drivers of propellant line temperatures, thus leading to large predicted temperature uncertainties and increased risks to the success of space missions. JPL has gained significant knowledge of these variances from Mars spacecraft designs over more than two decades. This knowledge has led to improved analytical approaches to bound existing designs, emphasis on increased temperature visibility for design verification and model correlation, as well as methods for improving the robustness of propellant line thermal designs.

Miller, Jennifer R.↗
NASA NTRS1974

Design of H2-O2 space shuttle APU. Volume 1: APU design

The H2-O2 space shuttle auxiliary power unit (APU) program is a NASA-Lewis effort aimed at hardware demonstration of the technology required for potential use on the space shuttle. It has been shown that a hydrogen-oxygen power unit (APU) system is an attractive alternate to the space shuttle baseline hydrazine APU system for minimum weight. It has the capability for meeting many of the heat sink requirements for the space shuttle vehicle, thereby reducing the amount of expendable evaporants required for cooling in the baseline APU. Volume 1 of this report covers preliminary design and analysis of the current reference system and detail design of the test version of this reference system. Combustor test results are also included. Volume 2 contains the results of the analysis of an initial version of the reference system and the computer printouts of system performance. The APU consists of subsystems for propellant feed and conditioning, turbopower, and control. Propellant feed and conditioning contains all heat exchangers, valves, and the combustor. The turbopower subsystem contains a two-stage partial-admission pressure-modulated, 400-hp, 63,000-rpm turbine, a 0-to 4-g lubrication system, and a gearbox with output pads for two hydraulic pumps and an alternator (alternator not included on test unit). The electronic control functions include regulation of speed and system temperatures; and start-and-stop sequences, overspeed (rpm) and temperature limits, failsafe provisions, and automatic shutdown provisions.

Harris, E.↗
NASA NTRS1978

Effect of silicone oil on nozzle heat transfer in solid propellant rockets

The feasibility of reducing heat transfer rates to inert nozzle walls in solid rockets by addition of silicone oil as a propellant ingredient is explored experimentally. A metallized control propellant and its counterpart with 1% silicone oil are fired in a 2 1/2-in. diameter end burning all-metal burner. Pressure-time histories are obtained; temperature-time data at the nozzle throat are obtained in some of the runs. The heat loss by the product gases to the cold inert walls of the nozzle lower the temperature of the gases leading to the condensation of aluminum oxide on the walls. This condensation is used as an index of the heat transfer rate to the nozzle. On a statistical basis it is found that the silicone propellant gives practically no condensation in 70% of the firings contrasted with 21% for its counterpart with no silicone oil in it. The gas temperature at the nozzle throat is experimentally seen to be generally higher with the silicone propellant than with its counterpart; this higher temperature is consistent with a preliminary analysis that is used here to relate the nozzle surface temperature to the heat transfer from the product gases.

Ramohalli, K.↗
NASA NTRS1937

Flight Tests of the Drag and Torque of the Propeller in Terminal-Velocity Dives

The drag and torque of a controllable propeller at various blade-angle settings, and under various diving conditions, were measured by indirect method on F6C-4 airplane in flight. The object of these tests were (1) to provide data on which calculations of the terminal velocity with a throttled engine and the accompanying engine speed could be based and (2) to determine the possibility of utilizing the propeller as an air brake to reduce the terminal velocity. The data obtained were used in the establishment of propeller charts, on the basis of which the terminal velocity and engine speed could be calculated for airplanes whose characteristics fall within the range of these tests. A method is given for the calculation of the terminal velocity with throttled engine and the engine speed.

Rhode, Ricahrd V↗
NASA NTRS1973

A feasibility study of developing toroidal tanks for a spinning spacecraft

A study was made to determine the feasibility of developing toroidal propellant tanks for a bipropellant (N204/MMH) propulsion system to be used in a proposed advanced Pioneer spin-stabilized vehicle intended for a Jupiter-orbiter and possibly a Saturn-orbiter mission. The rationale for considering the use of two toroidal tanks rather than the proposed use of four spherical tanks includes the belief that a more symmetrical distribution of propellant mass and a smaller variation in the position of the vehicle center-of-mass during propellant consumption would result, reducing requirements for attitude-control propellants, for balance weight, and for other weights associated with the dynamics of the spinning spacecraft. Results lead to the conclusion that a toroidal tank containing an effective, passive surface tension propellant acquisition device could be fabricated with available manufacturing methods and could be used interchangeably for either fuel or oxidizer.

Anderson, J. E.↗
NASA NTRS2018

Stationkeeping, Orbit Determination, and Attitude Control for Spacecraft in Near Rectilinear Halo Orbits

Final document is attached. From a Near Rectilinear Halo Orbit (NRHO), NASA's Gateway at the Moon is planned to serve as a proving ground and a staging location for human missions beyond Earth. Stationkeeping, Orbit Determination (OD), and attitude control are examined for uncrewed and crewed Gateway configurations. Orbit maintenance costs are investigated using finite maneuvers, considering skipped maneuvers and perturbations. OD analysis assesses DSN tracking and identifies OD challenges associated with the NRHO and crewed operations. The Gateway attitude profile is simulated to determine an effective equilibrium attitude. Attitude control propellant use and sizing of the required passive attitude control system are assessed.

Newman, Clark P.↗