Propelled abrasive grit applications for weed management in transitional corn grain production systems

USDA-ARS?s Scientific Manuscript database

Weed control is challenging to farmers who are transitioning from production systems that use synthetic herbicides to organic systems. A two-year field study examined weed control efficacy and corn grain yield of air-propelled corncob grit abrasion for in-row weed control. Grits were applied based o...

Reynolds numbers exhibit dramatic influence on directionality of movement of self-propelled systems.

PubMed

Zhao, Guanjia; Pumera, Martin

2012-05-14

Self-propelled artificial objects are at the current forefront of research. We demonstrate here that the motion directionality of millimetre sized self-propelled objects is highly dependent on the Reynolds numbers (Re) of the systems, with emphasis on the "intermediate" Re region (1-600). Our findings have strong implication on the motion controllability and predictability of these independent self-propelled systems. This journal is © the Owner Societies 2012

Register of specialized sources for information on selected fuels and oxidizers. [rocket propellants, bibliographies

NASA Technical Reports Server (NTRS)

Ludtke, P. R.

1975-01-01

Thirty-eight (38) organizations are listed and described that catalog and file information in their data systems on fuel and oxidizers. The fuels include hydrogen, methane and hydrazine-type fuels; the oxidizers include oxygen, fluorine, flox, nitrogen tetroxide and ozone. The type of available information covers thermophysical properties, propellant systems, propellant fires-control-extinguishment, propellant explosions, propellant combustion, propellant safety, and fluorine chemistry. These organizations have assembled and collated their information so that it will be useful in the solution of engineering problems.

14 CFR 35.23 - Propeller control system.

Code of Federal Regulations, 2013 CFR

2013-01-01

... between operating modes, performs the functions defined by the applicant throughout the declared operating... system imbedded software must be designed and implemented by a method approved by the Administrator that... software errors. (d) The propeller control system must be designed and constructed so that the failure or...

14 CFR 35.23 - Propeller control system.

Code of Federal Regulations, 2012 CFR

2012-01-01

... between operating modes, performs the functions defined by the applicant throughout the declared operating... system imbedded software must be designed and implemented by a method approved by the Administrator that... software errors. (d) The propeller control system must be designed and constructed so that the failure or...

14 CFR 35.23 - Propeller control system.

Code of Federal Regulations, 2014 CFR

2014-01-01

... between operating modes, performs the functions defined by the applicant throughout the declared operating... system imbedded software must be designed and implemented by a method approved by the Administrator that... software errors. (d) The propeller control system must be designed and constructed so that the failure or...

Approach Considerations in Aircraft with High-Lift Propeller Systems

NASA Technical Reports Server (NTRS)

Patterson, Michael D.; Borer, Nicholas K.

2017-01-01

NASA's research into distributed electric propulsion (DEP) includes the design and development of the X-57 Maxwell aircraft. This aircraft has two distinct types of DEP: wingtip propellers and high-lift propellers. This paper focuses on the unique opportunities and challenges that the high-lift propellers--i.e., the small diameter propellers distributed upstream of the wing leading edge to augment lift at low speeds--bring to the aircraft performance in approach conditions. Recent changes to the regulations related to certifying small aircraft (14 CFR x23) and these new regulations' implications on the certification of aircraft with high-lift propellers are discussed. Recommendations about control systems for high-lift propeller systems are made, and performance estimates for the X-57 aircraft with high-lift propellers operating are presented.

Compact and Integrated Liquid Bismuth Propellant Feed System

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Stanojev, Boris; Korman, Valentin; Gross, Jeffrey T.

2007-01-01

Operation of Hall thrusters with bismuth propellant has been shown to be a promising path toward high-power, high-performance, long-lifetime electric propulsion for spaceflight missions [1]. There has been considerable effort in the past three years aimed at resuscitating this promising technology and validating earlier experimental results indicating the advantages of a bismuth-fed Hall thruster. A critical element of the present effort is the precise metering of propellant to the thruster, since performance cannot be accurately assessed without an accurate accounting of mass flow rate. Earlier work used a pre./post-test propellant weighing scheme that did not provide any real-time measurement of mass flow rate while the thruster was firing, and makes subsequent performance calculations difficult. The motivation of the present work is to develop a precision liquid bismuth Propellant Management System (PMS) that provides hot, molten bismuth to the thruster while simultaneously monitoring in real-time the propellant mass flow rate. The system is a derivative of our previous propellant feed system [2], but the present system represents a more compact design. In addition, all control electronics are integrated into a single unit and designed to reside on a thrust stand and operate in the relevant vacuum environment where the thruster is operating, significantly increasing the present technology readiness level of liquid metal propellant feed systems. The design of various critical components in a bismuth PMS are described. These include the bismuth reservoir and pressurization system, 'hotspot' flow sensor, power system and integrated control system. Particular emphasis is given to selection of the electronics employed in this system and the methods that were used to isolate the power and control systems from the high-temperature portions of the feed system and thruster. Open loop calibration test results from the 'hotspot' flow sensor are reported, and results of integrated thruster/PMS tests demonstrate operation of the feed system in the relevant environment.

The methodology of variable management of propellant fuel consumption by jet-propulsion engines of a spacecraft

NASA Astrophysics Data System (ADS)

Kovtun, V. S.

2012-12-01

Traditionally, management of propellant fuel consumption on board of a spacecraft is only associated with the operation of jet-propulsion engines (JPE) that are actuator devices of motion control systems (MCS). The efficiency of propellant fuel consumption depends not only on the operation of the MCS, but also, to one extent or another, on all systems functioning on board of a spacecraft, and on processes that occur in them and involve conversion of variable management of propellant fuel consumption by JPEs as a constituent part of the control of the complex process of spacecraft flight.

Detail view of the "underside" of the Orbiter Maneuvering/Reaction Control ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Detail view of the "underside" of the Orbiter Maneuvering/Reaction Control Systems pod looking at the two spherical propellant tanks for the Reaction Control System, and the elongated propellant tanks for the Orbiter Maneuvering System. This view was taken at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Electrospray Thrusters for Attitude Control of a 1-U CubeSat

NASA Astrophysics Data System (ADS)

Timilsina, Navin

With a rapid increase in the interest in use of nanosatellites in the past decade, finding a precise and low-power-consuming attitude control system for these satellites has been a real challenge. In this thesis, it is intended to design and test an electrospray thruster system that could perform the attitude control of a 1-unit CubeSat. Firstly, an experimental setup is built to calculate the conductivity of different liquids that could be used as propellants for the CubeSat. Secondly, a Time-Of-Flight experiment is performed to find out the thrust and specific impulse given by these liquids and hence selecting the optimum propellant. On the other hand, a colloidal thruster system for a 1-U CubeSat is designed in Solidworks and fabricated using Lathe and CNC Milling Machine. Afterwards, passive propellant feeding is tested in this thruster system. Finally, the electronic circuit and wireless control system necessary to remotely control the CubeSat is designed and the final testing is performed. Among the propellants studied, Ethyl ammonium nitrate (EAN) was selected as the best propellant for the CubeSat. Theoretical design and fabrication of the thruster system was performed successfully and so was the passive propellant feeding test. The satellite was assembled for the final experiment but unfortunately the microcontroller broke down during the first test and no promising results were found out. However, after proving that one thruster works with passive feeding, it could be said that the ACS testing would have worked if we had performed vacuum compatibility tests for other components beforehand.

On-Board Propulsion System Analysis of High Density Propellants

NASA Technical Reports Server (NTRS)

Schneider, Steven J.

1998-01-01

The impact of the performance and density of on-board propellants on science payload mass of Discovery Program class missions is evaluated. A propulsion system dry mass model, anchored on flight-weight system data from the Near Earth Asteroid Rendezvous mission is used. This model is used to evaluate the performance of liquid oxygen, hydrogen peroxide, hydroxylammonium nitrate, and oxygen difluoride oxidizers with hydrocarbon and metal hydride fuels. Results for the propellants evaluated indicate that the state-of-art, Earth Storable propellants with high performance rhenium engine technology in both the axial and attitude control systems has performance capabilities that can only be exceeded by liquid oxygen/hydrazine, liquid oxygen/diborane and oxygen difluoride/diborane propellant combinations. Potentially lower ground operations costs is the incentive for working with nontoxic propellant combinations.

Propulsion System Development for the Iodine Satellite (iSAT) Demonstration Mission

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Peeples, Stephen R.; Seixal, Joao F.; Mauro, Stephanie L.; Lewis, Brandon L.; Jerman, Gregory A.; Calvert, Derek H.; Dankanich, John; Kamhawi, Hani; Hickman, Tyler A.;

Liquid Rocket Booster (LRB) for the Space Transportion System (STS) systems study. Appendix D: Trade study summary for the liquid rocket booster

NASA Technical Reports Server (NTRS)

1989-01-01

Trade studies plans for a number of elements in the Liquid Rocket Booster (LRB) component of the Space Transportation System (STS) are given in viewgraph form. Some of the elements covered include: avionics/flight control; avionics architecture; thrust vector control studies; engine control electronics; liquid rocket propellants; propellant pressurization systems; recoverable spacecraft; cryogenic tanks; and spacecraft construction materials.

Propellant Feed Subsystem for a 26 kW flight arcjet propulsion system

NASA Astrophysics Data System (ADS)

Vaughan, C. E.; Morris, J. P.

1993-06-01

The USAF arcjet ATTD program demanded the development of a low-cost ammonia Propellant Feed Subsystem (PFS). A flow rate of 240 +/- 5 mg/sec during a total of ten 15-min ammonia outflows was required for the flight mission. The precision of the flow tolerance required a departure from the design of previous ammonia propellant feed systems. Since a propellant management device was not used, thermocapillary forces were explored as a means to limit outflow of liquid phase ammonia. A high energy density feedline heater with an internal wick was developed to guarantee that only gas phase propellant would reach the arcjet. A digital control algorithm was developed to implement bang-bang control of mass flow rate metered by a sonic venturi. Development tests of this system have been completed. The system is capable of continuous gas phase outflows regardless of orientation. Integrated tests with the arcjet and power conditioning unit have also been successfully completed.

Controlled propulsion of artificial magnetic nanostructured propellers.

PubMed

Ghosh, Ambarish; Fischer, Peer

2009-06-01

For biomedical applications, such as targeted drug delivery and microsurgery, it is essential to develop a system of swimmers that can be propelled wirelessly in fluidic environments with good control. Here, we report the construction and operation of chiral colloidal propellers that can be navigated in water with micrometer-level precision using homogeneous magnetic fields. The propellers are made via nanostructured surfaces and can be produced in large numbers. The nanopropellers can carry chemicals, push loads, and act as local probes in rheological measurements.

Iodine Hall Thruster Propellant Feed System for a CubeSat

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Peeples, Steven

2014-01-01

The components required for an in-space iodine vapor-fed Hall effect thruster propellant management system are described. A laboratory apparatus was assembled and used to produce iodine vapor and control the flow through the application of heating to the propellant reservoir and through the adjustment of the opening in a proportional flow control valve. Changing of the reservoir temperature altered the flowrate on the timescale of minutes while adjustment of the proportional flow control valve changed the flowrate immediately without an overshoot or undershoot in flowrate with the requisite recovery time associated with thermal control systems. The flowrates tested spanned a range from 0-1.5 mg/s of iodine, which is sufficient to feed a 200-W Hall effect thruster.

TC-2 post Helios experiment data review. [postflight systems analysis of spacecraft performance

NASA Technical Reports Server (NTRS)

1975-01-01

Data are presented from a systems postflight analysis of the Centaur Launch Vehicle and Helios. Also given is a comparison of data from preflight analyses. Topics examined are: (1) propellant behavior; (2) helium usage; (3) propellant tank pressurization; (4) propellant tank thermodynamics; (5) component heating; thermal control; and thermal protection system; (6) main engine system; (7) H2O2 consumption; (8) boost pump post-meco performance; and (9) an overview of other systems.

Linear Test Bed. Volume 2: Test Bed No. 2. [linear aerospike test bed for thrust vector control

NASA Technical Reports Server (NTRS)

1974-01-01

Test bed No. 2 consists of 10 combustors welded in banks of 5 to 2 symmetrical tubular nozzle assemblies, an upper stationary thrust frame, a lower thrust frame which can be hinged, a power package, a triaxial combustion wave ignition system, a pneumatic control system, pneumatically actuated propellant valves, a purge and drain system, and an electrical control system. The power package consists of the Mark 29-F fuel turbopump, the Mark 29-0 oxidizer turbopump, a gas generator assembly, and propellant ducting. The system, designated as a linear aerospike system, was designed to demonstrate the feasibility of the concept and to explore technology related to thrust vector control, thrust vector optimization, improved sequencing and control, and advanced ignition systems. The propellants are liquid oxygen/liquid hydrogen. The system was designed to operate at 1200-psia chamber pressure at an engine mixture ratio of 5.5. With 10 combustors, the sea level thrust is 95,000 pounds.

Design Factors for Applying Cryogen Storage and Delivery Technology to Solar Thermal Propulsion

NASA Technical Reports Server (NTRS)

Millis, Marc G.

1996-01-01

Thermodynamic Vent System (TVS) and Multilayer Insulation (MLI) technology, originally developed for long term storage of cryogen propellants in microgravity, is ideally suited for propellant storage and delivery systems for solar thermal propulsion. With this technology the heat-induced pressure rise in the tank provides the propellant delivery pressure without the need for an auxiliary pressurant system, and propellant delivery is used to remove the excess heat to control tank pressure. The factors to consider in designing such a balanced system, are presented. An example of a minimum system design is presented along with examples of laboratory-tested hardware.

Low-g simulation testing of propellant systems using neutral buoyancy

NASA Technical Reports Server (NTRS)

Balzer, D. L.; Lake, R. J., Jr.

1972-01-01

A two liquid, neutral buoyancy technique is being used to simulate propellant behavior in a weightless environment. By equalizing the density of two immiscible liquids within a container (propellant tank), the effect of gravity at the liquid interface is balanced. Therefore the surface-tension forces dominate to control the liquid/liquid system configuration in a fashion analogous to a liquid/gas system in a zero gravity environment.

A new method for aerodynamic test of high altitude propellers

NASA Astrophysics Data System (ADS)

Gong, Xiying; Zhang, Lin

A ground test system is designed for aerodynamic performance tests of high altitude propellers. The system is consisted of stable power supply, servo motors, two-component balance constructed by tension-compression sensors, ultrasonic anemometer, data acquisition module. It is loaded on a truck to simulate propellers' wind-tunnel test for different wind velocities at low density circumstance. The graphical programming language LABVIEW for developing virtual instrument is used to realize the test system control and data acquisition. Aerodynamic performance test of a propeller with 6.8 m diameter was completed by using this system. The results verify the feasibility of the ground test method.

SS/RCS surface tension propellant acquisition/expulsion tankage technology program

NASA Technical Reports Server (NTRS)

1974-01-01

An evaluation of published propellant physical property data together with bubble point tests of fine-mesh screen in propellants, was conducted. The effort consisted of: (1) the collection and evaluation of pertinent physical property data for hydrazine (N2H4), monomethylhydrazine (MMH), and nitrogen tetroxide (N2O4); (2) testing to determine the effect of dissolved pressurant gas, temperature, purity, and system cleanliness or contamination on system bubble point, and (3) the compilation and publishing of both the literature and test results. The space shuttle reaction control system (SS/RCS) is a bipropellant system using N2O4 and MMH, while the auxiliary power system (SS/APU) employs monopropellant N2H4. Since both the RCS and the APU use a surface tension device for propellant acquisition, the propellant properties of interest are those which impact the design and operation of surface tension systems. Information on propellant density, viscosity, surface tension, and contact angle was collected, compiled, and evaluated.

Wetter for fine dry powder

DOEpatents

Hall, James E.; Williams, Everett H.

1977-01-01

A system for wetting fine dry powders such as bentonite clay with water or other liquids is described. The system includes a wetting tank for receiving water and a continuous flow of fine powder feed. The wetting tank has a generally square horizontal cross section with a bottom end closure in the shape of an inverted pyramid. Positioned centrally within the wetting tank is a flow control cylinder which is supported from the walls of the wetting tank by means of radially extending inclined baffles. A variable speed motor drives a first larger propeller positioned immediately below the flow control cylinder in a direction which forces liquid filling the tank to flow downward through the flow control cylinder and a second smaller propeller positioned below the larger propeller having a reverse pitch to oppose the flow of liquid being driven downward by the larger propeller.

Advanced Launch Vehicle Upper Stages Using Liquid Propulsion and Metallized Propellants

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan A.

1990-01-01

Metallized propellants are liquid propellants with a metal additive suspended in a gelled fuel or oxidizer. Typically, aluminum (Al) particles are the metal additive. These propellants provide increase in the density and/or the specific impulse of the propulsion system. Using metallized propellant for volume-and mass-constrained upper stages can deliver modest increases in performance for low earth orbit to geosynchronous earth orbit (LEO-GEO) and other earth orbital transfer missions. Metallized propellants, however, can enable very fast planetary missions with a single-stage upper stage system. Trade studies comparing metallized propellant stage performance with non-metallized upper stages and the Inertial Upper Stage (IUS) are presented. These upper stages are both one- and two-stage vehicles that provide the added energy to send payloads to altitudes and onto trajectories that are unattainable with only the launch vehicle. The stage designs are controlled by the volume and the mass constraints of the Space Transportation System (STS) and Space Transportation System-Cargo (STS-C) launch vehicles. The influences of the density and specific impulse increases enabled by metallized propellants are examined for a variety of different stage and propellant combinations.

Propelled abrasive grit for weed control in organic silage corn

USDA-ARS?s Scientific Manuscript database

Weed management in organic farming requires many strategies to accomplish acceptable control and maintain crop yields. This two-year field study used air propelled abrasive grit for in-row weed control in a silage corn system. Corncob grit was applied as a single application at corn vegetative growt...

Structureborne noise control in advanced turboprop aircraft

NASA Technical Reports Server (NTRS)

Loeffler, Irvin J.

1987-01-01

Structureborne noise is discussed as a contributor to propeller aircraft interior noise levels that are nonresponsive to the application of a generous amount of cabin sidewall acoustic treatment. High structureborne noise levels may jeopardize passenger acceptance of the fuel-efficient high-speed propeller transport aircraft designed for cruise at Mach 0.65 to 0.85. These single-rotation tractor and counter-rotation tractor and pusher propulsion systems will consume 15 to 30 percent less fuel than advanced turbofan systems. Structureborne noise detection methodologies and the importance of development of a structureborne noise sensor are discussed. A structureborne noise generation mechanism is described in which the periodic components or propeller swirl produce periodic torques and forces on downstream wings and airfoils that are propagated to the cabin interior as noise. Three concepts for controlling structureborne noise are presented: (1) a stator row swirl remover, (2) selection of a proper combination of blade numbers in the rotor/stator system of a single-rotation propeller, and the rotor/rotor system of a counter-rotation propeller, and (3) a tuned mechanical absorber.

A Users Guide for the NASA ANOPP Propeller Analysis System

NASA Technical Reports Server (NTRS)

Nguygen, L. Cathy; Kelly, Jeffrey J.

1997-01-01

The purpose of this report is to document improvements to the Propeller Analysis System of the Aircraft Noise Prediction Program (PAS-ANOPP) and to serve as a users guide. An overview of the functional modules and modifications made to the Propeller ANOPP system are described. Propeller noise predictions are made by executing a sequence of functional modules through the use of ANOPP control statements. The most commonly used ANOPP control statements are discussed with detailed examples demonstrating the use of each control statement. Originally, the Propeller Analysis System included the angle-of-attack only in the performance module. Recently, modifications have been made to also include angle-of-attack in the noise prediction module. A brief description of PAS prediction capabilities is presented which illustrate the input requirements necessary to run the code by way of ten templates. The purpose of the templates are to provide PAS users with complete examples which can be modified to serve their particular purposes. The examples include the use of different approximations in the computation of the noise and the effects of synchrophasing. Since modifications have been made to the original PAS-ANOPP, comparisons of the modified ANOPP and wind tunnel data are also included. Two appendices are attached at the end of this report which provide useful reference material. One appendix summarizes the PAS functional modules while the second provides a detailed discussion of the TABLE control statement.

Long life reaction control system design

NASA Astrophysics Data System (ADS)

Fanciullo, Thomas J.; Judd, Craig

1993-02-01

Future single stage to orbit systems will utilize oxygen/hydrogen propellants in their main propulsion means due to the propellant's high energy content and environmental acceptability. Operational effectiveness studies and life cycle cost studies have indicated that minimizing the number of different commodities on a given vehicle not only reduces cost, but reduces the ground span times in both the pre- and postflight operations. Therefore, oxygen and hydrogen should be used for the reaction controls systems, eliminating the need to deal with toxic or corrosive fluids. When the hydrogen scramjet powered NASP design development began in 1985, new system design studies considered overall integration of subsystems; in the context of that approach, O2/H2 reaction controls system were more than competitive with storable propellant systems and had the additional benefits of lower life cycle cost, rapid turnaround times, and O2 and H2 commodities for use throughout the vehicle. Similar benefits were derived in rocket-powered SSTO vehicles.

Development of a biowaste resistojet propulsion system

NASA Technical Reports Server (NTRS)

1973-01-01

The equipment, exclusive of thrustors, required to demonstrate the feasibility of a resistojet propulsion system for space station attitude control application using representative simulated crew biowaste propellants and available resistojet thrustors in the ground simulation tests is discussed. The overall objective of the program was to provide a biowaste resistojet prototype propellant management and control system sufficiently similar to the flight article to permit concept feasibility and system demonstration testing of interface compatibility, operational characteristics, and system flexibility.

Development of Static Balance Measurement and Correction Compound Platform for Single Blade of Controllable Pitch Propeller

NASA Astrophysics Data System (ADS)

Chao, Zhang; Shijie, Su; Yilin, Yang; Guofu, Wang; Chao, Wang

2017-11-01

Aiming at the static balance of the controllable pitch propeller (CPP), a high efficiency static balance method based on the double-layer structure of the measuring table and gantry robot is adopted to realize the integration of torque measurement and corrected polish for controllable pitch propeller blade. The control system was developed by Microsoft Visual Studio 2015, and a composite platform prototype was developed. Through this prototype, conduct an experiment on the complete process of torque measurement and corrected polish based on a 300kg class controllable pitch propeller blade. The results show that the composite platform can correct the static balance of blade with a correct, efficient and labor-saving operation, and can replace the traditional method on static balance of the blade.

Practical Methodology for the Inclusion of Nonlinear Slosh Damping in the Stability Analysis of Liquid-Propelled Space Vehicles

NASA Technical Reports Server (NTRS)

Ottander, John A.; Hall, Robert A.; Powers, Joseph F.

2017-01-01

One of the challenges of developing flight control systems for liquid-propelled space vehicles is ensuring stability and performance in the presence of parasitic minimally damped slosh dynamics in the liquid propellants. This can be especially difficult when the fundamental frequencies of the slosh motions are in proximity to the frequency used for vehicle control. The challenge is partially alleviated since the energy dissipation and effective damping in the slosh modes increases with amplitude. However, traditional launch vehicle control design methodology is performed with linearized systems using a fixed slosh damping corresponding to a slosh motion amplitude based on heritage values. This papers presents a method for performing the control design and analysis using damping at slosh amplitudes chosen based on the resulting limit cycle amplitude of the vehicle thrust vector system due to a control-slosh interaction under degraded phase and gain margin conditions.

Practical Methodology for the Inclusion of Nonlinear Slosh Damping in the Stability Analysis of Liquid-propelled Space Vehicles

NASA Technical Reports Server (NTRS)

Ottander, John A.; Hall, Robert A., Jr.; Powers, Joseph F.

2017-01-01

One of the challenges of developing flight control systems for liquid-propelled space vehicles is ensuring stability and performance in the presence of parasitic minimally damped slosh dynamics in the liquid propellants. This can be especially difficult when the fundamental frequencies of the slosh motions are in proximity to the frequency used for vehicle control. The challenge is partially alleviated since the energy dissipation and effective damping in the slosh modes increases with amplitude. However, traditional launch vehicle control design methodology is performed with linearized systems using a fixed slosh damping corresponding to a slosh motion amplitude based on heritage values. This papers presents a method for performing the control design and analysis using damping at slosh amplitudes chosen based on the resulting limit cycle amplitude of the vehicle thrust vector system due to a control-slosh interaction under degraded phase and gain margin conditions.

Two-Pendulum Model of Propellant Slosh in Europa Clipper PMD Tank

NASA Technical Reports Server (NTRS)

Ng, Wanyi; Benson, David

2017-01-01

Model propellant slosh for Europa Clipper using two pendulums such that controls engineers can predict slosh behavior during the mission. Importance of predicting propellant slosh; (1) Sloshing changes CM (center of mass) of spacecraft and exerts forces and torques on spacecraft. (2) Avoid natural frequencies of structures. (3) Size ACS (Attitude Control Systems) thrusters to counteract forces and torques. Can model sloshing fluid as two pendulums with specific parameters (mass, length, damping),

Perceived Annoyance to Noise Produced by a Distributed Electric Propulsion High Lift System

NASA Technical Reports Server (NTRS)

Palumbo, Dan; Rathsam, Jonathan; Christian, Andrew; Rafaelof, Menachem

2016-01-01

Results of a psychoacoustic test performed to understand the relative annoyance to noise produced by several configurations of a distributed electric propulsion high lift system are given. It is found that the number of propellers in the system is a major factor in annoyance perception. This is an intuitive result as annoyance increases, in general, with frequency, and, the blade passage frequency of the propellers increases with the number of propellers. Additionally, the data indicate that having some variation in the blade passage frequency from propeller-to-propeller is beneficial as it reduces the high tonality generated when all the propellers are spinning in synchrony at the same speed. The propellers can be set to spin at different speeds, but it was found that allowing the motor controllers to drift within 1% of nominal settings produced the best results (lowest overall annoyance). The methodology employed has been demonstrated to be effective in providing timely feedback to designers in the early stages of design development.

A Discussion of Two Challenges of Non-cooperative Satellite Refueling

NASA Technical Reports Server (NTRS)

Coll, Gregory C.; Aranyos, Thomas; Nufer, Brian M.; Kandula, Max; Tomasic, David J.

2015-01-01

There is interest from government and commercial aerospace communities in advancing propellant transfer technology for in-orbit refueling of satellites. This paper introduces two challenges to a Propellant Transfer System (PTS) under development for demonstration of non-cooperative satellite refueling. The PTS is being developed to transfer storable propellant (heritage hypergolic fuels and oxidizers as well as xenon) safely and reliably from one servicer satellite to a non-cooperative typical existing client satellite. NASA is in the project evaluation planning stages for conducting a first time on-orbit demonstration to an existing government asset. The system manages pressure, flow rate totalization, temperature and other parameters to control the condition of the propellant being transferred to the client. It keeps the propellant isolated while performing leak checks of itself and the client interface before transferring propellant. A major challenge is to design a safe, reliable system with some new technologies while maintaining a reasonable cost.

A Discussion of Two Challenges of Non-Cooperative Satellite Refueling

NASA Technical Reports Server (NTRS)

Coll, Gregory T.; Aranyos, Thomas J.; Nufer, Brian M.; Tomasic, David; Kandula, Max

2015-01-01

There is interest from government and commercial aerospace communities in advancing propellant transfer technology for in-orbit refueling of satellites. This paper introduces two challenges to a Propellant Transfer System (PTS) under development for demonstration of non-cooperative satellite refueling. The PTS is being developed to transfer storable propellant (heritage hypergolic fuels and oxidizers as well as xenon) safely and reliably from one servicer satellite to a non-cooperative typical existing client satellite. NASA is in the project evaluation planning stages for conducting a first time on-orbit demonstration to an existing government asset. The system manages pressure, flow rate totalization, temperature and other parameters to control the condition of the propellant being transferred to the client. It keeps the propellant isolated while performing leak checks of itself and the client interface before transferring propellant. A major challenge is to design a safe, reliable system with some new technologies while maintaining a reasonable cost.

In-space propellant systems safety. Volume 3: System safety analysis

NASA Technical Reports Server (NTRS)

1972-01-01

The primary objective was to examine from a system safety viewpoint in-space propellant logistic elements and operations to define the potential hazards and to recommend means to reduce, eliminate or control them. A secondary objective was to conduct trade studies of specific systems or operations to determine the safest of alternate approaches.

Propellantless Attitude Control of Solar Sail Technology Utilizing Reflective Control Devices

NASA Technical Reports Server (NTRS)

Munday, Jeremy

2016-01-01

Solar sails offer an opportunity for a CubeSatscale, propellant-free spacecraft technology that enables long-term and long-distance missions not possible with traditional methods. Solar sails operate using the transfer of linear momentum from photons of sunlight reflected from the surface of the sail. To propel the spacecraft, no mechanically moving parts, thrusters, or propellant are needed. However, attitude control, or orientation, is still performed using traditional methods involving reaction wheels and propellant ejection, which severely limit mission lifetime. For example, the current state of the art solutions employed by upcoming missions couple solar sails with a state of the art propellant ejection gas system. Here, the use of the gas thruster has limited the lifetime of the mission. To solve the limited mission lifetime problem, the Propellantless Attitude Control of Solar Sail Technology Utilizing Reflective Control Devices project team is working on propellantless attitude control using thin layers of material, an optical film, electrically switchable from transparent to reflective. The technology is based on a polymer-dispersed liquid crystal (PDLC), which allows this switch upon application of a voltage. This technology removes the need for propellant, which reduces weight and cost while improving performance and lifetime.

Wind energy conversion system

DOEpatents

Longrigg, Paul

1987-01-01

The wind energy conversion system includes a wind machine having a propeller connected to a generator of electric power, the propeller rotating the generator in response to force of an incident wind. The generator converts the power of the wind to electric power for use by an electric load. Circuitry for varying the duty factor of the generator output power is connected between the generator and the load to thereby alter a loading of the generator and the propeller by the electric load. Wind speed is sensed electro-optically to provide data of wind speed upwind of the propeller, to thereby permit tip speed ratio circuitry to operate the power control circuitry and thereby optimize the tip speed ratio by varying the loading of the propeller. Accordingly, the efficiency of the wind energy conversion system is maximized.

Flow Control of Liquid Metal Propellants for In-Space Electric Propulsion Systems

NASA Technical Reports Server (NTRS)

Bonds, Kevin W.; Polzin, Kurt A.

2010-01-01

Operation of Hall thrusters with bismuth propellant has been shown to be a promising path for development of high-power (140 kW per thruster), high performance (8000s I(sub sp at >70% efficiency) electric propulsion systems.

Autonomous Operations System: Development and Application

NASA Technical Reports Server (NTRS)

Toro Medina, Jaime A.; Wilkins, Kim N.; Walker, Mark; Stahl, Gerald M.

2016-01-01

Autonomous control systems provides the ability of self-governance beyond the conventional control system. As the complexity of mechanical and electrical systems increases, there develops a natural drive for developing robust control systems to manage complicated operations. By closing the bridge between conventional automated systems to knowledge based self-awareness systems, nominal control of operations can evolve into relying on safe critical mitigation processes to support any off-nominal behavior. Current research and development efforts lead by the Autonomous Propellant Loading (APL) group at NASA Kennedy Space Center aims to improve cryogenic propellant transfer operations by developing an automated control and health monitoring system. As an integrated systems, the center aims to produce an Autonomous Operations System (AOS) capable of integrating health management operations with automated control to produce a fully autonomous system.

Flow Control and Measurement in Electric Propulsion Systems: Towards an AIAA Reference Standard

NASA Technical Reports Server (NTRS)

Snyder, John Steven; Baldwin, Jeff; Frieman, Jason D.; Walker, Mitchell L. R.; Hicks, Nathan S.; Polzin, Kurt A.; Singleton, James T.

2013-01-01

Accurate control and measurement of propellant flow to a thruster is one of the most basic and fundamental requirements for operation of electric propulsion systems, whether they be in the laboratory or on flight spacecraft. Hence, it is important for the electric propulsion community to have a common understanding of typical methods for flow control and measurement. This paper addresses the topic of propellant flow primarily for the gaseous propellant systems which have dominated laboratory research and flight application over the last few decades, although other types of systems are also briefly discussed. While most flight systems have employed a type of pressure-fed flow restrictor for flow control, both thermal-based and pressure-based mass flow controllers are routinely used in laboratories. Fundamentals and theory of operation of these types of controllers are presented, along with sources of uncertainty associated with their use. Methods of calibration and recommendations for calibration processes are presented. Finally, details of uncertainty calculations are presented for some common calibration methods and for the linear fits to calibration data that are commonly used.

Method and apparatus to produce high specific impulse and moderate thrust from a fusion-powered rocket engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Cohen, Samuel A.; Pajer, Gary A.; Paluszek, Michael A.

A system and method for producing and controlling high thrust and desirable specific impulse from a continuous fusion reaction is disclosed. The resultant relatively small rocket engine will have lower cost to develop, test, and operate that the prior art, allowing spacecraft missions throughout the planetary system and beyond. The rocket engine method and system includes a reactor chamber and a heating system for heating a stable plasma to produce fusion reactions in the stable plasma. Magnets produce a magnetic field that confines the stable plasma. A fuel injection system and a propellant injection system are included. The propellant injectionmore » system injects cold propellant into a gas box at one end of the reactor chamber, where the propellant is ionized into a plasma. The propellant and fusion products are directed out of the reactor chamber through a magnetic nozzle and are detached from the magnetic field lines producing thrust.« less

Requirements for maintaining cryogenic propellants during planetary surface stays

NASA Technical Reports Server (NTRS)

Riccio, Joseph R.; Schoenberg, Richard J.

1991-01-01

Potential impacts on the planetary surface system infrastructure resulting from the use of liquid hydrogen and oxygen propellants for a stage and half lander are discussed. Particular attention is given to techniques which can be incorporated into the surface infrastructure and/or the vehicle to minimize the impact resulting from the use of these cryogens. Methods offered for reducing cryogenic propellant boiloff include modification of the lander to accommodate boiloff, incorporation of passive thermal control devices to the lander, addition of active propellant management, and use of alternative propellants.

Active Control of Cryogenic Propellants in Space

NASA Technical Reports Server (NTRS)

Notardonato, William

2011-01-01

A new era of space exploration is being planned. Exploration architectures under consideration require the long term storage of cryogenic propellants in space. This requires development of active control systems to mitigate the effect of heat leak. This work summarizes current state of the art, proposes operational design strategies and presents options for future architectures. Scaling and integration of active systems will be estimated. Ideal long range spacecraft systems will be proposed with Exploration architecture benefits considered.

Liquid Methane/Liquid Oxygen Propellant Conditioning Feed System (PCFS) Test Rigs

NASA Technical Reports Server (NTRS)

Skaff, A.; Grasl, S.; Nguyen, C.; Hockenberry S.; Schubert, J.; Arrington, L.; Vasek, T.

2008-01-01

As part of their Propulsion and Cryogenic Advanced Development (PCAD) program, NASA has embarked upon an effort to develop chemical rocket engines which utilize non-toxic, cryogenic propellants such as liquid oxygen (LO2) and liquid methane (LCH4). This effort includes the development and testing of a 100 lbf Reaction Control Engine (RCE) that will be used to evaluate the performance of a LO2/LCH4 rocket engine over a broad range of propellant temperatures and pressures. This testing will take place at NASA-Glenn Research Center's (GRC) Research Combustion Laboratory (RCL) test facility in Cleveland, OH, and is currently scheduled to begin in late 2008. While the initial tests will be performed at sea level, follow-on testing will be performed at NASA-GRC's Altitude Combustion Stand (ACS) for altitude testing. In support of these tests, Sierra Lobo, Inc. (SLI) has designed, developed, and fabricated two separate portable propellant feed systems under the Propellant Conditioning and Feed System (PCFS) task: one system for LCH4, and one for LO2. These systems will be capable of supplying propellants over a large range of conditions from highly densified to several hundred pounds per square inch (psi) saturated. This paper presents the details of the PCFS design and explores the full capability of these propellant feed systems.

Rocket propulsion elements - An introduction to the engineering of rockets (6th revised and enlarged edition)

NASA Astrophysics Data System (ADS)

Sutton, George P.

The subject of rocket propulsion is treated with emphasis on the basic technology, performance, and design rationale. Attention is given to definitions and fundamentals, nozzle theory and thermodynamic relations, heat transfer, flight performance, chemical rocket propellant performance analysis, and liquid propellant rocket engine fundamentals. The discussion also covers solid propellant rocket fundamentals, hybrid propellant rockets, thrust vector control, selection of rocket propulsion systems, electric propulsion, and rocket testing.

Numerical Modeling of Pressurization of Cryogenic Propellant Tank for Integrated Vehicle Fluid System

NASA Technical Reports Server (NTRS)

Majumdar, Alok K.; LeClair, Andre C.; Hedayat, Ali

2016-01-01

This paper presents a numerical model of pressurization of a cryogenic propellant tank for the Integrated Vehicle Fluid (IVF) system using the Generalized Fluid System Simulation Program (GFSSP). The IVF propulsion system, being developed by United Launch Alliance, uses boiloff propellants to drive thrusters for the reaction control system as well as to run internal combustion engines to develop power and drive compressors to pressurize propellant tanks. NASA Marshall Space Flight Center (MSFC) has been running tests to verify the functioning of the IVF system using a flight tank. GFSSP, a finite volume based flow network analysis software developed at MSFC, has been used to develop an integrated model of the tank and the pressurization system. This paper presents an iterative algorithm for converging the interface boundary conditions between different component models of a large system model. The model results have been compared with test data.

14 CFR 23.1149 - Propeller speed and pitch controls.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Propeller speed and pitch controls. 23.1149... Powerplant Controls and Accessories § 23.1149 Propeller speed and pitch controls. (a) If there are propeller... propeller; and (2) Simultaneous control of all propellers. (b) The controls must allow ready synchronization...

14 CFR 23.1149 - Propeller speed and pitch controls.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Propeller speed and pitch controls. 23.1149... Powerplant Controls and Accessories § 23.1149 Propeller speed and pitch controls. (a) If there are propeller... propeller; and (2) Simultaneous control of all propellers. (b) The controls must allow ready synchronization...

14 CFR 23.1149 - Propeller speed and pitch controls.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Propeller speed and pitch controls. 23.1149... Powerplant Controls and Accessories § 23.1149 Propeller speed and pitch controls. (a) If there are propeller... propeller; and (2) Simultaneous control of all propellers. (b) The controls must allow ready synchronization...

Combustion Processes in Solid Propellant Cracks

DTIC Science & Technology

1981-06-01

Ignition at the Closed End of an Inert Ctack . . ......................... 38 12. Block Diagram of Remotely-Controlled Ignition and Photography System ...41 13. Block Diagram of Data Acquisition System ... ........ .. 42 14. Measured Pressure-Time Traces for Crack...ignition system has been designed and fabricated. 5. Experimental firings with single-pore propellant grain have been conducted to study the effects of

Quantitative computer simulations of extraterrestrial processing operations

NASA Technical Reports Server (NTRS)

Vincent, T. L.; Nikravesh, P. E.

1989-01-01

The automation of a small, solid propellant mixer was studied. Temperature control is under investigation. A numerical simulation of the system is under development and will be tested using different control options. Control system hardware is currently being put into place. The construction of mathematical models and simulation techniques for understanding various engineering processes is also studied. Computer graphics packages were utilized for better visualization of the simulation results. The mechanical mixing of propellants is examined. Simulation of the mixing process is being done to study how one can control for chaotic behavior to meet specified mixing requirements. An experimental mixing chamber is also being built. It will allow visual tracking of particles under mixing. The experimental unit will be used to test ideas from chaos theory, as well as to verify simulation results. This project has applications to extraterrestrial propellant quality and reliability.

A Thruster Sub-System Module (TSSM) for solar electric propulsion

NASA Technical Reports Server (NTRS)

Sharp, G. R.

1975-01-01

Solar Electric Propulsion (SEP) is currently being studied for possible use in a number of near-earth and planetary missions. Thruster systems for these missions could be integrated directly into a spacecraft or modularized into a Thruster Sub-System Module (TSSM). A TSSM for electric propulsion missions would consist of a 30-cm ion thruster, thruster gimbal system, propellant storage and feed system, associated Power Processing Unit (PPU), thermal control system and complete supporting structure. The TSSM would be wholly self-contained and be essentially a plug-in or strap-on electric stage with simple mechanical, thermal, electrical and propellant interfaces. The TSSM described in this report is designed for a broad range of missions requiring from two to ten TSSM's mounted in a 2 by x configuration. The thermal control system is designed to accommodate waste heat from the power processor based on realistic efficiencies when the TSSM is operating from 0.7 to 3.5 AU's. The modules are 0.61 M (2 ft) wide by 2.29 M (7.5 ft) long and have a dry weight including propellant tank of 54.4 kg (120 lb). The propellant tank will hold 145.1 kg (320 lb) of mercury.

The hard start phenomena in hypergolic engines. Volume 2: Combustion characteristics of propellants and propellant combinations

NASA Technical Reports Server (NTRS)

Miron, Y.; Perlee, H. E.

1974-01-01

The combustion characteristics of hypergolic propellants are described. A research project was conducted to determine if the reaction control system engine propellants on Apollo spacecraft undergo explosive reaction when subjected to conditions present in the engine at the time of ignition. Combustion characteristics pertinent to the hard-start phenomenon are considered. The thermal stability of frozen mixtures of hydrazine-based fuels with nitrogen tetroxide was analyzed. Results of the tests are presented in the form of tables and graphs.

Design and development of Shuttle Get-Away-Special experiment G-0074. [off-load capability for a full-tank propellant acquisition system

NASA Technical Reports Server (NTRS)

Orton, G. F.

1984-01-01

An experiment to investigate more versatile, lower cost surface tension propellant acquisition approaches for future satellite and spacecraft propellant tanks is designed to demonstrate a propellant off-load capability for a full-tank gallery surface tension device, such as that employed in the shuttle reaction control subsystem, and demonstrate a low-cost refillable trap concept that could be used in future orbit maneuver propulsion systems for multiple engine restarts. A Plexiglas test tank, movie camera and lights, auxiliary liquid accumulator, control electronics, battery pack, and associated valving and plumbing are used. The test liquid is Freon 113, dyed blue for color movie coverage. The fully loaded experiments weighs 106 pounds and is to be installed in a NASA five-cubic-foot flight canister. Vibration tests, acoustic tests, and high and low temperature tests were performed to quality the experiment for flight.

14 CFR 23.1153 - Propeller feathering controls.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Propeller feathering controls. 23.1153... Powerplant Controls and Accessories § 23.1153 Propeller feathering controls. If there are propeller feathering controls installed, it must be possible to feather each propeller separately. Each control must...

14 CFR 23.1153 - Propeller feathering controls.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Propeller feathering controls. 23.1153... Powerplant Controls and Accessories § 23.1153 Propeller feathering controls. If there are propeller feathering controls installed, it must be possible to feather each propeller separately. Each control must...

14 CFR 23.1153 - Propeller feathering controls.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Propeller feathering controls. 23.1153... Powerplant Controls and Accessories § 23.1153 Propeller feathering controls. If there are propeller feathering controls installed, it must be possible to feather each propeller separately. Each control must...

14 CFR 23.1153 - Propeller feathering controls.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Propeller feathering controls. 23.1153... Powerplant Controls and Accessories § 23.1153 Propeller feathering controls. If there are propeller feathering controls installed, it must be possible to feather each propeller separately. Each control must...

14 CFR 23.1153 - Propeller feathering controls.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Propeller feathering controls. 23.1153... Powerplant Controls and Accessories § 23.1153 Propeller feathering controls. If there are propeller feathering controls installed, it must be possible to feather each propeller separately. Each control must...

Design and fabrication of a magnetic propulsion system for self-propelled capsule endoscope.

PubMed

Gao, Mingyuan; Hu, Chengzhi; Chen, Zhenzhi; Zhang, Honghai; Liu, Sheng

2010-12-01

This paper investigates design, modeling, simulation, and control issues related to self-propelled endoscopic capsule navigated inside the human body through external magnetic fields. A novel magnetic propulsion system is proposed and fabricated, which has great potential of being used in the field of noninvasive gastrointestinal endoscopy. Magnetic-analysis model is established and finite-element simulations as well as orthogonal design are performed for obtaining optimized mechanical and control parameters for generating appropriate external magnetic field. Simulated intestinal tract experiments are conducted, demonstrating controllable movement of the capsule under the developed magnetic propulsion system.

H2O2 space shuttle APU

NASA Technical Reports Server (NTRS)

1975-01-01

A cryogenic H2-O2 auxiliary power unit (APU) was developed and successfully demonstrated. It has potential application as a minimum weight alternate to the space shuttle baseline APU because of its (1) low specific propellant consumption and (2) heat sink capabilities that reduce the amount of expendable evaporants. A reference system was designed with the necessary heat exchangers, combustor, turbine-gearbox, valves, and electronic controls to provide 400 shp to two aircraft hydraulic pumps. Development testing was carried out first on the combustor and control valves. This was followed by development of the control subsystem including the controller, the hydrogen and oxygen control valves, the combustor, and a turbine simulator. The complete APU system was hot tested for 10 hr with ambient and cryogenic propellants. Demonstrated at 95 percent of design power was 2.25 lb/hp-hr. At 10 percent design power, specific propellant consumption was 4 lb/hp-hr with space simulated exhaust and 5.2 lb/hp-hr with ambient exhaust. A 10 percent specific propellant consumption improvement is possible with some seal modifications. It was demonstrated that APU power levels could be changed by several hundred horsepower in less than 100 msec without exceeding allowable turbine inlet temperatures or turbine speed.

Polar Satellite Launch Vehicle (PSLV) development programme in India

NASA Astrophysics Data System (ADS)

Janardhana, E.

The design of the Indian Polar Satellite Launch Vehicle (PSLV), for the launching (by 1990) of 1-1.5-tonne payloads into 900-km sun-synchronous orbit, is discussed, and the mission development program is described. The first stage is a solid propellant motor augmented by six solid strap-ons, and the second stage of liquid storable propellant has a high thrust gimballed engine. A high performance solid motor incorporates a flex nozzle for control as the third stage, and the fourth stage is a liquid propulsion system using N204 and MMH propellant with two regeneratively cooled engines. The vehicle equipment bay, housing the inertial guidance and control system, and the TTC system are located around the fourth stage for guidance and tracking with the associated ground segment until spacecraft ejection into orbit.

Physics Simulation Software for Autonomous Propellant Loading and Gas House Autonomous System Monitoring

NASA Technical Reports Server (NTRS)

Regalado Reyes, Bjorn Constant

2015-01-01

1. Kennedy Space Center (KSC) is developing a mobile launching system with autonomous propellant loading capabilities for liquid-fueled rockets. An autonomous system will be responsible for monitoring and controlling the storage, loading and transferring of cryogenic propellants. The Physics Simulation Software will reproduce the sensor data seen during the delivery of cryogenic fluids including valve positions, pressures, temperatures and flow rates. The simulator will provide insight into the functionality of the propellant systems and demonstrate the effects of potential faults. This will provide verification of the communications protocols and the autonomous system control. 2. The High Pressure Gas Facility (HPGF) stores and distributes hydrogen, nitrogen, helium and high pressure air. The hydrogen and nitrogen are stored in cryogenic liquid state. The cryogenic fluids pose several hazards to operators and the storage and transfer equipment. Constant monitoring of pressures, temperatures and flow rates are required in order to maintain the safety of personnel and equipment during the handling and storage of these commodities. The Gas House Autonomous System Monitoring software will be responsible for constantly observing and recording sensor data, identifying and predicting faults and relaying hazard and operational information to the operators.

pH-induced motion control of self-propelled oil droplets using a hydrolyzable gemini cationic surfactant.

PubMed

Miura, Shingo; Banno, Taisuke; Tonooka, Taishi; Osaki, Toshihisa; Takeuchi, Shoji; Toyota, Taro

2014-07-15

Self-propelled motion of micrometer-sized substances has drawn much attention as an autonomous transportation system. One candidate vehicle is a chemically driven micrometer-sized oil droplet. However, to the best of our knowledge, there has been no report of a chemical reaction system controlling the three-dimensional motion of oil droplets underwater. In this study, we developed a molecular system that controlled the self-propelled motion of 4-heptyloxybenzaldehyde oil droplets by using novel gemini cationic surfactants containing carbonate linkages (2G12C). We found that, in emulsions containing sodium hydroxide, the motion time of the self-propelled oil droplets was longer in the presence of 2G12C than in the presence of gemini cationic surfactants without carbonate linkages. Moreover, in 2G12C solution, oil droplets at rest underwent unidirectional, self-propelled motion in a gradient field toward a higher concentration of sodium hydroxide. Even though they stopped within several seconds, they restarted in the same direction. 2G12C was gradually hydrolyzed under basic conditions to produce a pair of the corresponding monomeric surfactants, which exhibit different interfacial properties from 2G12C. The prolonged and restart motion of the oil droplets were explained by the increase in the heterogeneity of the interfacial tension of the oil droplets.

Mars transit vehicle thermal protection system: Issues, options, and trades

NASA Technical Reports Server (NTRS)

Brown, Norman

1986-01-01

A Mars mission is characterized by different mission phases. The thermal control of cryogenic propellant in a propulsive vehicle must withstand the different mission environments. Long term cryogenic storage may be achieved by passive or active systems. Passive cryo boiloff management features will include multilayer insulation, vapor cooled shield, and low conductance structural supports and penetrations. Active boiloff management incorporates the use of a refrigeration system. Key system trade areas include active verses passive system boiloff management (with respect to safety, reliability, and cost) and propellant tank insulation optimizations. Technology requirements include refrigeration technology advancements, insulation performance during long exposure, and cryogenic fluid transfer system for mission vehicle propellant tanking during vehicle buildip in LEO.

Circulation control propellers for general aviation, including a BASIC computer program

NASA Technical Reports Server (NTRS)

Taback, I.; Braslow, A. L.; Butterfield, A. J.

1983-01-01

The feasibility of replacing variable pitch propeller mechanisms with circulation control (Coanada effect) propellers on general aviation airplanes was examined. The study used a specially developed computer program written in BASIC which could compare the aerodynamic performance of circulation control propellers with conventional propellers. The comparison of aerodynamic performance for circulation control, fixed pitch and variable pitch propellers is based upon the requirements for a 1600 kg (3600 lb) single engine general aviation aircraft. A circulation control propeller using a supercritical airfoil was shown feasible over a representative range of design conditions. At a design condition for high speed cruise, all three types of propellers showed approximately the same performance. At low speed, the performance of the circulation control propeller exceeded the performance for a fixed pitch propeller, but did not match the performance available from a variable pitch propeller. It appears feasible to consider circulation control propellers for single engine aircraft or multiengine aircraft which have their propellers on a common axis (tractor pusher). The economics of the replacement requires a study for each specific airplane application.

Average-passage simulation of counter-rotating propfan propulsion systems as applied to cruise missiles

NASA Technical Reports Server (NTRS)

Mulac, Richard A.; Schneider, Jon C.; Adamczyk, John J.

1989-01-01

Counter-rotating propfan (CRP) propulsion technologies are currently being evaluated as cruise missile propulsion systems. The aerodynamic integration concerns associated with this application are being addressed through the computational modeling of the missile body-propfan flowfield interactions. The work described in this paper consists of a detailed analysis of the aerodynamic interactions between the control surfaces and the propfan blades through the solution of the average-passage equation system. Two baseline configurations were studied, the control fins mounted forward of the counter-rotating propeller and the control fins mounted aft of the counter-rotating propeller. In both cases, control fin-propfan separation distance and control fin deflection angle were varied.

NASA Ares I Launch Vehicle Roll and Reaction Control Systems Design Status

NASA Technical Reports Server (NTRS)

Butt, Adam; Popp, Chris G.; Pitts, Hank M.; Sharp, David J.

2009-01-01

This paper provides an update of design status following the preliminary design review of NASA s Ares I first stage roll and upper stage reaction control systems. The Ares I launch vehicle has been chosen to return humans to the moon, mars, and beyond. It consists of a first stage five segment solid rocket booster and an upper stage liquid bi-propellant J-2X engine. Similar to many launch vehicles, the Ares I has reaction control systems used to provide the vehicle with three degrees of freedom stabilization during the mission. During launch, the first stage roll control system will provide the Ares I with the ability to counteract induced roll torque. After first stage booster separation, the upper stage reaction control system will provide the upper stage element with three degrees of freedom control as needed. Trade studies and design assessments conducted on the roll and reaction control systems include: propellant selection, thruster arrangement, pressurization system configuration, and system component trades. Since successful completion of the preliminary design review, work has progressed towards the critical design review with accomplishments made in the following areas: pressurant / propellant tank, thruster assembly, and other component configurations, as well as thruster module design, and waterhammer mitigation approach. Also, results from early development testing are discussed along with plans for upcoming system testing. This paper concludes by summarizing the process of down selecting to the current baseline configuration for the Ares I roll and reaction control systems.

A Study of Fluid Interface Configurations in Exploration Vehicle Propellant Tanks

NASA Technical Reports Server (NTRS)

Zimmerli, Gregory A.; Asipauskas, Marius; Chen, Yongkang; Weislogel, Mark M.

2010-01-01

The equilibrium shape and location of fluid interfaces in spacecraft propellant tanks while in low-gravity is of interest to system designers, but can be challenging to predict. The propellant position can affect many aspects of the spacecraft such as the spacecraft center of mass, response to thruster firing due to sloshing, liquid acquisition, propellant mass gauging, and thermal control systems. We use Surface Evolver, a fluid interface energy minimizing algorithm, to investigate theoretical equilibrium liquid-vapor interfaces for spacecraft propellant tanks similar to those that have been considered for NASA's new class of Exploration vehicles. The choice of tank design parameters we consider are derived from the NASA Exploration Systems Architecture Study report. The local acceleration vector employed in the computations is determined by estimating low-Earth orbit (LEO) atmospheric drag effects and centrifugal forces due to a fixed spacecraft orientation with respect to the Earth or Moon, and rotisserie-type spacecraft rotation. Propellant/vapor interface positions are computed for the Earth Departure Stage and Altair lunar lander descent and ascent stage tanks for propellant loads applicable to LEO and low-lunar orbit. In some of the cases investigated the vapor ullage bubble is located at the drain end of the tank, where propellant management device hardware is often located.

Numerical Modeling of Cavitating Venturi: A Flow Control Element of Propulsion System

NASA Technical Reports Server (NTRS)

Majumdar, Alok; Saxon, Jeff (Technical Monitor)

2002-01-01

In a propulsion system, the propellant flow and mixture ratio could be controlled either by variable area flow control valves or by passive flow control elements such as cavitating venturies. Cavitating venturies maintain constant propellant flowrate for fixed inlet conditions (pressure and temperature) and wide range of outlet pressures, thereby maintain constant, engine thrust and mixture ratio. The flowrate through the venturi reaches a constant value and becomes independent of outlet pressure when the pressure at throat becomes equal to vapor pressure. In order to develop a numerical model of propulsion system, it is necessary to model cavitating venturies in propellant feed systems. This paper presents a finite volume model of flow network of a cavitating venturi. The venturi was discretized into a number of control volumes and mass, momentum and energy conservation equations in each control volume are simultaneously solved to calculate one-dimensional pressure, density, and flowrate and temperature distribution. The numerical model predicts cavitations at the throat when outlet pressure was gradually reduced. Once cavitation starts, with further reduction of downstream pressure, no change in flowrate is found. The numerical predictions have been compared with test data and empirical equation based on Bernoulli's equation.

Mathematical model of marine diesel engine simulator for a new methodology of self propulsion tests

NASA Astrophysics Data System (ADS)

Izzuddin, Nur; Sunarsih, Priyanto, Agoes

2015-05-01

As a vessel operates in the open seas, a marine diesel engine simulator whose engine rotation is controlled to transmit through propeller shaft is a new methodology for the self propulsion tests to track the fuel saving in a real time. Considering the circumstance, this paper presents the real time of marine diesel engine simulator system to track the real performance of a ship through a computer-simulated model. A mathematical model of marine diesel engine and the propeller are used in the simulation to estimate fuel rate, engine rotating speed, thrust and torque of the propeller thus achieve the target vessel's speed. The input and output are a real time control system of fuel saving rate and propeller rotating speed representing the marine diesel engine characteristics. The self-propulsion tests in calm waters were conducted using a vessel model to validate the marine diesel engine simulator. The simulator then was used to evaluate the fuel saving by employing a new mathematical model of turbochargers for the marine diesel engine simulator. The control system developed will be beneficial for users as to analyze different condition of vessel's speed to obtain better characteristics and hence optimize the fuel saving rate.

Design of a Mars Airplane Propulsion System for the Aerial Regional-Scale Environmental Survey (ARES) Mission Concept

NASA Technical Reports Server (NTRS)

Kuhl, Christopher A.

2008-01-01

The Aerial Regional-Scale Environmental Survey (ARES) is a Mars exploration mission concept that utilizes a rocket propelled airplane to take scientific measurements of atmospheric, surface, and subsurface phenomena. The liquid rocket propulsion system design has matured through several design cycles and trade studies since the inception of the ARES concept in 2002. This paper describes the process of selecting a bipropellant system over other propulsion system options, and provides details on the rocket system design, thrusters, propellant tank and PMD design, propellant isolation, and flow control hardware. The paper also summarizes computer model results of thruster plume interactions and simulated flight performance. The airplane has a 6.25 m wingspan with a total wet mass of 185 kg and has to ability to fly over 600 km through the atmosphere of Mars with 45 kg of MMH / MON3 propellant.

14 CFR 25.1153 - Propeller feathering controls.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Propeller feathering controls. 25.1153... Accessories § 25.1153 Propeller feathering controls. (a) There must be a separate propeller feathering control for each propeller. The control must have means to prevent its inadvertent operation. (b) If...

14 CFR 25.1153 - Propeller feathering controls.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Propeller feathering controls. 25.1153... Accessories § 25.1153 Propeller feathering controls. (a) There must be a separate propeller feathering control for each propeller. The control must have means to prevent its inadvertent operation. (b) If...

14 CFR 25.1153 - Propeller feathering controls.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Propeller feathering controls. 25.1153... Accessories § 25.1153 Propeller feathering controls. (a) There must be a separate propeller feathering control for each propeller. The control must have means to prevent its inadvertent operation. (b) If...

14 CFR 25.1153 - Propeller feathering controls.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Propeller feathering controls. 25.1153... Accessories § 25.1153 Propeller feathering controls. (a) There must be a separate propeller feathering control for each propeller. The control must have means to prevent its inadvertent operation. (b) If...

14 CFR 25.1153 - Propeller feathering controls.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Propeller feathering controls. 25.1153... Accessories § 25.1153 Propeller feathering controls. (a) There must be a separate propeller feathering control for each propeller. The control must have means to prevent its inadvertent operation. (b) If...

14 CFR 25.1149 - Propeller speed and pitch controls.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Propeller speed and pitch controls. 25.1149... Accessories § 25.1149 Propeller speed and pitch controls. (a) There must be a separate propeller speed and... synchronization of all propellers. (d) The propeller speed and pitch controls must be to the right of, and at...

14 CFR 25.1149 - Propeller speed and pitch controls.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Propeller speed and pitch controls. 25.1149... Accessories § 25.1149 Propeller speed and pitch controls. (a) There must be a separate propeller speed and... synchronization of all propellers. (d) The propeller speed and pitch controls must be to the right of, and at...

Investigation of Propellant Sloshing and Zero Gravity Equilibrium for the Orion Service Module Propellant Tanks

NASA Astrophysics Data System (ADS)

Kreppel, Samantha

A scaled model of the downstream Orion service module propellant tank was constructed to asses the propellant dynamics under reduced and zero-gravity conditions. Flight and ground data from the experiment is currently being used to validate computational models of propel-lant dynamics in Orion-class propellant tanks. The high fidelity model includes the internal structures of the propellant management device (PMD) and the mass-gauging probe. Qualita-tive differences between experimental and CFD data are understood in terms of fluid dynamical scaling of inertial effects in the scaled system. Propellant configurations in zero-gravity were studied at a range of fill-fractions and the settling time for various docking maneuvers was determined. A clear understanding of the fluid dynamics within the tank is necessary to en-sure proper control of the spacecraft's flight and to maintain safe operation of this and future service modules. Understanding slosh dynamics in partially-filled propellant tanks is essential to assessing spacecraft stability.

Propellant Savings during Soyuz Undock from the International Space Station

NASA Technical Reports Server (NTRS)

Turett, Fiona

2016-01-01

As a vehicle continuously orbiting Earth for over a decade, the International Space Station (ISS) must be conscious of ways to conserve consumables to maximize the efficiency of cargo flights to ISS. One such consumable is propellant. As part of an ongoing effort to minimize propellant usage onboard ISS and use control moment gyroscopes as much as possible for ISS control, an effort was made in late 2014 to allow Soyuz manned vehicle undockings without requiring the use of thrusters. This method, which has been used for four Soyuz undockings, saves up to 160 kg of propellant each year. Fiona completed a B.S. is Mechanical Engienering at Washington University in St. Louis in 2009, after which she moved to Houston, TX to begin working at NASA Johnson Space Center. She currently works in the Flight Operations Directorate as an ADCO (Attitude Determination and Control Officer) flight controller and MCG (Motion Control Group) instructor. Her responsibilities include operating the motion control systems of the ISS in Mission Control, interfacing with Russian colleagues, mentoring and teaching flight controller trainees, and training astronauts for their missions to ISS.

Iodine Hall Thruster Propellant Feed System for a CubeSat

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.

2014-01-01

There has been significant work recently in the development of iodine-fed Hall thrusters for in-space propulsion applications.1 The use of iodine as a propellant provides many advantages over present xenon-gas-fed Hall thruster systems. Iodine is a solid at ambient temperature (no pressurization required) and has no special handling requirements, making it safe for secondary flight opportunities. It has exceptionally high ?I sp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing system level advantages over mid-term high power electric propulsion options. Iodine provides thrust and efficiency that are comparable to xenonfed Hall thrusters while operating in the same discharge current and voltage regime, making it possible to leverage the development of flight-qualified xenon Hall thruster power processing units for the iodine application. Work at MSFC is presently aimed at designing, integrating, and demonstrating a flight-like iodine feed system suitable for the Hall thruster application. This effort represents a significant advancement in state-of-the-art. Though Iodine thrusters have demonstrated high performance with mission enabling potential, a flight-like feed system has never been demonstrated and iodine compatible components do not yet exist. Presented in this paper is the end-to-end integrated feed system demonstration. The system includes a propellant tank with active feedback-control heating, fill and drain interfaces, latching and proportional flow control valves (PFCV), flow resistors, and flight-like CubeSat power and control electronics. Hardware is integrated into a CubeSat-sized structure, calibrated and tested under vacuum conditions, and operated under under hot-fire conditions using a Busek BHT-200 thruster designed for iodine. Performance of the system is evaluated thorugh accurate measurement of thrust and a calibrated of mass flow rate measurement, which is a function of reservoir temperature/pressure, the flow resistors, and the setting of the PFCV. The calibration is performed using independent flow control monitoring techniques, providing an in situ measure of the flowrate as a function of controllable parameters. The reservoir temperature controls the iodine sublimation rate, providing propellant to ths thruster by pressurizing the propellant feed system to approx.1-2 psi. Control of the temperature and the PFCV are used to maintain reservoir pressure and keep the thruster discharge current constant.

14 CFR 25.1149 - Propeller speed and pitch controls.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Propeller speed and pitch controls. 25.1149... Accessories § 25.1149 Propeller speed and pitch controls. (a) There must be a separate propeller speed and pitch control for each propeller. (b) The controls must be grouped and arranged to allow— (1) Separate...

14 CFR 25.1149 - Propeller speed and pitch controls.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Propeller speed and pitch controls. 25.1149... Accessories § 25.1149 Propeller speed and pitch controls. (a) There must be a separate propeller speed and pitch control for each propeller. (b) The controls must be grouped and arranged to allow— (1) Separate...

14 CFR 25.1149 - Propeller speed and pitch controls.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Propeller speed and pitch controls. 25.1149... Accessories § 25.1149 Propeller speed and pitch controls. (a) There must be a separate propeller speed and pitch control for each propeller. (b) The controls must be grouped and arranged to allow— (1) Separate...

Control of Propellant Lead/Lag to the LAE in the AXAF Propulsion System

NASA Technical Reports Server (NTRS)

Casillas, A. R.; Eninger, J.; Joseph, G.; Kenney, J.; Trinidad, M.

1998-01-01

Control of the rate at which hypergolic propellants are supplied to a rocket engine prior to ignition is critically important. Potentially damaging explosions may result from excessive lead of either propellant into the combustion chamber. Because the injector fill process is governed by the engine as well as the propellant feed system design, proper management of this issue must take both into consideration. This was recognized early in the development of TRW's Advanced Columbium-Liquid Apogee Engine (LAE), which was flight-qualified in 1996 to maneuver the Advanced X-Ray Astrophysics Facility (AXAF) spacecraft into orbit. The LAE runs on hydrazine and nitrogen tetroxide (MON-3) at a nominal mixture ratio of 1.0. This paper describes the comprehensive test program conducted to ensure reliable startup operation of the LAE in the AYAF propulsion system. The most significant factors affecting chamber fuel lead were found to be: (1) engine location, (2) propellant saturation level, (3) amount of undissolved gas in the lines, and (4) off- nominal tank pressures. Hot-fire tests at a chamber fuel lead range over and above that expected for the LAEs in AXAF demonstrated extremely tolerant behavior of the engine. AY-AF is scheduled for launch on NASA's STS-93 in December 1998.

Propulsion Risk Reduction Activities for Non-Toxic Cryogenic Propulsion

NASA Technical Reports Server (NTRS)

Smith, Timothy D.; Klem, Mark D.; Fisher, Kenneth

2010-01-01

The Propulsion and Cryogenics Advanced Development (PCAD) Project s primary objective is to develop propulsion system technologies for non-toxic or "green" propellants. The PCAD project focuses on the development of non-toxic propulsion technologies needed to provide necessary data and relevant experience to support informed decisions on implementation of non-toxic propellants for space missions. Implementation of non-toxic propellants in high performance propulsion systems offers NASA an opportunity to consider other options than current hypergolic propellants. The PCAD Project is emphasizing technology efforts in reaction control system (RCS) thruster designs, ascent main engines (AME), and descent main engines (DME). PCAD has a series of tasks and contracts to conduct risk reduction and/or retirement activities to demonstrate that non-toxic cryogenic propellants can be a feasible option for space missions. Work has focused on 1) reducing the risk of liquid oxygen/liquid methane ignition, demonstrating the key enabling technologies, and validating performance levels for reaction control engines for use on descent and ascent stages; 2) demonstrating the key enabling technologies and validating performance levels for liquid oxygen/liquid methane ascent engines; and 3) demonstrating the key enabling technologies and validating performance levels for deep throttling liquid oxygen/liquid hydrogen descent engines. The progress of these risk reduction and/or retirement activities will be presented.

Propulsion Risk Reduction Activities for Nontoxic Cryogenic Propulsion

NASA Technical Reports Server (NTRS)

Smith, Timothy D.; Klem, Mark D.; Fisher, Kenneth L.

2010-01-01

The Propulsion and Cryogenics Advanced Development (PCAD) Project s primary objective is to develop propulsion system technologies for nontoxic or "green" propellants. The PCAD project focuses on the development of nontoxic propulsion technologies needed to provide necessary data and relevant experience to support informed decisions on implementation of nontoxic propellants for space missions. Implementation of nontoxic propellants in high performance propulsion systems offers NASA an opportunity to consider other options than current hypergolic propellants. The PCAD Project is emphasizing technology efforts in reaction control system (RCS) thruster designs, ascent main engines (AME), and descent main engines (DME). PCAD has a series of tasks and contracts to conduct risk reduction and/or retirement activities to demonstrate that nontoxic cryogenic propellants can be a feasible option for space missions. Work has focused on 1) reducing the risk of liquid oxygen/liquid methane ignition, demonstrating the key enabling technologies, and validating performance levels for reaction control engines for use on descent and ascent stages; 2) demonstrating the key enabling technologies and validating performance levels for liquid oxygen/liquid methane ascent engines; and 3) demonstrating the key enabling technologies and validating performance levels for deep throttling liquid oxygen/liquid hydrogen descent engines. The progress of these risk reduction and/or retirement activities will be presented.

Propulsion and Cryogenics Advanced Development (PCAD) Project Propulsion Technologies for the Lunar Lander

NASA Technical Reports Server (NTRS)

Klem, Mark D.; Smith, Timothy D.

2008-01-01

The Propulsion and Cryogenics Advanced Development (PCAD) Project in the Exploration Technology Development Program is developing technologies as risk mitigation for Orion and the Lunar Lander. An integrated main and reaction control propulsion system has been identified as a candidate for the Lunar Lander Ascent Module. The propellants used in this integrated system are Liquid Oxygen (LOX)/Liquid Methane (LCH4) propellants. A deep throttle pump fed Liquid Oxygen (LOX)/Liquid Hydrogen (LH2) engine system has been identified for the Lunar Lander Descent Vehicle. The propellant combination and architecture of these propulsion systems are novel and would require risk reduction prior to detailed design and development. The PCAD Project addresses the technology requirements to obtain relevant and necessary test data to further the technology maturity of propulsion hardware utilizing these propellants. This plan and achievements to date will be presented.

Initial steps toward automation of a propellant processor

NASA Technical Reports Server (NTRS)

Schallhorn, Paul; Ramohalli, Kumar

1990-01-01

This paper presents the results from an experimental study aimed at ultimately automating the mixing of propellants in order to minimize unintended variations usually attributed to human error. The water heater and delivery system of a one-pint Baker-Perkins (APV) vertical mixer are automated with computer control. Various innovations are employed to introduce economy and low thermal inertia. Some of these include twin heaters/reservoirs instead of one large reservoir, a compact water mixer for achieving the desired temperature quickly, and thorough insulation of the entire water system. The completed system is tested during two propellant mixes. The temperature uniformly is proven through careful measurements employing several local thermocouples.

Liquid Bismuth Feed System for Electric Propulsion

NASA Technical Reports Server (NTRS)

Markusic, T. E.; Polzin, K. A.; Stanojev, B. J.

2006-01-01

Operation of Hall thrusters with bismuth propellant has been shown to be a promising path toward high-power, high-performance, long-lifetime electric propulsion for spaceflight missions. For example, the VHITAL project aims td accurately, experimentally assess the performance characteristics of 10 kW-class bismuth-fed Hall thrusters - in order to validate earlier results and resuscitate a promising technology that has been relatively dormant for about two decades. A critical element of these tests will be the precise metering of propellant to the thruster, since performance cannot be accurately assessed without an accurate accounting of mass flow rate. Earlier work used a pre/post-test propellant weighing scheme that did not provide any real-time measurement of mass flow rate while the thruster was firing, and makes subsequent performance calculations difficult. The motivation of the present work was to develop a precision liquid bismuth Propellant Management System (PMS) that provides real-time propellant mass flow rate measurement and control, enabling accurate thruster performance measurements. Additionally, our approach emphasizes the development of new liquid metal flow control components and, hence, will establish a basis for the future development of components for application in spaceflight. The design of various critical components in a bismuth PMS are described - reservoir, electromagnetic pump, hotspot flow sensor, and automated control system. Particular emphasis is given to material selection and high-temperature sealing techniques. Open loop calibration test results are reported, which validate the systems capability to deliver bismuth at mass flow rates ranging from 10 to 100 mg/sec with an uncertainty of less than +/- 5%. Results of integrated vaporizer/liquid PMS tests demonstrate all of the necessary elements of a complete bismuth feed system for electric propulsion.

46 CFR 154.904 - Inert gas system: Controls.

Code of Federal Regulations, 2011 CFR

2011-10-01

... 46 Shipping 5 2011-10-01 2011-10-01 false Inert gas system: Controls. 154.904 Section 154.904... STANDARDS FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design, Construction and Equipment Atmospheric Control in Cargo Containment Systems § 154.904 Inert gas system: Controls. The inert gas system...

Cryogenics and the Human Exploration of Mars

NASA Technical Reports Server (NTRS)

Salerno, Louis J.; Kittel, Peter; Rasky, Daniel J. (Technical Monitor)

1997-01-01

Current plans within NASA involve extending the human exploration of space from low earth orbit into the solar system, with the first human exploration of Mars presently planned in 2011. Integral to all hum Mars mission phases is cryogenic fluid management. Cryogenic fluids will be required both as propellant and for In-Situ Resource Utilization (ISRU). Without safe and efficient cryogen storage human Mars missions will not be possible. Effective control and handling of cryogenic fluids is the key to affordable Mars missions, and advancing active thermal control technology is synergistic with all of NASA's exploration initiatives and with existing and future instrument cooling programs, including MTPE and Origins. Present mission scenarios for human exploration require cryogenic propellant storage for up to 1700 days and for up to 60 metric tons. These requirements represent increases of an order of magnitude over previous storage masses and lifetimes. The key cryogenic terminology areas to be addressed in human Mars missions are long-term propellant storage, cryogenic refrigeration, cryogenic liquefaction, and zero gravity fluid management. Long-term storage for the thermal control of cryogenic propellants is best accomplished with a mix of passive and active technologies. Passive technologies such as advanced multilayer insulation (MLI) concepts will be combined with the development of active coolers (cryogenic refrigerators). Candidates for long-life active cooling applications include Reverse Turbo-Brayton, Stirling, and Pulse-Tube coolers. The integration of passive and active technologies will form a hybrid system optimized to minimize the launch mass while preserving the cryogenic propellants. Since cryogenic propellants are the largest mass that Mars missions must launch from earth, even a modest reduction in the percentage of propellant carried results in a significant weight saving. This paper will present a brief overview of cryogenic fluid management technology as it applies to the current human Mars mission scenarios.

Zero boil-off system testing

NASA Astrophysics Data System (ADS)

Plachta, D. W.; Johnson, W. L.; Feller, J. R.

2016-03-01

Cryogenic propellants such as liquid hydrogen (LH2) and liquid oxygen (LO2) are a part of NASA's future space exploration plans due to their high specific impulse for rocket motors of upper stages. However, the low storage temperatures of LH2 and LO2 cause substantial boil-off losses for long duration missions. These losses can be eliminated by incorporating high performance cryocooler technology to intercept heat load to the propellant tanks and modulating the cryocooler temperature to control tank pressure. The technology being developed by NASA is the reverse turbo-Brayton cycle cryocooler and its integration to the propellant tank through a distributed cooling tubing network coupled to the tank wall. This configuration was recently tested at NASA Glenn Research Center in a vacuum chamber and cryoshroud that simulated the essential thermal aspects of low Earth orbit, its vacuum and temperature. This test series established that the active cooling system integrated with the propellant tank eliminated boil-off and robustly controlled tank pressure.

Zero Boil-Off System Testing

NASA Technical Reports Server (NTRS)

Plachta, D. W.; Johnson, W. L.; Feller, J. R.

2015-01-01

Cryogenic propellants such as liquid hydrogen (LH2) and liquid oxygen (LO2) are a part of NASA's future space exploration plans due to their high specific impulse for rocket motors of upper stages. However, the low storage temperatures of LH2 and LO2 cause substantial boil-off losses for long duration missions. These losses can be eliminated by incorporating high performance cryocooler technology to intercept heat load to the propellant tanks and modulating the cryocooler temperature to control tank pressure. The technology being developed by NASA is the reverse turbo-Brayton cycle cryocooler and its integration to the propellant tank through a distributed cooling tubing network coupled to the tank wall. This configuration was recently tested at NASA Glenn Research Center in a vacuum chamber and cryoshroud that simulated the essential thermal aspects of low Earth orbit, its vacuum and temperature. This test series established that the active cooling system integrated with the propellant tank eliminated boil-off and robustly controlled tank pressure.

Software For Graphical Representation Of A Network

NASA Technical Reports Server (NTRS)

Mcallister, R. William; Mclellan, James P.

1993-01-01

System Visualization Tool (SVT) computer program developed to provide systems engineers with means of graphically representing networks. Generates diagrams illustrating structures and states of networks defined by users. Provides systems engineers powerful tool simplifing analysis of requirements and testing and maintenance of complex software-controlled systems. Employs visual models supporting analysis of chronological sequences of requirements, simulation data, and related software functions. Applied to pneumatic, hydraulic, and propellant-distribution networks. Used to define and view arbitrary configurations of such major hardware components of system as propellant tanks, valves, propellant lines, and engines. Also graphically displays status of each component. Advantage of SVT: utilizes visual cues to represent configuration of each component within network. Written in Turbo Pascal(R), version 5.0.

Design and testing of the Space Station Freedom Propellant Tank Assembly

NASA Technical Reports Server (NTRS)

Dudley, D. D.; Thonet, T. A.; Goforth, A. M.

1992-01-01

Propellant storage and management functions for the Propulsion Module of the U.S. Space Station Freedom are provided by the Propellant Tank Assembly (PTA). The PTA consists of a surface-tension type propellant acquisition device contained within a welded titanium pressure vessel. The PTA design concept was selected with high reliability and low program risk as primary goals in order to meet stringent NASA structural, expulsion, fracture control and reliability requirements. The PTA design makes use of Shuttle Orbital Maneuvering System and Peacekeeper Propellant Storage Assembly design and analysis techniques. This paper summarizes the PTA design solution and discusses the underlying detailed analyses. In addition, design verification and qualification test activities are discussed.

The Iodine Satellite (iSat) Propellant Feed System - Design and Demonstration

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Seixal, Joao F.; Mauro, Stephanie; Burt, Adam O.; Martinez, Armando; Peeples, Steven R.

2017-01-01

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cm3 and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high Delta V maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. One of the systems under development to support such a technology is the propellant feed system, which must be capable of storing solid iodine propellant, applying heat to sublime the stored solid into the vapor phase, and then control the flow of low-pressure gaseous iodine to both the thruster and cathode. In a test conducted in 2016, a first-generation iodine propellant feed system was integrated with a cathode and Hall thruster. While this test had to be terminated, the feed system in this first test was able to support both cathode and integrated cathode and thruster operation prior to the termination of the test. In the present paper, we describe work performed since that initial integrated test. The effort uses lessons learned from the previous integrated test, retiring risk associated with the iodine propellant feed system, answering open design-space questions, and demonstrating iodine flow control in an integrated system. The work is undertaken at both the component level and then at the integrated subsystem level to systematically improve the feed system design, improving the hardware fidelity so the appearance and operation of the system are as flight-like as possible. At the component level, the work focuses on the propellant tank, the feed system tubing, the valves used to control the flow to the cathode and thruster, and the heaters that maintain the temperature of the flowpaths and keep iodine from redepositing and clogging the system. Work on the propellant reservoir focuses on fabricating a tank that matches the geometry of the flight design, which allows for the identification of flight tank fabrication issues that may arise and permits thermal testing of a tank possessing the same size and thermal mass as the flight design, which can be used to anchor thermal modeling of the component. This is critical for finalizing the tank heater power requirements that feed into the heater design. All metallic materials in the feed system are hastelloy or Inconel, as these materials are resistant to chemical attack by the highly-reactive iodine vapor. The tubing in the iodine feed system must possess ports to permit a neutral gas purge of the system that clear impurities after iodine is loaded into the propellant tank. A procedure is discussed whereby these ports are crimped and sealed after the purge process is completed so as to not re-expose the iodine system to air. The valves are a critical component for control of the flow to the thruster and the cathode. Significant effort has gone into upgrading the materials of the valves to make them more resistant to chemical attack and into developing an understanding of the use of these valves during the startup and operation of the cathode and thruster. The heaters that line the entire feed system are designed to draw minimal power from the power processing unit (PPU) while still having the capacity to maintain all the feed system components at the temperatures required to discourage iodine deposition inside components downstream of the propellant tank exit. The heaters possess two separate resistive traces, giving the design redundancy should a failure occur in the primary heater circuit of one of the heater zones. The task of operating a feed system in conjunction with a thruster and cathode is undertaken in a series of sub-steps. The system is first assembled and operated on xenon gas, using the valves for cathode startup and thruster control based on measurement of the discharge current. After startup and control on xenon are demonstrated, the thruster will be transitioned to iodine operation, demonstrating thruster startup and feed system control while using a xenon-fed cathode. Finally, the last step is to integrate an iodine-compatible cathode with the system, demonstrate autonomous cathode start-up with open-loop control and thruster start-up with closed-loop control for multiple cycles.

Multimode marine engine room simulation system based on field bus technology

NASA Astrophysics Data System (ADS)

Zheng, Huayao; Deng, Linlin; Guo, Yi

2003-09-01

Developing multi mode MER (Marine Engine Room) Labs is the main work in Marine Simulation Center, which is the key lab of Communication Ministry of China. It includes FPP (Fixed Pitch Propeller) and CPP (Controllable Pitch Propeller) mode MER simulation systems, integrated electrical propulsion mode MER simulation system, physical mode MER lab, etc. FPP mode simulation system, which was oriented to large container ship, had been completed since 1999, and got second level of Shanghai Municipal Science and Technical Progress award. This paper mainly introduces the recent development and achievements of Marine Simulation Center. Based on the Lon Works field bus, the structure characteristics and control strategies of completely distributed intelligent control network are discussed. The experiment mode of multi-nodes field bus detection and control system is described. Besides, intelligent fault diagnosis technology about some mechatronics integration control systems explored is also involved.

A method of calculating the performance of controllable propellers with sample computations

NASA Technical Reports Server (NTRS)

Hartman, Edwin P

1934-01-01

This paper contains a series of calculations showing how the performance of controllable propellers may be derived from data on fixed-pitch propellers given in N.A.C.A. Technical Report No. 350, or from similar data. Sample calculations are given which compare the performance of airplanes with fixed-pitch and with controllable propellers. The gain in performance with controllable propellers is shown to be largely due to the increased power available, rather than to an increase in efficiency. Controllable propellers are of particular advantage when used with geared and with supercharged engines. A controllable propeller reduces the take-off run, increases the rate of climb and the ceiling, but does not increase the high speed, except when operating above the design altitude of the previously used fixed-pitch propeller or when that propeller was designed for other than high speed.

Optimal Propellant Maneuver Flight Demonstrations on ISS

NASA Technical Reports Server (NTRS)

Bhatt, Sagar; Bedrossian, Nazareth; Longacre, Kenneth; Nguyen, Louis

2013-01-01

In this paper, first ever flight demonstrations of Optimal Propellant Maneuver (OPM), a method of propulsive rotational state transition for spacecraft controlled using thrusters, is presented for the International Space Station (ISS). On August 1, 2012, two ISS reorientations of about 180deg each were performed using OPMs. These maneuvers were in preparation for the same-day launch and rendezvous of a Progress vehicle, also a first for ISS visiting vehicles. The first maneuver used 9.7 kg of propellant, whereas the second used 10.2 kg. Identical maneuvers performed without using OPMs would have used approximately 151.1kg and 150.9kg respectively. The OPM method is to use a pre-planned attitude command trajectory to accomplish a rotational state transition. The trajectory is designed to take advantage of the complete nonlinear system dynamics. The trajectory choice directly influences the cost of the maneuver, in this case, propellant. For example, while an eigenaxis maneuver is kinematically the shortest path between two orientations, following that path requires overcoming the nonlinear system dynamics, thereby increasing the cost of the maneuver. The eigenaxis path is used for ISS maneuvers using thrusters. By considering a longer angular path, the path dependence of the system dynamics can be exploited to reduce the cost. The benefits of OPM for the ISS include not only reduced lifetime propellant use, but also reduced loads, erosion, and contamination from thrusters due to fewer firings. Another advantage of the OPM is that it does not require ISS flight software modifications since it is a set of commands tailored to the specific attitude control architecture. The OPM takes advantage of the existing ISS control system architecture for propulsive rotation called USTO control mode1. USTO was originally developed to provide ISS Orbiter stack attitude control capability for a contingency tile-repair scenario, where the Orbiter is maneuvered using its robotic manipulator relative to the ISS. Since 2005 USTO has been used for nominal ISS operations.

Mathematical model of marine diesel engine simulator for a new methodology of self propulsion tests

DOE Office of Scientific and Technical Information (OSTI.GOV)

Izzuddin, Nur; Sunarsih,; Priyanto, Agoes

As a vessel operates in the open seas, a marine diesel engine simulator whose engine rotation is controlled to transmit through propeller shaft is a new methodology for the self propulsion tests to track the fuel saving in a real time. Considering the circumstance, this paper presents the real time of marine diesel engine simulator system to track the real performance of a ship through a computer-simulated model. A mathematical model of marine diesel engine and the propeller are used in the simulation to estimate fuel rate, engine rotating speed, thrust and torque of the propeller thus achieve the targetmore » vessel’s speed. The input and output are a real time control system of fuel saving rate and propeller rotating speed representing the marine diesel engine characteristics. The self-propulsion tests in calm waters were conducted using a vessel model to validate the marine diesel engine simulator. The simulator then was used to evaluate the fuel saving by employing a new mathematical model of turbochargers for the marine diesel engine simulator. The control system developed will be beneficial for users as to analyze different condition of vessel’s speed to obtain better characteristics and hence optimize the fuel saving rate.« less

Extended temperature range ACPS thruster investigation

NASA Technical Reports Server (NTRS)

Blubaugh, A. L.; Schoenman, L.

1974-01-01

The successful hot fire demonstration of a pulsing liquid hydrogen/liquid oxygen and gaseous hydrogen/liquid oxygen attitude control propulsion system thruster is described. The test was the result of research to develop a simple, lightweight, and high performance reaction control system without the traditional requirements for extensive periods of engine thermal conditioning, or the use of complex equipment to convert both liquid propellants to gas prior to delivery to the engine. Significant departures from conventional injector design practice were employed to achieve an operable design. The work discussed includes thermal and injector manifold priming analyses, subscale injector chilldown tests, and 168 full scale and 550 N (1250 lbF) rocket engine tests. Ignition experiments, at propellant temperatures ranging from cryogenic to ambient, led to the generation of a universal spark ignition system which can reliably ignite an engine when supplied with liquid, two phase, or gaseous propellants. Electrical power requirements for spark igniter are very low.

77 FR 42677 - Special Conditions: General Electric CT7-2E1 Turboshaft Engine

Federal Register 2010, 2011, 2012, 2013, 2014

2012-07-20

... Mihail, ANE-111, Engine and Propeller Directorate, Aircraft Certification Service, 12 New England..., ANE-7 Engine and Propeller Directorate, Aircraft Certification Service, 12 New England Executive Park... additional requirements for the rating's definition, overspeed, controls system, and endurance test because...

Differential Draining of Parallel-Fed Propellant Tanks in Morpheus and Apollo Flight

NASA Technical Reports Server (NTRS)

Hurlbert, Eric; Guardado, Hector; Hernandez, Humberto; Desai, Pooja

2015-01-01

Parallel-fed propellant tanks are an advantageous configuration for many spacecraft. Parallel-fed tanks allow the center of gravity (cg) to be maintained over the engine(s), as opposed to serial-fed propellant tanks which result in a cg shift as propellants are drained from tank one tank first opposite another. Parallel-fed tanks also allow for tank isolation if that is needed. Parallel tanks and feed systems have been used in several past vehicles including the Apollo Lunar Module. The design of the feedsystem connecting the parallel tank is critical to maintain balance in the propellant tanks. The design must account for and minimize the effect of manufacturing variations that could cause delta-p or mass flowrate differences, which would lead to propellant imbalance. Other sources of differential draining will be discussed. Fortunately, physics provides some self-correcting behaviors that tend to equalize any initial imbalance. The question concerning whether or not active control of propellant in each tank is required or can be avoided or not is also important to answer. In order to provide data on parallel-fed tanks and differential draining in flight for cryogenic propellants (as well as any other fluid), a vertical test bed (flying lander) for terrestrial use was employed. The Morpheus vertical test bed is a parallel-fed propellant tank system that uses passive design to keep the propellant tanks balanced. The system is operated in blow down. The Morpheus vehicle was instrumented with a capacitance level sensor in each propellant tank in order to measure the draining of propellants in over 34 tethered and 12 free flights. Morpheus did experience an approximately 20 lb/m imbalance in one pair of tanks. The cause of this imbalance will be discussed. This paper discusses the analysis, design, flight simulation vehicle dynamic modeling, and flight test of the Morpheus parallel-fed propellant. The Apollo LEM data is also examined in this summary report of the flight data.

Liquid Bismuth Propellant Management System for the Very High Specific Impulse Thruster with Anode Layer

NASA Technical Reports Server (NTRS)

Polzin, K. A.; Markusic, T. E.; Stanojev, B. J.

2007-01-01

Two prototype bismuth propellant feed systems were constructed and operated in conjunction with a propellant vaporizer. One system provided bismuth to a vaporizer using gas pressurization but did not include a means to measure the flow rate. The second system incorporated an electromagnetic pump to provide fine control of the hydrostatic pressure and a new type of in-line flow sensor that was developed for accurate, real-time measurement of the mass flow rate. High-temperature material compatibility was a driving design requirement for the pump and flow sensor, leading to the selection of Macor for the main body of both components. Posttest inspections of both components revealed no degradation of the material. The gas pressurization system demonstrated continuous pressure control over a range from zero to 200 torr. In separate proof-of-concept experiments, the electromagnetic pump produced a linear pressure rise as a function of current that compared favorably with theoretical pump pressure predictions, producing a pressure rise of 10 kPa at 30 A. Preliminary flow sensor operation indicated a bismuth flow rate of 6 mg/s with an uncertainty of plus or minus 6%. An electronics suite containing a real-time controller was successfully used to control the entire system, simultaneously monitoring all power supplies and performing data acquisition duties.

Ultrasonic propulsion

NASA Astrophysics Data System (ADS)

Allison, Eric

In this investigation, a propulsion system is introduced for propelling and guiding an object through a fluid. Thrust for forward motion and for turning is produced by acoustic waves generated by piezoelectric ultrasonic transducers. The principle of operation of the transducers is described, and methods are presented for the design of the entire system, including the transducers, signal generator, guidance and control system, and the power source. A wirelessly controlled proof-of-concept device was constructed. This device demonstrates the operation and practicality of the propulsion and guidance systems and illustrates that they may be employed in situations where the use of conventional propulsive devices such as propellers or jets is unfeasible.

Active local control of propeller-aircraft run-up noise.

PubMed

Hodgson, Murray; Guo, Jingnan; Germain, Pierre

2003-12-01

Engine run-ups are part of the regular maintenance schedule at Vancouver International Airport. The noise generated by the run-ups propagates into neighboring communities, disturbing the residents. Active noise control is a potentially cost-effective alternative to passive methods, such as enclosures. Propeller aircraft generate low-frequency tonal noise that is highly compatible with active control. This paper presents a preliminary investigation of the feasibility and effectiveness of controlling run-up noise from propeller aircraft using local active control. Computer simulations for different configurations of multi-channel active-noise-control systems, aimed at reducing run-up noise in adjacent residential areas using a local-control strategy, were performed. These were based on an optimal configuration of a single-channel control system studied previously. The variations of the attenuation and amplification zones with the number of control channels, and with source/control-system geometry, were studied. Here, the aircraft was modeled using one or two sources, with monopole or multipole radiation patterns. Both free-field and half-space conditions were considered: for the configurations studied, results were similar in the two cases. In both cases, large triangular quiet zones, with local attenuations of 10 dB or more, were obtained when nine or more control channels were used. Increases of noise were predicted outside of these areas, but these were minimized as more control channels were employed. By combining predicted attenuations with measured noise spectra, noise levels after implementation of an active control system were estimated.

Cryogenic Propellant Storage and Transfer Technology Demonstration For Long Duration In-Space Missions

NASA Technical Reports Server (NTRS)

Meyer, Michael L.; Motil, Susan M.; Kortes, Trudy F.; Taylor, William J.; McRight, Patrick S.

2012-01-01

The high specific impulse of cryogenic propellants can provide a significant performance advantage for in-space transfer vehicles. The upper stages of the Saturn V and various commercial expendable launch vehicles have used liquid oxygen and liquid hydrogen propellants; however, the application of cryogenic propellants has been limited to relatively short duration missions due to the propensity of cryogens to absorb environmental heat resulting in fluid losses. Utilizing advanced cryogenic propellant technologies can enable the efficient use of high performance propellants for long duration missions. Crewed mission architectures for beyond low Earth orbit exploration can significantly benefit from this capability by developing realistic launch spacing for multiple launch missions, by prepositioning stages and by staging propellants at an in-space depot. The National Aeronautics and Space Administration through the Office of the Chief Technologist is formulating a Cryogenic Propellant Storage and Transfer Technology Demonstration Mission to mitigate the technical and programmatic risks of infusing these advanced technologies into the development of future cryogenic propellant stages or in-space propellant depots. NASA is seeking an innovative path for human space exploration, which strengthens the capability to extend human and robotic presence throughout the solar system. This mission will test and validate key cryogenic technological capabilities and has the objectives of demonstrating advanced thermal control technologies to minimize propellant loss during loiter, demonstrating robust operation in a microgravity environment, and demonstrating efficient propellant transfer on orbit. The status of the demonstration mission concept development, technology demonstration planning and technology maturation activities in preparation for flight system development are described.

Electromagnetic Formation Flight (EMFF) for Sparse Aperture Arrays

NASA Technical Reports Server (NTRS)

Kwon, Daniel W.; Miller, David W.; Sedwick, Raymond J.

2004-01-01

Traditional methods of actuating spacecraft in sparse aperture arrays use propellant as a reaction mass. For formation flying systems, propellant becomes a critical consumable which can be quickly exhausted while maintaining relative orientation. Additional problems posed by propellant include optical contamination, plume impingement, thermal emission, and vibration excitation. For these missions where control of relative degrees of freedom is important, we consider using a system of electromagnets, in concert with reaction wheels, to replace the consumables. Electromagnetic Formation Flight sparse apertures, powered by solar energy, are designed differently from traditional propulsion systems, which are based on V. This paper investigates the design of sparse apertures both inside and outside the Earth's gravity field.

14 CFR 23.1149 - Propeller speed and pitch controls.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Propeller speed and pitch controls. 23.1149... Powerplant Controls and Accessories § 23.1149 Propeller speed and pitch controls. (a) If there are propeller speed or pitch controls, they must be grouped and arranged to allow— (1) Separate control of each...

14 CFR 23.1149 - Propeller speed and pitch controls.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Propeller speed and pitch controls. 23.1149... Powerplant Controls and Accessories § 23.1149 Propeller speed and pitch controls. (a) If there are propeller speed or pitch controls, they must be grouped and arranged to allow— (1) Separate control of each...

Combustion performance and heat transfer characterization of LOX/hydrocarbon type propellants

NASA Technical Reports Server (NTRS)

Gross, R. S.

1980-01-01

A sound data base was established by analytically and experimentally generating basic regenerative cooling, combustion performance, combustion stability, and combustion chamber heat transfer parameters for LOX/HC propellants, with specific application to second generation orbit maneuvering and reaction control systems (OMS/RCS) for the Space Shuttle Orbiter.

Development of HAN-based Liquid Propellant Thruster

NASA Astrophysics Data System (ADS)

Hisatsune, K.; Izumi, J.; Tsutaya, H.; Furukawa, K.

2004-10-01

Many of propellants that are applied to the conventional spacecraft propulsion system are toxic propellants. Because of its toxicity, considering the environmental pollution or safety on handling, it will be necessary to apply the "green" propellant to the spacecraft propulsion system. The purpose of this study is to apply HAN based liquid propellant (LP1846) to mono propellant thruster. Compared to the hydrazine that is used in conventional mono propellant thruster, HAN based propellant is not only lower toxic but also can obtain higher specific impulse. Moreover, HAN based propellant can be decomposed by the catalyst. It means there are the possibility of applying to the mono propellant thruster that can leads to the high reliability of the propulsion system.[1],[2] However, there are two technical subjects, to apply HAN based propellant to the mono propellant thruster. One is the high combustion temperature. The catalyst will be damaged under high temperature condition. The other is the low catalytic activity. It is the serious problem on application of HAN based propellant to the mono propellant thruster that is used for attitude control of spacecraft. To improve the catalytic activity of HAN based propellant, it is necessary to screen the best catalyst for HAN based propellant. The adsorption analysis is conducted by Monte Carlo Simulation to screen the catalyst metal for HAN and TEAN. The result of analysis shows the Iridium is the best catalyst metal for HAN and TEAN. Iridium is the catalyst metal that is used at conventional mono propellant thruster catalyst Shell405. Then, to confirm the result of analysis, the reaction test about catalyst is conducted. The result of this test is the same as the result of adsorption analysis. That means the adsorption analysis is effective in screening the catalyst metal. At the evaluating test, the various types of carrier of catalyst are also compared to Shell 405 to improve catalytic activity. The test result shows the inorganic porous media is superior to Shell405 in catalytic activity. Next, the catalyst life with HAN based propellant (LP1846) is evaluated. The Shell405 and inorganic porous media catalyst are compared at the life test. The test result shows the inorganic porous media catalyst is superior to Shell405 in catalyst life. In this paper, the detail of the result of adsorption analysis and evaluating test are reported.

Plasma characterization for application in ballistics

DOE Office of Scientific and Technical Information (OSTI.GOV)

Katulka, G.; Nusca, M.; White, K.

1996-12-31

There is currently a strong motivation for improving the existing performance of fielded military gun systems. For that objective, research over the past several years has been carried out in an effort to enhance performance by addition of energy into the gun chamber by way of a plasma generator. This energy addition, referred to as Electro-thermal Chemical (ETC) propulsion, can be readily controlled electrically where it can be used to ignite the chamber`s energetic material, enhance the total energy, and control the interior process through control of the propellant combustion. To realize the potential advantages of this system it ismore » important to characterize the plasma generator in terms of (a) the impedance characteristics and its relationship to the pulse forming network used to generate the plasma, (b) the plasma output energy components such as radiation and convection in both time and space, (c) the details of the hydrodynamic interactions of the plasma with the propelling charge bed in the gun chamber and, (d) the direct effect of the plasma on the propellant reactions. Experimental studies have been carried out to study the effect of the plasma radiation on the propellant characteristics related to combustion.« less

Design, fabrication and test of a liquid hydrogen titanium honeycomb cryogenic test tank for use as a reusable launch vehicle main propellant tank

NASA Astrophysics Data System (ADS)

Stickler, Patrick B.; Keller, Peter C.

1998-01-01

Reusable Launch Vehicles (RLV's) utilizing LOX\\LH2 as the propellant require lightweight durable structural systems to meet mass fraction goals and to reduce overall systems operating costs. Titanium honeycomb sandwich with flexible blanket TPS on the windward surface is potentially the lightest-weight and most operable option. Light weight is achieved in part because the honeycomb sandwich tank provides insulation to its liquid hydrogen contents, with no need for separate cryogenic insulation, and in part because the high use temperature of titanium honeycomb reduces the required surface area of re-entry thermal protection systems. System operability is increased because TPS needs to be applied only to surfaces where temperatures exceed approximately 650 K. In order to demonstrate the viability of a titanium sandwich constructed propellant tank, a technology demonstration program was conducted including the design, fabrication and testing of a propellant tank-TPS system. The tank was tested in controlled as well as ambient environments representing ground hold conditions for a RLV main propellant tank. Data collected during each test run was used to validate predictions for air liquefaction, outside wall temperature, boil-off rates, frost buildup and its insulation effects, and the effects of placing a thermal protection system blanket on the external surface. Test results indicated that titanium honeycomb, when used as a RLV propellant tank material, has great promise as a light-weight structural system.

Advanced Ground Systems Maintenance Cryogenics Test Lab Control System Upgrade Project

NASA Technical Reports Server (NTRS)

Harp, Janice Leshay

2014-01-01

This project will outfit the Simulated Propellant Loading System (SPLS) at KSC's Cryogenics Test Laboratory with a new programmable logic control system. The control system upgrade enables the Advanced Ground Systems Maintenace Element Integration Team and other users of the SPLS to conduct testing in a controls environment similar to that used at the launch pad.

SMS engineering design report

NASA Technical Reports Server (NTRS)

1976-01-01

The engineering design for the Shuttle Missions Simulator is presented in sections, with each section representing a subsystem development activity. Subsystems covered include: electrical power system; mechanical power system; main propellant and external tank; solid rocket booster; reaction control system; orbital maneuvering system; guidance, navigation, and control; data processing system; mission control center interface; and image display system.

Experiments on reduction of propeller induced interior noise by active control of cylinder vibration

NASA Technical Reports Server (NTRS)

Fuller, C. R.; Jones, J. D.

1987-01-01

The feasibility of reducing interior noise caused by advanced turbo propellers by controlling the vibration of aircraft fuselages was investigated by performing experiments in an anechoic chamber with an aircraft model test rig and apparatus. It was found that active vibration control provides reasonable global attenuation of interior noise levels for the cases of resonant (at 576 Hz) and forced (at 708 Hz) system response. The controlling mechanism behind the effect is structural-acoustic coupling between the shell and the contained field, termed interface modal filtering.

Development of a Passively Varying Pitch Propeller

NASA Astrophysics Data System (ADS)

Heinzen, Stearns Beamon

Small general aviation aircraft and unmanned aerial systems are often equipped with sophisticated navigation, control, and other avionics, but retain propulsion systems consisting of retrofitted radio control and ultralight equipment. Consequently, new high performance airframes often rely on relatively primitive propulsive technology. This trend is beginning to shift with recent advances in small turboprop engines, fuel injected reciprocating engines, and improved electric technologies. Although these systems are technologically advanced, they are often paired with standard fixed pitch propellers. To fully realize the potential of these aircraft and the new generation of engines, small propellers which can efficiently transmit power over wide flight envelopes and a variety of power settings must be developed. This work demonstrates a propeller which passively adjusts to incoming airflow at a low penalty to aircraft weight and complexity. This allows the propeller to operate in an efficient configuration over a wide flight envelope, and can prevent blade stall in low-velocity / highly-loaded thrust cases and over-speeding at high flight speeds. The propeller incorporates blades which pivot freely on a radial axis and are aerodynamically tailored to attain and maintain a pitch angle yielding favorable local blade angles of attack, matched to changing inflow conditions. This blade angle is achieved through the use of reflexed airfoils designed for a positive pitching moment, comparable to those used on many tailless flying wings. By setting the axis of rotation at a point forward of the blade aerodynamic center, the blades will naturally adjust to a predetermined positive lift 'trim' condition. Then, as inflow conditions change, the blade angle will automatically pivot to maintain the same angle with respect to incoming air. Computational, wind tunnel, and flight test results indicate that the extent of efficient propeller operation can be increased dramatically as compared to the fixed pitch propellers currently used on most light aircraft and small unmanned systems, making significant improvements in aircraft performance possible. These improvements may yield aircraft with reduced takeoff distances, improved climb rates, increased range and endurance, and higher top speeds, without sacrificing on-design performance.

Extending the life and recycle capability of earth storable propellant systems.

NASA Technical Reports Server (NTRS)

Schweickert, T. F.

1972-01-01

Rocket propulsion systems for reusable vehicles will be required to operate reliably for a large number of missions with a minimum of maintenance and a fast turnaround. For the space shuttle reaction control system to meet these requirements, current and prior related system failures were examined for their impact on reuse and, where warranted, component design and/or system configuration changes were defined for improving system service life. It was found necessary to change the pressurization component arrangement used on many single-use applications in order to eliminate a prevalent check valve failure mode and to incorporate redundant expulsion capability in propellant tank designs to achieve the necessary system reliability. Material flaws in pressurant and propellant tanks were noted to have a significant effect on tank cycle life. Finally, maintenance considerations dictated a modularized systems approach, allowing the system to be removed from the vehicle for service and repair at a remote site.

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

NASA Technical Reports Server (NTRS)

Plachta, David W.; Tucker, Stephen; Hoffman, David J.

1993-01-01

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.

PC programs for the prediction of the linear stability behavior of liquid propellant propulsion systems and application to current MSFC rocket engine test programs, volume 1

NASA Technical Reports Server (NTRS)

Doane, George B., III; Armstrong, W. C.

1990-01-01

Research on propulsion stability (chugging and acoustic modes), and propellant valve control was investigated. As part of the activation of the new liquid propulsion test facilities, it is necessary to analyze total propulsion system stability. To accomplish this, several codes were built to run on desktop 386 machines. These codes enable one to analyze the stability question associated with the propellant feed systems. In addition, further work was adapted to this computing environment and furnished along with other codes. This latter inclusion furnishes those interested in high frequency oscillatory combustion behavior (that does not couple to the feed system) a set of codes for study of proposed liquid rocket engines.

Liquid Acquisition Strategies for Exploration Missions: Current Status 2010

NASA Technical Reports Server (NTRS)

Chato, David J.

2010-01-01

NASA is currently developing the propulsion system concepts for human exploration missions to the lunar surface. The propulsion concepts being investigated are considering the use of cryogenic propellants for the low gravity portion of the mission, that is, the lunar transit, lunar orbit insertion, lunar descent and the rendezvous in lunar orbit with a service module after ascent from the lunar surface. These propulsion concepts will require the vapor free delivery of the cryogenic propellants stored in the propulsion tanks to the exploration vehicles main propulsion system (MPS) engines and reaction control system (RCS) engines. Propellant management devices (PMD s) such as screen channel capillary liquid acquisition devices (LAD s), vanes and sponges currently are used for earth storable propellants in the Space Shuttle Orbiter OMS and RCS applications and spacecraft propulsion applications but only very limited propellant management capability exists for cryogenic propellants. NASA has begun a technology program to develop LAD cryogenic fluid management (CFM) technology through a government in-house ground test program of accurately measuring the bubble point delta-pressure for typical screen samples using LO2, LN2, LH2 and LCH4 as test fluids at various fluid temperatures and pressures. This presentation will document the CFM project s progress to date in concept designs, as well ground testing results.

KSC-04pd0950

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the left-hand Orbital Maneuvering System (OMS) pod is lowered toward the orbiter Discovery for installation. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0948

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the left-hand Orbital Maneuvering System (OMS) pod is maneuvered toward the engine interfaces on the orbiter Discovery for installation. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

Study of liquid oxygen/liquid hydrogen auxiliary propulsion systems for the space tug

NASA Technical Reports Server (NTRS)

Nichols, J. F.

1975-01-01

Design concepts are considered that permit use of a liquid-liquid (as opposed to gas-gas) oxygen/hydrogen thrust chamber for attitude control and auxiliary propulsion thrusters on the space tug. The best of the auxiliary propulsion system concepts are defined and their principal characteristics, including cost as well as operational capabilities, are established. Design requirements for each of the major components of the systems, including thrusters, are developed at the conceptual level. The competitive concepts considered use both dedicated (separate tanks) and integrated (propellant from main propulsion tanks) propellant supply. The integrated concept is selected as best for the space tug after comparative evaluation against both cryogenic and storable propellant dedicated systems. A preliminary design of the selected system is established and recommendations for supporting research and technology to further the concept are presented.

Numerical Modeling of Self-Pressurization and Pressure Control by Thermodynamic Vent System in a Cryogenic Tank

NASA Technical Reports Server (NTRS)

Majumdar, Alok; Valenzuela, Juan; LeClair, Andre; Moder, Jeff

2015-01-01

This paper presents a numerical model of a system-level test bed - the multipurpose hydrogen test bed (MHTB) using Generalized Fluid System Simulation Program (GFSSP). MHTB is representative in size and shape of a fully integrated space transportation vehicle liquid hydrogen (LH2) propellant tank and was tested at Marshall Space Flight Center (MSFC) to generate data for cryogenic storage. GFSSP is a finite volume based network flow analysis software developed at MSFC and used for thermo-fluid analysis of propulsion systems. GFSSP has been used to model the self-pressurization and ullage pressure control by Thermodynamic Vent System (TVS). A TVS typically includes a Joule-Thompson (J-T) expansion device, a two-phase heat exchanger, and a mixing pump and spray to extract thermal energy from the tank without significant loss of liquid propellant. Two GFSSP models (Self-Pressurization & TVS) were separately developed and tested and then integrated to simulate the entire system. Self-Pressurization model consists of multiple ullage nodes, propellant node and solid nodes; it computes the heat transfer through Multi-Layer Insulation blankets and calculates heat and mass transfer between ullage and liquid propellant and ullage and tank wall. TVS model calculates the flow through J-T valve, heat exchanger and spray and vent systems. Two models are integrated by exchanging data through User Subroutines of both models. The integrated models results have been compared with MHTB test data of 50% fill level. Satisfactory comparison was observed between test and numerical predictions.

46 CFR 113.37-5 - General requirements.

Code of Federal Regulations, 2013 CFR

2013-10-01

... equipped with fixed pitch propellers must have on the navigating bridge and at the engineroom control station a propeller speed and direction indicator for each shaft. (b) A vessel equipped with controllable pitch propellers must have on the navigating bridge and at the engineroom control station a propeller...

46 CFR 113.37-5 - General requirements.

Code of Federal Regulations, 2014 CFR

2014-10-01

... equipped with fixed pitch propellers must have on the navigating bridge and at the engineroom control station a propeller speed and direction indicator for each shaft. (b) A vessel equipped with controllable pitch propellers must have on the navigating bridge and at the engineroom control station a propeller...

46 CFR 113.37-5 - General requirements.

Code of Federal Regulations, 2012 CFR

2012-10-01

... equipped with fixed pitch propellers must have on the navigating bridge and at the engineroom control station a propeller speed and direction indicator for each shaft. (b) A vessel equipped with controllable pitch propellers must have on the navigating bridge and at the engineroom control station a propeller...

46 CFR 154.1842 - Cargo system: Controls and alarms.

Code of Federal Regulations, 2011 CFR

2011-10-01

... 46 Shipping 5 2011-10-01 2011-10-01 false Cargo system: Controls and alarms. 154.1842 Section 154... SAFETY STANDARDS FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Operations § 154.1842 Cargo system: Controls and alarms. The master shall ensure that the cargo emergency shut-down system and the...

46 CFR 154.1842 - Cargo system: Controls and alarms.

Code of Federal Regulations, 2010 CFR

2010-10-01

... 46 Shipping 5 2010-10-01 2010-10-01 false Cargo system: Controls and alarms. 154.1842 Section 154... SAFETY STANDARDS FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Operations § 154.1842 Cargo system: Controls and alarms. The master shall ensure that the cargo emergency shut-down system and the...

Bistable (latching) solenoid actuated propellant isolation valve

NASA Technical Reports Server (NTRS)

Wichmann, H.; Deboi, H. H.

1979-01-01

The design, fabrication, assembly and test of a development configuration bistable (latching) solenoid actuated propellant isolation valve suitable for the control hydrazine and liquid fluorine to an 800 pound thrust rocket engine is described. The valve features a balanced poppet, utilizing metal bellows, a hard poppet/seat interface and a flexure support system for the internal moving components. This support system eliminates sliding surfaces, thereby rendering the valve free of self generated particles.

40 CFR 1042.505 - Testing engines using discrete-mode or ramped-modal duty cycles.

Code of Federal Regulations, 2014 CFR

2014-07-01

... used with (or intended to be used with) fixed-pitch propellers, propeller-law auxiliary engines, and... with) controllable-pitch propellers or with electrically coupled propellers, unless these engines are... engines that are used with (or intended to be used with) controllable-pitch propellers or with...

Nonlinear Modeling and Control of a Propellant Mixer

NASA Technical Reports Server (NTRS)

Barbieri, Enrique; Richter, Hanz; Figueroa, Fernando

2003-01-01

A mixing chamber used in rocket engine combustion testing at NASA Stennis Space Center is modeled by a second order nonlinear MIMO system. The mixer is used to condition the thermodynamic properties of cryogenic liquid propellant by controlled injection of the same substance in the gaseous phase. The three inputs of the mixer are the positions of the valves regulating the liquid and gas flows at the inlets, and the position of the exit valve regulating the flow of conditioned propellant. The outputs to be tracked and/or regulated are mixer internal pressure, exit mass flow, and exit temperature. The outputs must conform to test specifications dictated by the type of rocket engine or component being tested downstream of the mixer. Feedback linearization is used to achieve tracking and regulation of the outputs. It is shown that the system is minimum-phase provided certain conditions on the parameters are satisfied. The conditions are shown to have physical interpretation.

Cryogenic Technology Development for Exploration Missions

NASA Technical Reports Server (NTRS)

Chato, David J.

2007-01-01

This paper reports the status and findings of different cryogenic technology research projects in support of the President s Vision for Space Exploration. The exploration systems architecture study is reviewed for cryogenic fluid management needs. It is shown that the exploration architecture is reliant on the cryogenic propellants of liquid hydrogen, liquid oxygen and liquid methane. Needs identified include: the key technologies of liquid acquisition devices, passive thermal and pressure control, low gravity mass gauging, prototype pressure vessel demonstration, active thermal control; as well as feed system testing, and Cryogenic Fluid Management integrated system demonstration. Then five NASA technology projects are reviewed to show how these needs are being addressed by technology research. Projects reviewed include: In-Space Cryogenic Propellant Depot; Experimentation for the Maturation of Deep Space Cryogenic Refueling Technology; Cryogenic Propellant Operations Demonstrator; Zero Boil-Off Technology Experiment; and Propulsion and Cryogenic Advanced Development. Advances are found in the areas of liquid acquisition of liquid oxygen, mass gauging of liquid oxygen via radio frequency techniques, computational modeling of thermal and pressure control, broad area cooling thermal control strategies, flight experiments for resolving low gravity issues of cryogenic fluid management. Promising results are also seen for Joule-Thomson pressure control devices in liquid oxygen and liquid methane and liquid acquisition of methane, although these findings are still preliminary.

14 CFR 23.1027 - Propeller feathering system.

Code of Federal Regulations, 2011 CFR

2011-01-01

... Oil System § 23.1027 Propeller feathering system. (a) If the propeller feathering system uses engine... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Propeller feathering system. 23.1027... made to prevent sludge or other foreign matter from affecting the safe operation of the propeller...

14 CFR 23.1027 - Propeller feathering system.

Code of Federal Regulations, 2013 CFR

2013-01-01

... Oil System § 23.1027 Propeller feathering system. (a) If the propeller feathering system uses engine... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Propeller feathering system. 23.1027... made to prevent sludge or other foreign matter from affecting the safe operation of the propeller...

14 CFR 23.1027 - Propeller feathering system.

Code of Federal Regulations, 2014 CFR

2014-01-01

... Oil System § 23.1027 Propeller feathering system. (a) If the propeller feathering system uses engine... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Propeller feathering system. 23.1027... made to prevent sludge or other foreign matter from affecting the safe operation of the propeller...

14 CFR 23.1027 - Propeller feathering system.

Code of Federal Regulations, 2012 CFR

2012-01-01

... Oil System § 23.1027 Propeller feathering system. (a) If the propeller feathering system uses engine... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Propeller feathering system. 23.1027... made to prevent sludge or other foreign matter from affecting the safe operation of the propeller...

14 CFR 23.1027 - Propeller feathering system.

Code of Federal Regulations, 2010 CFR

2010-01-01

... Oil System § 23.1027 Propeller feathering system. (a) If the propeller feathering system uses engine... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Propeller feathering system. 23.1027... made to prevent sludge or other foreign matter from affecting the safe operation of the propeller...

Inflight performance of the Ulysses reaction control system

NASA Technical Reports Server (NTRS)

McGarry, Andrew; Berry, William; Parker, David

1997-01-01

The Ulysses spacecraft has been exploring the heliosphere since October 1990 in a six-year polar orbit. Despite varying operational demands, the pressure-fed monopropellant hydrazine reaction control system (RCS) has experienced few problems. The observed anomalies, having minimal operational impact, include plume impingement effects, electrical power overload effects and hydrazine gas generation effects. These anomalies are presented and discussed, with emphasis on the first observation of gas in the hydrazine propellant. The relatively low gas generation rate is attributed to: the use of high purity hydrazine; the configuration of the spin-stabilized spacecraft; the extensive use of titanium alloys; and the efficiency of the thermal control of the propellant tank which maintains a temperature of 21 C.

Iodine Satellite

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Dankanich, John; Martinez, Andres; Petro, Andrew

2015-01-01

The Iodine Satellite (iSat) spacecraft will be the first CubeSat to demonstrate high change in velocity from a primary propulsion system by using Hall thruster technology and iodine as a propellant. The mission will demonstrate CubeSat maneuverability, including plane change, altitude change and change in its closest approach to Earth to ensure atmospheric reentry in less than 90 days. The mission is planned for launch in fall 2017. Hall thruster technology is a type of electric propulsion. Electric propulsion uses electricity, typically from solar panels, to accelerate the propellant. Electric propulsion can accelerate propellant to 10 times higher velocities than traditional chemical propulsion systems, which significantly increases fuel efficiency. To enable the success of the propulsion subsystem, iSat will also demonstrate power management and thermal control capabilities well beyond the current state-of-the-art for spacecraft of its size. This technology is a viable primary propulsion system that can be used on small satellites ranging from about 22 pounds (10 kilograms) to more than 1,000 pounds (450 kilograms). iSat's fuel efficiency is ten times greater and its propulsion per volume is 100 times greater than current cold-gas systems and three times better than the same system operating on xenon. iSat's iodine propulsion system consists of a 200 watt (W) Hall thruster, a cathode, a tank to store solid iodine, a power processing unit (PPU) and the feed system to supply the iodine. This propulsion system is based on a 200 W Hall thruster developed by Busek Co. Inc., which was previously flown using xenon as the propellant. Several improvements have been made to the original system to include a compact PPU, targeting greater than 80 percent reduction in mass and volume of conventional PPU designs. The cathode technology is planned to enable heaterless cathode conditioning, significantly increasing total system efficiency. The feed system has been designed to include iodine compatible control valves with internal heaters and temperature sensors to coincide with the iodine-compatible thruster. A key advantage to using iodine as a propellant is that it may be stored in the tank as an unpressurized solid on the ground and before flight operations. During operations, the tank is heated to vaporize the propellant. Iodine vapor is then routed through custom flow control valves to control mass flow to the thruster and cathode assembly. The thruster then ionizes the vapor and accelerates it via magnetic and electrostatic fields, resulting in high specific impulse, characteristic of a highly efficient propulsion system. The iSat spacecraft is a 12-unit (12U) CubeSat with dimensions of about 8 inches x 8 inches x 12 inches (20 centimeters x 20 centimeters x 30 centimeters). The spacecraft frame will be constructed from aluminum with a finish to prevent iodine-driven corrosion. The iSat spacecraft includes full three-axis control and will leverage heat generated by spacecraft components and radiators for a passive thermal control system. After the CubeSat has successfully detached from its launch vehicle, it will deploy its solar panels, correct for tip-off and maintain attitude control before ground contact. An initial check-out period of two weeks is planned for testing all subsystems. The spacecraft will charge the power system while in sunlight, using momentum wheels and magnetic torque rods to rotate the vehicle to the required attitude.

Concept and performance study of turbocharged solid propellant ramjet

NASA Astrophysics Data System (ADS)

Li, Jiang; Liu, Kai; Liu, Yang; Liu, Shichang

2018-06-01

This study proposes a turbocharged solid propellant ramjet (TSPR) propulsion system that integrates a turbocharged system consisting of a solid propellant (SP) air turbo rocket (ATR) and the fuel-rich gas generator of a solid propellant ramjet (SPR). First, a suitable propellant scheme was determined for the TSPR. A solid hydrocarbon propellant is used to generate gas for driving the turbine, and a boron-based fuel-rich propellant is used to provide fuel-rich gas to the afterburner. An appropriate TSPR structure was also determined. The TSPR's thermodynamic cycle was analysed to prove its theoretical feasibility. The results showed that the TSPR's specific cycle power was larger than those of SP-ATR and SPR and thermal efficiency was slightly less than that of SP-ATR. Overall, TSPR showed optimal performance in a wide flight envelope. The specific impulses and specific thrusts of TSPR, SP-ATR, and SPR in the flight envelope were calculated and compared. TSPR's flight envelope roughly overlapped that of SP-ATR, its specific impulse was larger than that of SP-ATR, and its specific thrust was larger than those of SP-ATR and SPR. Attempts to improve the TSPR off-design performance prompted our proposal of a control plan for off-design codes in which both the turbocharger corrected speed and combustor excess gas coefficient are kept constant. An off-design performance model was established by analysing the TSPR working process. We concluded that TSPR with a constant corrected speed had wider flight envelope, higher thrust, and higher specific impulse than TSPR with a constant physical speed determined by calculating the performance of off-design TSPR codes under different control plans. The results of this study can provide a reference for further studies on TSPRs.

Pressurization, Pneumatic, and Vent Subsystems of the X-34 Main Propulsion System

NASA Technical Reports Server (NTRS)

Hedayat, A.; Steadman, T. E.; Brown, T. M.; Knight, K. C.; White, C. E., Jr.; Champion, R. H., Jr.

1998-01-01

In pressurization systems, regulators and orifices are use to control the flow of the pressurant. For the X-34 Main Propulsion System, three pressurization subsystem design configuration options were considered. In the first option, regulators were used while in the other options, orifices were considered. In each design option, the vent/relief system must be capable of relieving the pressurant flow without allowing the tank pressure to rise above proof, therefore, impacts on the propellant tank vent system were investigated and a trade study of the pressurization system was conducted. The analysis indicated that design option using regulators poses least risk. Then, a detailed transient thermal/fluid analysis of the recommended pressurization system was performed. Helium usage, thermodynamic conditions, and overpressurization of each propellant tank were evaluated. The pneumatic and purge subsystem is used for pneumatic valve actuation, Inter-Propellant Seal purges, Engine Spin Start, and engine purges at the required interface pressures, A transient analysis of the pneumatic and purge subsystem provided helium usage and flow rates to Inter-Propellant Seal and engine interfaces. Fill analysis of the helium bottles of pressurization and pneumatic subsystems during ground operation was performed. The required fill time and the stored

Thrust Control Loop Design for Electric-Powered UAV

NASA Astrophysics Data System (ADS)

Byun, Heejae; Park, Sanghyuk

2018-04-01

This paper describes a process of designing a thrust control loop for an electric-powered fixed-wing unmanned aerial vehicle equipped with a propeller and a motor. In particular, the modeling method of the thrust system for thrust control is described in detail and the propeller thrust and torque force are modeled using blade element theory. A relation between current and torque of the motor is obtained using an experimental setup. Another relation between current, voltage and angular velocity is also obtained. The electric motor and the propeller dynamics are combined to model the thrust dynamics. The associated trim and linearization equations are derived. Then, the thrust dynamics are coupled with the flight dynamics to allow a proper design for the thrust loop in the flight control. The proposed method is validated by an application to a testbed UAV through simulations and flight test.

Concept of planetary gear system to control fluid mixture ratio

NASA Technical Reports Server (NTRS)

Mcgroarty, J. D.

1966-01-01

Mechanical device senses and corrects for fluid flow departures from the selected flow ratio of two fluids. This system has been considered for control of rocket engine propellant mixture control but could find use wherever control of the flow ratio of any two fluids is desired.

Surface tension propellant control for Viking 75 Orbiter

NASA Technical Reports Server (NTRS)

Dowdy, M. W.; Hise, R. E.; Peterson, R. G.; Debrock, S. C.

1976-01-01

The paper describes the selection, development and qualification of the surface tension system and includes results of low-g drop tower tests of scale models, 1-g simulation tests of low-g large ullage settling and liquid withdrawal, structural qualification tests, and propellant surface tension/contact angle studies. Subscale testing and analyses were used to evaluate the ability of the system to maintain or recover the desired propellant orientation following possible disturbances during the Viking mission. This effort included drop tower tests to demonstrate that valid wick paths exist for moving any displaced propellant back over the tank outlet. Variations in surface tension resulting from aging, temperature, and lubricant contamination were studied and the effects of surface finish, referee fluid exposure, aging, and lubricant contamination on contact angle were assessed. Results of movies of typical subscale drop tower tests and full scale slosh tests are discussed.

Eight cm technology thruster development. [structurally integrated ion thruster for attitude control and stationkeeping of synchronous satellites

NASA Technical Reports Server (NTRS)

Hyman, J., Jr.

1974-01-01

A structural integrated ion thruster with 8-cm beam diameter (SIT-8) was developed for attitude control and stationkeeping of synchronous satellites. As optimized, the system demonstrates a thrust T=1.14 mlb (not corrected for beam V sub B = 1200 V (I sub sp = 2200 sec) total propellant utilization efficiency nu sub u = 59.8% (is approximately 72% without auxiliary pulse-igniter electrode), and electrical efficiency n sub E 61.9%. The thruster incorporates a wire-mesh anode and tantalum cover surfaces to control discharge chamber flake formation and employs an auxiliary pulse-igniter electrode for hollow-cathode ignition. When the SIT-8 is integrated with the compatible SIT-5 propellant tankage, the system envelope is 35 cm long by 13 cm flange bolt circle with a mass of 9.8 kg including 6.8 kg of mercury propellant. Two thrust vectoring systems which generate beam deflections in two orthogonal directions were also developed under the program and tested with the 8-cm thruster. One system vectors the beam over + or - 10 degrees by gimbaling of the entire thruster (not including tankage), while the other system vectors the beam over + or - 7 degrees by translating the accel electrode relative to the screen electrode.

A solenoid failure detection system for cold gas attitude control jet valves

NASA Technical Reports Server (NTRS)

Johnston, P. A.

1970-01-01

The development of a solenoid valve failure detection system is described. The technique requires the addition of a radioactive gas to the propellant of a cold gas jet attitude control system. Solenoid failure is detected with an avalanche radiation detector located in the jet nozzle which senses the radiation emitted by the leaking radioactive gas. Measurements of carbon monoxide leakage rates through a Mariner type solenoid valve are presented as a function of gas activity and detector configuration. A cylindrical avalanche detector with a factor of 40 improvement in leak sensitivity is proposed for flight systems because it allows the quantity of radioactive gas that must be added to the propellant to be reduced to a practical level.

Subsystem Analysis/Optimization for the X-34 Main Propulsion System

NASA Technical Reports Server (NTRS)

McDonald, J. P.; Hedayat, A.; Brown, T. M.; Knight, K. C.; Champion, R. H., Jr.

1998-01-01

The Orbital Sciences Corporation X-34 vehicle demonstrates technologies and operations key to future reusable launch vehicles. The general flight performance goal of this unmanned rocket plane is Mach 8 flight at an altitude of 250,000 feet. The Main Propulsion System (MPS) supplies liquid propellants to the main engine, which provides the primary thrust for attaining mission goals. Major MPS design and operational goals are aircraft-like ground operations, quick turnaround between missions, and low initial/operational costs. Analyses related to optimal MPS subsystem design are reviewed in this paper. A pressurization system trade weighs maintenance/reliability concerns against those for safety in a comparison of designs using pressure regulators versus orifices to control pressurant flow. A propellant dump/feed system analysis weighs the issues of maximum allowable vehicle landing weight, trajectory, and MPS complexity to arrive at a final configuration for propellant dump/feed systems.

Spacecraft and mission design for the SP-100 flight experiment

NASA Technical Reports Server (NTRS)

Deininger, William D.; Vondra, Robert J.

1988-01-01

The design and performance of a spacecraft employing arcjet nuclear electric propulsion, suitable for use in the SP-100 Space Reactor Power System (SRPS) Flight Experiment, are outlined. The vehicle design is based on a 93 kW(e) ammonia arcjet system operating at an experimentally measured specific impulse of 1031 s and an efficiency of 42.3 percent. The arcjet/gimbal assemblies, power conditioning subsystem, propellant feed system, propulsion system thermal control, spacecraft diagnostic instrumentation, and the telemetry requirements are described. A 100 kW(e) SRPS is assumed. The spacecraft mass is baselined at 5675 kg excluding the propellant and propellant feed system. Four mission scenarios are described which are capable of demonstrating the full capability of the SRPS. The missions considered include spacecraft deployment to possible surveillance platform orbits, a spacecraft storage mission, and an orbit raising round trip corresponding to possible orbit transfer vehicle (OTV) missions.

Subcooling Cryogenic Propellants for Long Duration Space Exploration

NASA Technical Reports Server (NTRS)

Mustafi, Shuvo; Canavan, Edgar; Johnson, Wesley; Kutter, Bernard; Shull, Jeff

2009-01-01

The use of cryogenic propellants such as hydrogen and oxygen is crucial for exploration of the solar system because of their superior specific impulse capability. Future missions may require vehicles with the flexibility to remain in orbit or travel in space for months, necessitating long-term storage of these cryogens. One powerful technique for easing the challenge of cryogenic fluid storage is to remove energy from tlie cryogenic propellant by isobaricly subcooling them below their normal boiling point prior to launch. The isobaric subcooling of the cryogenic propellant will be performed by using a cold pressurant to maintain the tank pressure while the cryogen's temperature is simultaneously reduced. After launch, even with the use of the best insulation systems, heat will leak into the cold cryogenic propellant tank. However, the large heat capacity available in highly subcooled cryogenic propellants allows them to absorb the energy that leaks into the tank until the cryogen reaches its operational thermodynamic condition. During this period of heating of the subcooled cryogen there will be no loss of the propellant due to venting for pressure control. This simple technique can extend the operational life of a spacecraft or an orbital cryogenic depot many months with minimal mass penalty. Subcooling technologies for cryogenic propellants would thus provide the Exploration Systems Mission Directorate with an enhanced level of mission flexibility. However, there are a few challenges associated with subcooling cryogenic propellants since compact subcooling ground support equipment has not been demonstrated. This paper explores the beneficial impact of subcooling cryogenic propellants on the launch pad for long-term cryogenic propellant storage in space and proposes a novel method for implementing subcooling of cryogenic propellants for spacecraft such as the Ares V Earth Departure Stage (EDS). Analysis indicates that with a careful strategy to handle the subcooled cryogen it would be possible to store cryogenic propellants in space for many months without venting. A concept for subcooling the cryogenic propellant relatively quickly and inexpensively on the launch pad - the thermodynamic cryogen subcooler (TCS) - will be presented. Important components of the TCS and an associated subcooled cryogen tank (SCT) will be discussed in this paper. Results from a preliminary thermodynamic model of the performance of a TCS for an EDS sized hydrogen tank will also be presented.

General airplane performance

NASA Technical Reports Server (NTRS)

Rockfeller, W C

1939-01-01

Equations have been developed for the analysis of the performance of the ideal airplane, leading to an approximate physical interpretation of the performance problem. The basic sea-level airplane parameters have been generalized to altitude parameters and a new parameter has been introduced and physically interpreted. The performance analysis for actual airplanes has been obtained in terms of the equivalent ideal airplane in order that the charts developed for use in practical calculations will for the most part apply to any type of engine-propeller combination and system of control, the only additional material required consisting of the actual engine and propeller curves for propulsion unit. Finally, a more exact method for the calculation of the climb characteristics for the constant-speed controllable propeller is presented in the appendix.

46 CFR 154.903 - Inert gas systems: General.

Code of Federal Regulations, 2011 CFR

2011-10-01

... STANDARDS FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design, Construction and Equipment Atmospheric Control in Cargo Containment Systems § 154.903 Inert gas systems: General. (a) Inert gas carried...

Zero Boil-Off System Testing

NASA Technical Reports Server (NTRS)

Plachta, David W.; Johnson, Wesley L.; Feller, Jeffrey R.

2015-01-01

Cryogenic propellants such as liquid hydrogen (LH2) and liquid oxygen (LO2) are a part of NASA's future space exploration due to their high specific impulse for rocket motors of upper stages suitable for transporting 10s to 100s of metric tons of payload mass to destinations outside of low earth orbit and for their return. However, the low storage temperatures of LH2 and LO2 cause substantial boil-off losses for missions with durations greater than several months. These losses can be eliminated by incorporating high performance cryocooler technology to intercept heat load to the propellant tanks and modulating the cryocooler to control tank pressure. The active thermal control technology being developed by NASA is the reverse turbo-Brayton cycle cryocooler and its integration to the propellant tank through a distributed cooling tubing network coupled to the tank wall. This configuration was recently tested at NASA Glenn Research Center, in a vacuum chamber and cryoshroud that simulated the essential thermal aspects of low Earth orbit, its vacuum and temperature. Testing consisted of three passive tests with the active cryocooler system off, and 7 active tests, with the cryocooler powered up. The test matrix included zero boil-off tests performed at 90 full and 25 full, and several demonstrations at excess cooling capacity and reduced cooling capacity. From this, the tank pressure response with varied cryocooler power inputs was determined. This test series established that the active cooling system integrated with the propellant tank eliminated boil-off and robustly controlled tank pressure.

Experimental active control of sound in the ATR 42

NASA Astrophysics Data System (ADS)

Paonessa, A.; Sollo, A.; Paxton, M.; Purver, M.; Ross, C. F.

Passenger comfort is becoming day by day an important issue for the market of the regional turboprop aircraft and also for the future high speed propeller driven aircraft. In these aircraft the main contribution to the passenger annoyance is due to the propeller noise blade passing frequency (BPF) and its harmonics. In the recent past a detailed theoretical and experimental work has been done by Alenia Aeronautica in order to reduce the noise level in the ATR aircraft passenger cabin by means of conventional passive treatments: synchrophasing of propellers, dynamic vibration absorbers, structural reinforcements, damping materials. The application of these treatments has been introduced on production aircraft with a remarkable improvement of noise comfort but with a significant weight increase. For these reasons, a major technology step is required for reaching passenger comfort comparable to that of jet aircraft with the minimum weight increase. The most suitable approach to this problem has been envisaged in the active noise control which consists in generating an anti-sound field in the passenger cabin to reduce the noise at propeller BPF and its harmonics. The attenuation is reached by means of a control system which acquires information about the cabin noise distribution and the propeller speed during flight and simultaneously generates the signals to drive the speakers.

KSC-04pd0942

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the left-hand Orbital Maneuvering System (OMS) pod (seen from the back) is lifted off its transporter. The OMS pod will be installed on the orbiter Discovery. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0949

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers stand by as the left-hand Orbital Maneuvering System (OMS) pod is maneuvered toward the engine interfaces on the orbiter Discovery for installation. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0944

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the left-hand Orbital Maneuvering System (OMS) pod is lifted at an angle from the transporter below. The OMS pod will be installed on the orbiter Discovery. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0947

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the left-hand Orbital Maneuvering System (OMS) pod (top of photo) is poised behind the engine interfaces on the orbiter Discovery for installation. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0946

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers on an upper level watch as the left-hand Orbital Maneuvering System (OMS) pod is lifted high to maneuver it toward the orbiter Discovery for installation. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0943

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers check the lifting of the left-hand Orbital Maneuvering System (OMS) pod. The OMS pod will be installed on the orbiter Discovery. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0941

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the left-hand Orbital Maneuvering System (OMS) pod (seen from the front) is lifted off its transporter. The OMS pod will be installed on the orbiter Discovery. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0945

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, a worker on an upper level watches as the left-hand Orbital Maneuvering System (OMS) pod is lifted high to maneuver it toward the orbiter Discovery for installation. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

Cryogenic Fluid Management Technology Development Roadmaps

NASA Technical Reports Server (NTRS)

Stephens, J. R.; Johnson, W. L.

2017-01-01

Advancement in Cryogenic Fluid Management (CFM) Technologies is essential for achieving NASA's future long duration missions. Propulsion systems utilizing cryogens are necessary to achieve mission success. Current State Of the Art (SOA) CFM technologies enable cryogenic propellants to be stored for several hours. However, some envisioned mission architectures require cryogens to be stored for two years or longer. The fundamental roles of CFM technologies are long term storage of cryogens, propellant tank pressure control and propellant delivery. In the presence of heat, the cryogens will "boil-off" over time resulting in excessive pressure buildup, off-nominal propellant conditions, and propellant loss. To achieve long term storage and tank pressure control, the CFM elements will intercept and/or remove any heat from the propulsion system. All functions are required to perform both with and without the presence of a gravitational field. Which CFM technologies are required is a function of the cryogens used, mission architecture, vehicle design and propellant tank size. To enable NASA's crewed mission to the Martian surface, a total of seventeen CFM technologies have been identified to support an In-Space Stage and a Lander/Ascent Vehicle. Recognizing that FY2020 includes a Decision Point regarding the In-Space Stage Architecture, a set of CFM Technology Development Roadmaps have been created identifying the current Technology Readiness Level (TRL) of each element, current technology "gaps", and existing technology development efforts. The roadmaps include a methodical approach and schedule to achieve a flight demonstration in FY2023, hence maturing CFM technologies to TRL 7 for infusion into the In-Space Stage Preliminary Design.

Research on Modeling of Propeller in a Turboprop Engine

NASA Astrophysics Data System (ADS)

Huang, Jiaqin; Huang, Xianghua; Zhang, Tianhong

2015-05-01

In the simulation of engine-propeller integrated control system for a turboprop aircraft, a real-time propeller model with high-accuracy is required. A study is conducted to compare the real-time and precision performance of propeller models based on strip theory and lifting surface theory. The emphasis in modeling by strip theory is focused on three points as follows: First, FLUENT is adopted to calculate the lift and drag coefficients of the propeller. Next, a method to calculate the induced velocity which occurs in the ground rig test is presented. Finally, an approximate method is proposed to obtain the downwash angle of the propeller when the conventional algorithm has no solution. An advanced approximation of the velocities induced by helical horseshoe vortices is applied in the model based on lifting surface theory. This approximate method will reduce computing time and remain good accuracy. Comparison between the two modeling techniques shows that the model based on strip theory which owns more advantage on both real-time and high-accuracy can meet the requirement.

Cryogenic Propellant Management Device: Conceptual Design Study

NASA Technical Reports Server (NTRS)

Wollen, Mark; Merino, Fred; Schuster, John; Newton, Christopher

2010-01-01

Concepts of Propellant Management Devices (PMDs) were designed for lunar descent stage reaction control system (RCS) and lunar ascent stage (main and RCS propulsion) missions using liquid oxygen (LO2) and liquid methane (LCH4). Study ground rules set a maximum of 19 days from launch to lunar touchdown, and an additional 210 days on the lunar surface before liftoff. Two PMDs were conceptually designed for each of the descent stage RCS propellant tanks, and two designs for each of the ascent stage main propellant tanks. One of the two PMD types is a traditional partial four-screen channel device. The other type is a novel, expanding volume device which uses a stretched, flexing screen. It was found that several unique design features simplified the PMD designs. These features are (1) high propellant tank operating pressures, (2) aluminum tanks for propellant storage, and (3) stringent insulation requirements. Consequently, it was possible to treat LO2 and LCH4 as if they were equivalent to Earth-storable propellants because they would remain substantially subcooled during the lunar mission. In fact, prelaunch procedures are simplified with cryogens, because any trapped vapor will condense once the propellant tanks are pressurized in space.

SPINS-IND: Pellet injector for fuelling of magnetically confined fusion systems.

PubMed

Gangradey, R; Mishra, J; Mukherjee, S; Panchal, P; Nayak, P; Agarwal, J; Saxena, Y C

2017-06-01

Using a Gifford-McMahon cycle cryocooler based refrigeration system, a single barrel hydrogen pellet injection (SPINS-IND) system is indigenously developed at Institute for Plasma Research, India. The injector is based on a pipe gun concept, where a pellet formed in situ in the gun barrel is accelerated to high speed using high pressure light propellant gas. The pellet size is decided by considering the Greenwald density limit and its speed is decided by considering a neutral gas shielding model based scaling law. The pellet shape is cylindrical of dimension (1.6 mm ℓ × 1.8 mm φ). For pellet ejection and acceleration, a fast opening valve of short opening duration is installed at the breech of the barrel. A three-stage differential pumping system is used to restrict the flow of the propellant gas into the plasma vacuum vessel. Diagnostic systems such as light gate and fast imaging camera (240 000 frames/s) are employed to measure the pellet speed and size, respectively. A trigger circuit and a programmable logic controller based integrated control system developed on LabVIEW enables to control the pellet injector remotely. Using helium as a propellant gas, the pellet speed is varied in the range 650 m/s-800 m/s. The reliability of pellet formation and ejection is found to be more than 95%. This paper describes the details of SPINS-IND and its test results.

SPINS-IND: Pellet injector for fuelling of magnetically confined fusion systems

NASA Astrophysics Data System (ADS)

Gangradey, R.; Mishra, J.; Mukherjee, S.; Panchal, P.; Nayak, P.; Agarwal, J.; Saxena, Y. C.

2017-06-01

Using a Gifford-McMahon cycle cryocooler based refrigeration system, a single barrel hydrogen pellet injection (SPINS-IND) system is indigenously developed at Institute for Plasma Research, India. The injector is based on a pipe gun concept, where a pellet formed in situ in the gun barrel is accelerated to high speed using high pressure light propellant gas. The pellet size is decided by considering the Greenwald density limit and its speed is decided by considering a neutral gas shielding model based scaling law. The pellet shape is cylindrical of dimension (1.6 mm ℓ × 1.8 mm φ). For pellet ejection and acceleration, a fast opening valve of short opening duration is installed at the breech of the barrel. A three-stage differential pumping system is used to restrict the flow of the propellant gas into the plasma vacuum vessel. Diagnostic systems such as light gate and fast imaging camera (240 000 frames/s) are employed to measure the pellet speed and size, respectively. A trigger circuit and a programmable logic controller based integrated control system developed on LabVIEW enables to control the pellet injector remotely. Using helium as a propellant gas, the pellet speed is varied in the range 650 m/s-800 m/s. The reliability of pellet formation and ejection is found to be more than 95%. This paper describes the details of SPINS-IND and its test results.

Combustion performance and heat transfer characterization of LOX/hydrocarbon type propellants. Task 3: Data dump

NASA Technical Reports Server (NTRS)

Hart, S. W.

1982-01-01

A preliminary characterization of Orbital Maneuvering System (OMS) and Reaction Control System (RCS) engine point designs over a range of thrust and chamber pressure for several hydrocarbon fuels is reported. OMS and RCS engine point designs were established in two phases comprising baseline and parametric designs. Interface pressures, performance and operating parameters, combustion chamber cooling and turboprop requirements, component weights and envelopes, and propellant conditioning requirements for liquid to vapor phase engine operation are defined.

Engineering Model Propellant Feed System Development for an Iodine Hall Thruster Demonstration Mission

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.

2016-01-01

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cu cm and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high (Delta)v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ?Isp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (< 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp -6) torr at -75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. In the present paper, we describe the design and testing of the engineering model propellant feed system for iSAT (see Fig. 1). The feed system is based around an iodine propellant reservoir and two proportional control valves (PFCVs) that meter the iodine flow to the cathode and anode. The flow is split upstream of the PFCVs to both components can be fed from a common reservoir. Testing of the reservoir is reported to demonstrate that the design is capable of delivering the required propellant flow rates to operate the thruster. The tubing and reservoir are fabricated from hastelloy to resist corrosion by the heated gaseous iodine propellant. The reservoir, tubing, and PFCVs are heated to ensure the sublimed propellant will not re-deposit within the feed system. Heating is accomplished through a number of individual zones to control the overall power expended on heating the system and insulation is employed to minimize the amount of power used to heat the system prior to thruster operation.

Negative feedback system reduces pump oscillations

NASA Technical Reports Server (NTRS)

Rosenmann, W.

1967-01-01

External negative feedback system counteracts low frequency oscillations in rocket engine propellant pumps. The system uses a control piston to sense pump discharge fluid on one side and a gas pocket on the other.

Effectiveness of Population Health Management Using the Propeller Health Asthma Platform: A Randomized Clinical Trial.

PubMed

Merchant, Rajan K; Inamdar, Rubina; Quade, Robert C

2016-01-01

Telehealth strategies for asthma have focused primarily on adherence to controller medications. Telemonitoring of short-acting β-agonist (SABA) focuses on patterns of use and may allow more timely action to avert exacerbations. Studies assessing this approach are lacking. This pragmatic controlled study was designed to measure real-world effectiveness of the Propeller Health Asthma Platform to reduce use of SABA and improve asthma control. A total of 495 patients were enrolled in parallel arms (1:1) for 12 months of monitoring SABA use. Intervention group (IG) patients received access to and feedback from the Propeller Health system. Routine care (RC) patients were outfitted with sensors but did not receive feedback. Physicians were able to monitor the status of their patients in the IG and receive proactive notifications. The daily mean number of SABA uses per person decreased by 0.41 for the IG and by 0.31 for RC between the first week and the remainder of the study period (P < .001 for the difference between groups). Similarly, the proportion of SABA-free days increased 21% for the IG and 17% for RC (P < .01 for the difference between groups). Asthma Control Test (ACT) scores were not significantly different between arms in the entire study population, but adults with initially uncontrolled ACT scores showed a significantly larger improvement in the proportion with controlled asthma in IG versus RC (63% controlled in the study period vs 49%, respectively; P < .05 comparing the 2 improvements). Compared with RC, the study arm monitoring SABA use with the Propeller Health system significantly decreased SABA use, increased SABA-free days, and improved ACT scores (the latter among adults initially lacking asthma control). Copyright © 2016 American Academy of Allergy, Asthma & Immunology. Published by Elsevier Inc. All rights reserved.

Cryogenic Propellant Feed System Analytical Tool Development

NASA Technical Reports Server (NTRS)

Lusby, Brian S.; Miranda, Bruno M.; Collins, Jacob A.

2011-01-01

The Propulsion Systems Branch at NASA s Lyndon B. Johnson Space Center (JSC) has developed a parametric analytical tool to address the need to rapidly predict heat leak into propellant distribution lines based on insulation type, installation technique, line supports, penetrations, and instrumentation. The Propellant Feed System Analytical Tool (PFSAT) will also determine the optimum orifice diameter for an optional thermodynamic vent system (TVS) to counteract heat leak into the feed line and ensure temperature constraints at the end of the feed line are met. PFSAT was developed primarily using Fortran 90 code because of its number crunching power and the capability to directly access real fluid property subroutines in the Reference Fluid Thermodynamic and Transport Properties (REFPROP) Database developed by NIST. A Microsoft Excel front end user interface was implemented to provide convenient portability of PFSAT among a wide variety of potential users and its ability to utilize a user-friendly graphical user interface (GUI) developed in Visual Basic for Applications (VBA). The focus of PFSAT is on-orbit reaction control systems and orbital maneuvering systems, but it may be used to predict heat leak into ground-based transfer lines as well. PFSAT is expected to be used for rapid initial design of cryogenic propellant distribution lines and thermodynamic vent systems. Once validated, PFSAT will support concept trades for a variety of cryogenic fluid transfer systems on spacecraft, including planetary landers, transfer vehicles, and propellant depots, as well as surface-based transfer systems. The details of the development of PFSAT, its user interface, and the program structure will be presented.

A feasibility study of developing toroidal tanks for a spinning spacecraft

NASA Technical Reports Server (NTRS)

Anderson, J. E.; Fester, D. A.

1973-01-01

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.

Development of hybrid method for the prediction of underwater propeller noise

NASA Astrophysics Data System (ADS)

Seol, Hanshin; Suh, Jung-Chun; Lee, Soogab

2005-11-01

Noise reduction and control is an important problem in the performance of underwater acoustic systems and in the habitability of the passenger ship for crew and passenger. Furthermore, sound generated by a propeller is critical in underwater detection and it is often related to the survivability of the vessel especially for military purpose. This paper presents a numerical study on the non-cavitating and blade sheet cavitation noises of the underwater propeller. A brief summary of numerical method with verification and results are presented. The noise is predicted using time-domain acoustic analogy. The flow field is analyzed with potential-based panel method, and then the time-dependent pressure and sheet cavity volume data are used as the input for Ffowcs Williams-Hawkings formulation to predict the far-field acoustics. Noise characteristics are presented according to noise sources and conditions. Through this study, the dominant noise source of the underwater propeller is analyzed, which will provide a basis for proper noise control strategies.

Space storable propulsion components development

NASA Technical Reports Server (NTRS)

Hagler, R., Jr.

1982-01-01

The current development status of components to control the flow of propellants (liquid fluorine and hydrazine) in a demonstration space storable propulsion system is discussed. The criteria which determined the designs for the pressure regulator, explosive-actuated valves, propellant shutoff valve, latching solenoid-actuated valve and propellant filter are presented. The test philosophy that was followed during component development is outlined. The results from compatibility demonstrations for reusable connectors, flange seals, and CRES/Ti-6Al4V transition tubes and the evaluations of processes for welding (hand-held TIG, automated TIG, and EB), cleaning for fluorine service, and decontamination after fluorine exposure are described.

Non-Toxic Orbiter Maneuvering System (OMS) and Reaction Control System

NASA Technical Reports Server (NTRS)

Hurlbert, Eric A.; Nicholson, Leonard S. (Technical Monitor)

1999-01-01

NASA is pursuing the technology and advanced development of a non-toxic (NT) orbital maneuvering system (OMS) and reaction control system (RCS) for shuttle upgrades, RLV, and reusable first stages. The primary objectives of the shuttle upgrades program are improved safety, improved reliability, reduced operations time and cost, improved performance or capabilities, and commonality with future space exploration needs. Non-Toxic OMS/RCS offers advantages in each of these categories. A non-toxic OMS/RCS eliminates the ground hazards and the flight safety hazards of the toxic and corrosive propellants. The cost savings for ground operations are over $24M per year for 7 flights, and the savings increase with increasing flight rate up to $44M per year. The OMS/RCS serial processing time is reduced from 65 days to 13 days. The payload capability can be increased up to 5100 Ibms. The non-toxic OMS/RCS also provides improved space station reboost capability up to 20 nautical miles over the current toxic system of 14 nautical miles. A NT OMS/RCS represents a clear advancement in the SOA over MMH/NTO. Liquid oxygen and ethanol are clean burning, high-density propellants that provide a high degree of commonality with other spacecraft subsystems including life support, power, and thermal control, and with future human exploration and development of space missions. The simple and reliable pressure-fed design uses sub-cooled liquid oxygen at 250 to 350 psia, which allows a propellant to remain cryogenic for longer periods of time. The key technologies are thermal insulation and conditioning techniques are used to maintain the sub-cooling. Phase I successfully defined the system architecture, designed an integrated OMS/RCS propellant tank, analyzed the feed system, built and tested the 870 lbf RCS thrusters, and tested the 6000 lbf OMS engine. Phase 11 is currently being planned for the development and test of full-scale prototype of the system in 1999 and 2000

40 CFR 1042.505 - Testing engines using discrete-mode or ramped-modal duty cycles.

Code of Federal Regulations, 2013 CFR

2013-07-01

... used with) controllable-pitch propellers or with electrically coupled propellers, unless these engines... engines that are used with (or intended to be used with) controllable-pitch propellers or with electrically coupled propellers. Use this duty cycle also for variable-speed propulsion marine engines that are...

40 CFR 1042.505 - Testing engines using discrete-mode or ramped-modal duty cycles.

Code of Federal Regulations, 2011 CFR

2011-07-01

... used with) controllable-pitch propellers or with electrically coupled propellers, unless these engines... engines that are used with (or intended to be used with) controllable-pitch propellers or with electrically coupled propellers. Use this duty cycle also for variable-speed propulsion marine engines that are...

40 CFR 1042.505 - Testing engines using discrete-mode or ramped-modal duty cycles.

Code of Federal Regulations, 2012 CFR

2012-07-01

... used with) controllable-pitch propellers or with electrically coupled propellers, unless these engines... engines that are used with (or intended to be used with) controllable-pitch propellers or with electrically coupled propellers. Use this duty cycle also for variable-speed propulsion marine engines that are...

75 FR 13238 - Special Conditions: McCauley Propeller Systems, Model Propeller 3D15C1401/C80MWX-X

Federal Register 2010, 2011, 2012, 2013, 2014

2010-03-19

...-X AGENCY: Federal Aviation Administration (FAA), DOT. ACTION: Notice of proposed special conditions... for McCauley Propeller Systems for model propeller 3D15C1401/C80MWX-X. We are withdrawing the notice... McCauley Propeller Systems for model propeller 3D15C1401/C80MWX-X (71 FR 43674). On November 29, 2004...

NEXT Single String Integration Test Results

NASA Technical Reports Server (NTRS)

Soulas, George C.; Patterson, Michael J.; Pinero, Luis; Herman, Daniel A.; Snyder, Steven John

2010-01-01

As a critical part of NASA's Evolutionary Xenon Thruster (NEXT) test validation process, a single string integration test was performed on the NEXT ion propulsion system. The objectives of this test were to verify that an integrated system of major NEXT ion propulsion system elements meets project requirements, to demonstrate that the integrated system is functional across the entire power processor and xenon propellant management system input ranges, and to demonstrate to potential users that the NEXT propulsion system is ready for transition to flight. Propulsion system elements included in this system integration test were an engineering model ion thruster, an engineering model propellant management system, an engineering model power processor unit, and a digital control interface unit simulator that acted as a test console. Project requirements that were verified during this system integration test included individual element requirements ; integrated system requirements, and fault handling. This paper will present the results of these tests, which include: integrated ion propulsion system demonstrations of performance, functionality and fault handling; a thruster re-performance acceptance test to establish baseline performance: a risk-reduction PMS-thruster integration test: and propellant management system calibration checks.

14 CFR 25.1027 - Propeller feathering system.

Code of Federal Regulations, 2014 CFR

2014-01-01

... AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Oil System § 25.1027 Propeller feathering system. (a) If the propeller feathering system depends on engine oil, there must be means to trap... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Propeller feathering system. 25.1027...

14 CFR 25.1027 - Propeller feathering system.

Code of Federal Regulations, 2011 CFR

2011-01-01

... AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Oil System § 25.1027 Propeller feathering system. (a) If the propeller feathering system depends on engine oil, there must be means to trap... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Propeller feathering system. 25.1027...

14 CFR 25.1027 - Propeller feathering system.

Code of Federal Regulations, 2010 CFR

2010-01-01

... AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Oil System § 25.1027 Propeller feathering system. (a) If the propeller feathering system depends on engine oil, there must be means to trap... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Propeller feathering system. 25.1027...

14 CFR 25.1027 - Propeller feathering system.

Code of Federal Regulations, 2012 CFR

2012-01-01

... AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Oil System § 25.1027 Propeller feathering system. (a) If the propeller feathering system depends on engine oil, there must be means to trap... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Propeller feathering system. 25.1027...

14 CFR 25.1027 - Propeller feathering system.

Code of Federal Regulations, 2013 CFR

2013-01-01

... AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Oil System § 25.1027 Propeller feathering system. (a) If the propeller feathering system depends on engine oil, there must be means to trap... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Propeller feathering system. 25.1027...

Laminated chemical and physical micro-jet actuators based on conductive media

NASA Astrophysics Data System (ADS)

Gadiraju, Priya D.

2008-04-01

This dissertation presents the development of electrically-powered, lamination-based microactuators for the realization of large arrays of high impulse and short duration micro-jets with potential applications in the field of micro-electro-mechanical systems (MEMS). Microactuators offer unique control opportunities by converting the input electrical or chemical energy stored in a propellant into useful mechanical energy. This small and precise control obtained can potentially be applied towards aerodynamic control and transdermal drug delivery applications. This thesis work discusses the feasibility of using microactuators for two such applications: Control of the motion of a spinning projectile by utilizing the chemically-driven microjets ejected from the actuators, and enhancement of the permeability properties of skin by selectively ablating the stratum corneum layer of skin using the physical microjets ejected from the actuators. This enhanced permeability of skin can later be used for the delivery of high molecular weight drugs for transdermal drug delivery. The development of electrically powered microactuators starts by fabricating an array of radially firing microactuators using lamination-based microfabrication techniques that potentially enable batch fabrication at low cost. The microactuators of this thesis consist of three main parts: a micro chamber in which the propellant is stored; two electrode structures through which electrical energy is supplied to the propellant; and a micro nozzle through which the propellant or released gases from the propellant are expanded as a jet. Once the actuators are fabricated, they are integrated with MEMS-process-compatible propellants and optimized so as to produce instantaneous ignition of the propellant. This instantaneous ignition is achieved either by making the propellant itself conductive, thus, passing an electric current directly through the propellant; or by discharging an arc across the propellant by placing it between two closely spaced electrodes. The first concept is demonstrated for the application of projectile maneuvering where energetic solid propellant is used in generating a high velocity gaseous jet and the second concept is demonstrated for transdermal drug delivery application where a rapid physical jet of a non-energetic propellant is generated. In the case of chemical-based microactuators, the feasibility of using conductive solid propellant based actuators for maneuvering a 25 mm bluff body projectile spinning at 600 Hz is presented. Several conductive solid propellants are developed and characterized for their electrical conductivity and required ignition energy. Finally, the propellant integrated microactuators are characterized for performance in terms of impulse delivered, thrust generated and duration of the jet. These experimental results are then compared to predicted results from simulations. In the case of physical based microactuators, the feasibility of using released physical jets from the microactuator array for transdermal drug delivery application is presented. Several bio-compatible and FDA-approved liquids are used as propellants and are characterized in terms of thrusts delivered and duration of the released jets. These thermo-mechanical jets are then used to expose skin locally so as to create micro conduits in the stratum corneum layer of skin. Both thermal effects and thermo-mechanical effects of the jet on exposed skin are studied. For both cases, histology of exposed skin is presented and its permeability to drug analog molecules is studied.

Fluid Distribution for In-space Cryogenic Propulsion

NASA Technical Reports Server (NTRS)

Lear, William

2005-01-01

The ultimate goal of this task is to enable the use of a single supply of cryogenic propellants for three distinct spacecraft propulsion missions: main propulsion, orbital maneuvering, and attitude control. A fluid distribution system is sought which allows large propellant flows during the first two missions while still allowing control of small propellant flows during attitude control. Existing research has identified the probable benefits of a combined thermal management/power/fluid distribution system based on the Solar Integrated Thermal Management and Power (SITMAP) cycle. Both a numerical model and an experimental model are constructed in order to predict the performance of such an integrated thermal management/propulsion system. This research task provides a numerical model and an experimental apparatus which will simulate an integrated thermal/power/fluid management system based on the SITMAP cycle, and assess its feasibility for various space missions. Various modifications are done to the cycle, such as the addition of a regeneration process that allows heat to be transferred into the working fluid prior to the solar collector, thereby reducing the collector size and weight. Fabri choking analysis was also accounted for. Finally the cycle is to be optimized for various space missions based on a mass based figure of merit, namely the System Mass Ratio (SMR). -. 1 he theoretical and experimental results from these models are be used to develop a design code (JETSIT code) which is able to provide design parameters for such a system, over a range of cooling loads, power generation, and attitude control thrust levels. The performance gains and mass savings will be compared to those of existing spacecraft systems.

Integrated Liquid Bismuth Propellant Feed System

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Markusic, Thomas E.; Stanojev, Boris J.

2006-01-01

A prototype bismuth propellant feed and control system was constructed and tested. An electromagnetic pump was used in this system to provide fine control of the hydrostatic pressure, and a new type of in-line flow sensor was developed to provide an accurate, real-time measurement of the mass flow rate. High-temperature material compatibility was a driving design requirement for the pump and flow sensor, leading to the selection of macor for the main body of both components. Post-test inspections of both components revealed no cracks or leaking in either. In separate proof-of-concept experiments, the pump produced a linear pressure rise as a function of current that compared favorably with theoretical pump pressure predictions, with a pressure of 10 kPa at 30 A. Flow sensing was successfully demonstrated in a bench-top test using gallium as a substitute liquid metal. A real-time controller was successfully used to control the entire system, simultaneously monitoring all power supplies and performing data acquisition duties.

Numerical study on the influence of boss cap fins on efficiency of controllable-pitch propeller

NASA Astrophysics Data System (ADS)

Xiong, Ying; Wang, Zhanzhi; Qi, Wanjiang

2013-03-01

Numerical simulation is investigated to disclose how propeller boss cap fins (PBCF) operate utilizing Reynolds-averaged Navier-Stokes (RANS) method. In addition, exploration of the influencing mechanism of PBCF on the open water efficiency of one controllable-pitch propeller is analyzed through the open water characteristic curves, blade surface pressure distribution and hub streamline distribution. On this basis, the influence of parameters including airfoil profile, diameter, axial position of installation and circumferential installation angle on the open water efficiency of the controllable-pitch propeller is investigated. Numerical results show: for the controllable-pitch propeller, the thrust generated is at the optimum when the radius of boss cap fins is 1.5 times of propeller hub with an optimal installation position in the axial direction, and its optimal circumferential installation position is the midpoint of the extension line of the front and back ends of two adjacent propeller roots in the front of fin root. Under these optimal parameters, the gain of open water efficiency of the controllable-pitch propeller with different advance velocity coefficients is greater than 0.01, which accounts for approximately an increase of 1%-5% of open water efficiency.

Computer-Controlled Image Anaysis of Solid Propellant Combustion Holograms Using a Quantimet 720 and a PDP-11.

DTIC Science & Technology

1985-09-01

TND 1 96 PIN11. L 4. c. j;. NAVAL POSTGRADUATE SCHOOL Monterey, California NOV 19 19853 THESIS COMPUTER-CONTROLLED IMAGE ANALYSIS OF SOLID PROPELLANT...Controlled Image Analysis of Master’s Thesis Solid Propellant Combustion Holograms September, 1985 Using a Quantimet 720 and a PDP-11 S. PERFORMING ORG...unlimited Computer-Controlled Image Analysis of Solid Propellant * - Combustion Holograms Using a Quantimet 720 and a PDP-11 by Marvin Philip Shook

Subcooling for Long Duration In-Space Cryogenic Propellant Storage

NASA Technical Reports Server (NTRS)

Mustafi, Shuvo; Johnson, Wesley; Kashani, Ali; Jurns, John; Kutter, Bernard; Kirk, Daniel; Shull, Jeff

2010-01-01

Cryogenic propellants such as hydrogen and oxygen are crucial for exploration of the solar system because of their superior specific impulse capability. Future missions may require vehicles to remain in space for months, necessitating long-term storage of these cryogens. A Thermodynamic Cryogen Subcooler (TCS) can ease the challenge of cryogenic fluid storage by removing energy from the cryogenic propellant through isobaric subcooling of the cryogen below its normal boiling point prior to launch. The isobaric subcooling of the cryogenic propellant will be performed by using a cold pressurant to maintain the tank pressure while the cryogen's temperature is simultaneously reduced using the TCS. The TCS hardware will be integrated into the launch infrastructure and there will be no significant addition to the launched dry mass. Heat leaks into all cryogenic propellant tanks, despite the use of the best insulation systems. However, the large heat capacity available in the subcooled cryogenic propellants allows the energy that leaks into the tank to be absorbed until the cryogen reaches its operational thermodynamic condition. During this period of heating of the subcooled cryogen there will be minimal loss of the propellant due to venting for pressure control. This simple technique can extend the operational life of a spacecraft or an orbital cryogenic depot for months with minimal mass penalty. In fact isobaric subcooling can more than double the in-space hold time of liquid hydrogen compared to normal boiling point hydrogen. A TCS for cryogenic propellants would thus provide an enhanced level of mission flexibility. Advances in the important components of the TCS will be discussed in this paper.

Development of a Ground Operations Demonstration Unit for Liquid Hydrogen at Kennedy Space Center

NASA Astrophysics Data System (ADS)

Notardonato, W. U.

NASA operations for handling cryogens in ground support equipment have not changed substantially in 50 years, despite major technology advances in the field of cryogenics. NASA loses approximately 50% of the hydrogen purchased because of a continuous heat leak into ground and flight vessels, transient chill down of warm cryogenic equipment, liquid bleeds, and vent losses. NASA Kennedy Space Center (KSC) needs to develop energy-efficient cryogenic ground systems to minimize propellant losses, simplify operations, and reduce cost associated with hydrogen usage. The GODU LH2 project will design, assemble, and test a prototype storage and distribution system for liquid hydrogen that represents an advanced end-to-end cryogenic propellant system for a ground launch complex. The project has multiple objectives and will culminate with an operational demonstration of the loading of a simulated flight tank with densified propellants. The system will be unique because it uses an integrated refrigeration and storage system (IRAS) to control the state of the fluid. The integrated refrigerator is the critical feature enabling the testing of the following three functions: zero-loss storage and transfer, propellant densification/conditioning, and on-site liquefaction. This paper will discuss the test objectives, the design of the system, and the current status of the installation.

Tank System Integrated Model: A Cryogenic Tank Performance Prediction Program

NASA Technical Reports Server (NTRS)

Bolshinskiy, L. G.; Hedayat, A.; Hastings, L. J.; Sutherlin, S. G.; Schnell, A. R.; Moder, J. P.

2017-01-01

Accurate predictions of the thermodynamic state of the cryogenic propellants, pressurization rate, and performance of pressure control techniques in cryogenic tanks are required for development of cryogenic fluid long-duration storage technology and planning for future space exploration missions. This Technical Memorandum (TM) presents the analytical tool, Tank System Integrated Model (TankSIM), which can be used for modeling pressure control and predicting the behavior of cryogenic propellant for long-term storage for future space missions. Utilizing TankSIM, the following processes can be modeled: tank self-pressurization, boiloff, ullage venting, mixing, and condensation on the tank wall. This TM also includes comparisons of TankSIM program predictions with the test data andexamples of multiphase mission calculations.

Propeller torque load and propeller shaft torque response correlation during ice-propeller interaction

NASA Astrophysics Data System (ADS)

Polić, Dražen; Ehlers, Sören; Æsøy, Vilmar

2017-03-01

Ships use propulsion machinery systems to create directional thrust. Sailing in ice-covered waters involves the breaking of ice pieces and their submergence as the ship hull advances. Sometimes, submerged ice pieces interact with the propeller and cause irregular fluctuations of the torque load. As a result, the propeller and engine dynamics become imbalanced, and energy propagates through the propulsion machinery system until equilibrium is reached. In such imbalanced situations, the measured propeller shaft torque response is not equal to the propeller torque. Therefore, in this work, the overall system response is simulated under the ice-related torque load using the Bond graph model. The energy difference between the propeller and propeller shaft is estimated and related to their corresponding mechanical energy. Additionally, the mechanical energy is distributed among modes. Based on the distribution, kinetic and potential energy are important for the correlation between propeller torque and propeller shaft response.

Numerical modeling of self-pressurization and pressure control by a thermodynamic vent system in a cryogenic tank

NASA Astrophysics Data System (ADS)

Majumdar, Alok; Valenzuela, Juan; LeClair, Andre; Moder, Jeff

2016-03-01

This paper presents a numerical model of a system-level test bed-the multipurpose hydrogen test bed (MHTB) using the Generalized Fluid System Simulation Program (GFSSP). MHTB is representative in size and shape of a space transportation vehicle liquid hydrogen propellant tank, and ground-based testing was performed at NASA Marshall Space Flight Center (MSFC) to generate data for cryogenic storage. GFSSP is a finite volume-based network flow analysis software developed at MSFC and used for thermofluid analysis of propulsion systems. GFSSP has been used to model the self-pressurization and ullage pressure control by the Thermodynamic Vent System (TVS). A TVS typically includes a Joule-Thompson (J-T) expansion device, a two-phase heat exchanger (HEX), and a mixing pump and liquid injector to extract thermal energy from the tank without significant loss of liquid propellant. For the MHTB tank, the HEX and liquid injector are combined into a vertical spray bar assembly. Two GFSSP models (Self-Pressurization and TVS) were separately developed and tested and then integrated to simulate the entire system. The Self-Pressurization model consists of multiple ullage nodes, a propellant node, and solid nodes; it computes the heat transfer through multilayer insulation blankets and calculates heat and mass transfer between the ullage and liquid propellant and the ullage and tank wall. A TVS model calculates the flow through a J-T valve, HEX, and spray and vent systems. Two models are integrated by exchanging data through User Subroutines of both models. Results of the integrated models have been compared with MHTB test data at a 50% fill level. Satisfactory comparison was observed between tests and numerical predictions.

Magnetically Actuated Propellant Orientation, Controlling Fluids in a Low-Gravity Environment

NASA Technical Reports Server (NTRS)

Martin, James J.; Holt, James B.

2000-01-01

Cryogenic fluid management (CFM) is a technology area common to virtually every space transportation propulsion concept envisioned. Storage, supply, transfer and handling of sub-critical cryogenic fluids are basic capabilities that have long been needed by multiple programs and the need is expected to continue in the future. The use of magnetic fields provides another method, which could replace or augment current/traditional approaches, potentially simplifying vehicle operational constraints. The magnetically actuated propellant orientation (MAPO) program effort focused on the use of magnetic fields to control fluid motion as it relates to positioning (i.e. orientation and acquisition) of a paramagnetic substance such as LO2. Current CFM state- of-the-art systems used to control and acquire propellant in low gravity environments rely on liquid surface tension devices which employ vanes, fine screen mesh channels and baskets. These devices trap and direct propellant to areas where it's needed and have been used routinely with storable (non-cryogenic) propellants. However, almost no data exists r,egarding their operation in cryogenics and the use of such devices confronts designers with a multitude of significant technology issues. Typical problems include a sensitivity to screen dry out (due to thermal loads and pressurant gas) and momentary adverse accelerations (generated from either internal or external sources). Any of these problems can potentially cause the acquisition systems to ingest or develop vapor and fail. The use of lightweight high field strength magnets may offer a valuable means of augmenting traditional systems potentially mitigating or at least easing operational requirements. Two potential uses of magnetic fields include: 1) strategically positioning magnets to keep vent ports clear of liquid (enabling low G vented fill operations), and 2) placing magnets in the center or around the walls of the tank to create an insulating vapor pocket (between the liquid and the tank wall) which could effectively lower heat transfer to the liquid (enabling increased storage time).

Colloid Microthruster Feed System Development for Fine Pointing and Drag-Free Control of Multi-Year Astronomical Observatories

NASA Astrophysics Data System (ADS)

Ziemer, John; Mueller, J.; Spence, D.; Hruby, V.

2014-01-01

A new Colloid Microthruster feed system, including a propellant tank and redundant Microvalves, is being developed for fine pointing and drag-free operations of multi-year astronomical observatories under the PCOS SAT program. Almost all Gravitational Wave Observatory (GWO) concepts require microthrusters to maintain a drag-free environment for the inertial sensor instrument to meet the mission science objectives. The current state-of-the-art microthruster in the US is the Busek Colloid Micro-Newton Thruster (CMNT) originally developed under the New Millennium Program for the Space Technology 7 (ST7) and ESA's LISA Pathfinder (LPF) technology demonstration mission. The ST7 CMNT design includes a bellows propellant storage tank that is sized to provide up to 90 days of maximum thrust (30 µN). The new propellant tank is based on a blow-down, metal-diaphragm spherical tank design with enough capacity for a 5-year GWO mission. The new feed system will also include the third generation of Busek’s Microvalve, currently being developed under a NASA Phase II SBIR. The Microvalve is responsible for the picoliter per second control of the propellant from the tank to the thruster head, demanding parts with micron-level tolerances, critical alignments, and challenging acceptance test protocols. This microthruster system could also be considered for replacement of reaction wheels for slewing and fine pointing of other astronomical observatories, including Exo-Planet Observatory concepts. The goal of the PCOS SAT effort is to raise the new system to TRL 5 with performance and environmental testing within the next two years.

Swimming droplets driven by a surface wave

PubMed Central

Ebata, Hiroyuki; Sano, Masaki

2015-01-01

Self-propelling motion is ubiquitous for soft active objects such as crawling cells, active filaments, and liquid droplets moving on surfaces. Deformation and energy dissipation are required for self-propulsion of both living and non-living matter. From the perspective of physics, searching for universal laws of self-propelled motions in a dissipative environment is worthwhile, regardless of the objects' details. In this article, we propose a simple experimental system that demonstrates spontaneous migration of a droplet under uniform mechanical agitation. As we vary control parameters, spontaneous symmetry breaking occurs sequentially, and cascades of bifurcations of the motion arise. Equations describing deformable particles and hydrodynamic simulations successfully describe all of the observed motions. This system should enable us to improve our understanding of spontaneous motions of self-propelled objects. PMID:25708871

Swimming droplets driven by a surface wave

NASA Astrophysics Data System (ADS)

Ebata, Hiroyuki; Sano, Masaki

2015-02-01

Self-propelling motion is ubiquitous for soft active objects such as crawling cells, active filaments, and liquid droplets moving on surfaces. Deformation and energy dissipation are required for self-propulsion of both living and non-living matter. From the perspective of physics, searching for universal laws of self-propelled motions in a dissipative environment is worthwhile, regardless of the objects' details. In this article, we propose a simple experimental system that demonstrates spontaneous migration of a droplet under uniform mechanical agitation. As we vary control parameters, spontaneous symmetry breaking occurs sequentially, and cascades of bifurcations of the motion arise. Equations describing deformable particles and hydrodynamic simulations successfully describe all of the observed motions. This system should enable us to improve our understanding of spontaneous motions of self-propelled objects.

Baseline spacecraft and mission design for the SP-100 flight experiment

NASA Technical Reports Server (NTRS)

Deininger, William D.; Vondra, Robert J.

1989-01-01

The design and performance of a spacecraft employing arcjet nuclear electric propulsion, suitable for use in the SP-100 Space Reactor Power System (SRPS) Flight Experiment, are outlined. The vehicle design is based on a 93 kWe ammonia arcjet system operating at an experimentally-measured specific impulse of 1030 s and an efficiency of 42 percent. The arcjet/gimbal assemblies, power conditioning subsystem, propellant feed system, propulsion system thermal control, spacecraft diagnostic instrumentation, and the telemetry requirements are described. A 100 kWe SRPS is assumed. The total spacecraft mass is baselined at 5675 kg excluding the propellant and propellant feed system. Four mission scenarios are described which are capable of demonstrating the full capability of the SRPS. The missions considered include spacecraft deployment to possible surveillance platform orbits, a spacecraft storage mission and an orbit raising round trip corresponding to possible orbit transfer vehicle missions. Launches from Kennedy Space Center using the Titan IV expendable launch vehicle are assumed.

Design and Development of a Salbutamol Intake Detector for Low Respiratory Treatment

NASA Astrophysics Data System (ADS)

Vui Hin, Tsen; Ilyani Ramli, Nur

2017-08-01

This paper proposed a new salbutamol intake detector design using asthma spacer and gas sensor. The device enable real time monitoring of propellant level inhaled by the infant which will decrease the recovery time of the asthma attack. Microcontroller Arduino UNO is program to control the input and output of the system. MQ6 gas sensor detecting the propellant Hydrofluoroalkane from the metered dose inhaler (MDI) canister and demonstrated the level of propellant inhaled on the LCD in real time. MQ6 gas sensor suitable used to detect concentration of propellant inside the asthma spacer due to it is low sensitive to natural gas where include the carbon dioxide exhaled by the infant. Besides this, MQ6 gas sensor also highly sensitive to propane and the preview aerosol inventor mentioned propane as propellant which used for MDI to push the salbutamol out from MDI canister. Therefore, MQ6 gas sensor is suitable to detect propellant inside asthma spacer. The output voltage of MQ6 in initial state where no propellant inside asthma spacer is between 0.55V and 0.65V. Furthermore, when the MDI canister is been pressed, the concentration of propellant is increased and the output voltage of MQ6 gas sensor also increased in ranged between 1.1V and 1.2V.

Giant number fluctuations in self-propelled particles without alignment

NASA Astrophysics Data System (ADS)

Fily, Yaouen; Henkes, Silke; Marchetti, M. Cristina

2012-02-01

Giant number fluctuations are a ubiquitous property of active systems. They were predicted using a generic continuum description of active nematics, and have been observed in simulations of Vicsek-type models and in experiments on vibrated granular layers and swimming bacteria. In all of these systems, there is an alignment interaction among the self-propelled units, either imposed as a rule, or arising from hydrodynamic or other medium-mediated couplings. Here we report numerical evidence of giant number fluctuations in a minimal model of self-propelled disks in two dimensions in the absence of any alignment mechanism. The direction of self-propulsion evolves via rotational diffusion and the particles interact solely via a finite range repulsive soft potential. It can be shown that in this system self propulsion is equivalent to a non Markovian noise whose correlation time is controlled by the amplitude of the orientational noise.

Evolution of the 1-mlb mercury ion thruster subsystem

NASA Technical Reports Server (NTRS)

Kerslake, W. R.; Banks, B. A.

1978-01-01

The developmental history, performance, and major lifetests of each component of the present 1-mlb (4.5 mN) thruster system are traced over the past 10 years. The 1-mlb thruster subsystem consists of an 8 cm diameter ion thruster mounted on 2 axis gimbals, a mercury propellant tank, a power electronics unit, a controller/digital interface unit, and necessary electrical harnesses plus propellant tankage and feed lines.

Hydrogen peroxide propulsion for smaller satellites (SSC98-VIII-1)

DOE Office of Scientific and Technical Information (OSTI.GOV)

Whitehead, J C

1998-07-13

As satellite designs shrink, providing maneuvering and control capability falls outside the realm of available propulsion technology. While cold gas has been used on the smallest satellites, hydrogen peroxide propellant is suggested as the next step in performance and cost before hydrazine. Minimal toxicity and a small scale enable benchtop propellant preparation and development testing. Progress toward low-cost thrusters and self-pressurizing tank systems is described.

14 CFR 25.1155 - Reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Reverse thrust and propeller pitch settings... Powerplant Controls and Accessories § 25.1155 Reverse thrust and propeller pitch settings below the flight regime. Each control for reverse thrust and for propeller pitch settings below the flight regime must...

14 CFR 25.1155 - Reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Reverse thrust and propeller pitch settings... Powerplant Controls and Accessories § 25.1155 Reverse thrust and propeller pitch settings below the flight regime. Each control for reverse thrust and for propeller pitch settings below the flight regime must...

14 CFR 25.1155 - Reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Reverse thrust and propeller pitch settings... Powerplant Controls and Accessories § 25.1155 Reverse thrust and propeller pitch settings below the flight regime. Each control for reverse thrust and for propeller pitch settings below the flight regime must...

14 CFR 25.1155 - Reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Reverse thrust and propeller pitch settings... Powerplant Controls and Accessories § 25.1155 Reverse thrust and propeller pitch settings below the flight regime. Each control for reverse thrust and for propeller pitch settings below the flight regime must...

14 CFR 25.1155 - Reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Reverse thrust and propeller pitch settings... Powerplant Controls and Accessories § 25.1155 Reverse thrust and propeller pitch settings below the flight regime. Each control for reverse thrust and for propeller pitch settings below the flight regime must...

Anticipated Effectiveness of Active Noise Control in Propeller Aircraft Interiors as Determined by Sound Quality Tests

NASA Technical Reports Server (NTRS)

Powell, Clemans A.; Sullivan, Brenda M.

2004-01-01

Two experiments were conducted, using sound quality engineering practices, to determine the subjective effectiveness of hypothetical active noise control systems in a range of propeller aircraft. The two tests differed by the type of judgments made by the subjects: pair comparisons in the first test and numerical category scaling in the second. Although the results of the two tests were in general agreement that the hypothetical active control measures improved the interior noise environments, the pair comparison method appears to be more sensitive to subtle changes in the characteristics of the sounds which are related to passenger preference.

14. INTERIOR VIEW OF PROPELLER STAND CONTROL ROOM. WrightPatterson ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

14. INTERIOR VIEW OF PROPELLER STAND CONTROL ROOM. - Wright-Patterson Air Force Base, Area B, Building No. 20A, Propeller Test Complex, Seventh Street, from E to G Streets, Dayton, Montgomery County, OH

Integrated Pressure-Fed Liquid Oxygen / Methane Propulsion Systems - Morpheus Experience, MARE, and Future Applications

NASA Technical Reports Server (NTRS)

Hurlbert, Eric; Morehead, Robert; Melcher, John C.; Atwell, Matt

2016-01-01

An integrated liquid oxygen (LOx) and methane propulsion system where common propellants are fed to the reaction control system and main engines offers advantages in performance, simplicity, reliability, and reusability. LOx/Methane provides new capabilities to use propellants that are manufactured on the Mars surface for ascent return and to integrate with power and life support systems. The clean burning, non-toxic, high vapor pressure propellants provide significant advantages for reliable ignition in a space vacuum, and for reliable safing or purging of a space-based vehicle. The NASA Advanced Exploration Systems (AES) Morpheus lander demonstrated many of these key attributes as it completed over 65 tests including 15 flights through 2014. Morpheus is a prototype of LOx/Methane propellant lander vehicle with a fully integrated propulsion system. The Morpheus lander flight demonstrations led to the proposal to use LOx/Methane for a Discovery class mission, named Moon Aging Regolith Experiment (MARE) to land an in-situ science payload for Southwest Research Institute on the Lunar surface. Lox/Methane is extensible to human spacecraft for many transportation elements of a Mars architecture. This paper discusses LOx/Methane propulsion systems in regards to trade studies, the Morpheus project experience, the MARE NAVIS (NASA Autonomous Vehicle for In-situ Science) lander, and future possible applications. The paper also discusses technology research and development needs for Lox/Methane propulsion systems.

X-33 Attitude Control System Design for Ascent, Transition, and Entry Flight Regimes

NASA Technical Reports Server (NTRS)

Hall, Charles E.; Gallaher, Michael W.; Hendrix, Neal D.

1998-01-01

The Vehicle Control Systems Team at Marshall Space Flight Center, Systems Dynamics Laboratory, Guidance and Control Systems Division is designing under a cooperative agreement with Lockheed Martin Skunkworks, the Ascent, Transition, and Entry flight attitude control system for the X-33 experimental vehicle. Ascent flight control begins at liftoff and ends at linear aerospike main engine cutoff (NECO) while Transition and Entry flight control begins at MECO and concludes at the terminal area energy management (TAEM) interface. TAEM occurs at approximately Mach 3.0. This task includes not only the design of the vehicle attitude control systems but also the development of requirements for attitude control system components and subsystems. The X-33 attitude control system design is challenged by a short design cycle, the design environment (Mach 0 to about Mach 15), and the X-33 incremental test philosophy. The X-33 design-to-launch cycle of less than 3 years requires a concurrent design approach while the test philosophy requires design adaptation to vehicle variations that are a function of Mach number and mission profile. The flight attitude control system must deal with the mixing of aerosurfaces, reaction control thrusters, and linear aerospike engine control effectors and handle parasitic effects such as vehicle flexibility and propellant sloshing from the uniquely shaped propellant tanks. The attitude control system design is, as usual, closely linked to many other subsystems and must deal with constraints and requirements from these subsystems.

Transitions induced by speed in self-propelled particles system with attractive interactions

NASA Astrophysics Data System (ADS)

Cambui, Dorilson. S.; Rosas, Alexandre

2018-05-01

In this work, we consider a system of self-propelled particles with attractive interactions in two dimensions. The model presents an order-disorder transition with the speed playing the role of the control parameter. In order to characterize the transition, we investigate the behavior of the order parameter and the Binder cumulant as a function of the speed. Our main finding is that the transition can be either continuous or discontinuous depending on two parameter of the model: the strength of the noise and the radius of attraction.

An analysis of bipropellant neutralization for spacecraft refueling operations

NASA Technical Reports Server (NTRS)

Kauffman, David

1987-01-01

Refueling of satellites on orbit with storable propellants will involve venting part or all of the pressurant gas from the propellant tanks. This gas will be saturated with propellant vapor, and it may also have significant amounts of entrained fine droplets of propellant. The two most commonly used bipropellants, monomethyl hydrazine (MMH) and nitrogen tetroxide (N2O4), are highly reactive and toxic. Various possible ways of neutralizing the vented propellants are examined. The amount of propellant vented in a typical refueling operation is shown to be in the range of 0.2 to 5% of the tank capacity. Four potential neutralization schemes are examined: chemical decomposition, chemical reaction, condensation and adsorption. Chemical decomposition to essentially inert materials is thermodynamically feasible for both MMH and N2O4. It would be the simplest and easiest neutralization method to implement. Chemical decomposition would require more complex control. Condensation would require a refrigeration system and a very efficent phase separator. Adsorption is likely to be much heavier. A preliminary assessment of the four neutralization shemes is presented, along with suggested research and development plans.

Electronic circuitry development in a micropyrotechnic system for micropropulsion applications

NASA Astrophysics Data System (ADS)

Puig-Vidal, Manuel; Lopez, Jaime; Miribel, Pere; Montane, Enric; Lopez-Villegas, Jose M.; Samitier, Josep; Rossi, Carole; Camps, Thierry; Dumonteuil, Maxime

2003-04-01

An electronic circuitry is proposed and implemented to optimize the ignition process and the robustness of a microthruster. The principle is based on the integration of propellant material within a micromachined system. The operational concept is simply based on the combustion of an energetic propellant stored in a micromachined chamber. Each thruster contains three parts (heater, chamber, nozzle). Due to the one shot characteristic, microthrusters are fabricated in 2D array configuration. For the functioning of this kind of system, one critical point is the optimization of the ignition process as a function of the power schedule delivered by electronic devices. One particular attention has been paid on the design and implementation of an electronic chip to control and optimize the system ignition. Ignition process is triggered by electrical power delivered to a polysilicon resistance in contact with the propellant. The resistance is used to sense the temperature on the propellant which is in contact. Temperature of the microthruster node before the ignition is monitored via the electronic circuitry. A pre-heating process before ignition seems to be a good methodology to optimize the ignition process. Pre-heating temperature and pre-heating time are critical parameters to be adjusted. Simulation and experimental results will deeply contribute to improve the micropyrotechnic system. This paper will discuss all these point.

EXPERIMENTAL PERFORMANCE OF A CONTROLLABLE-PITCH SUPERCAVITATING PROPELLER.

DTIC Science & Technology

Studies were made of cavitation performance and open-water characteristics of a controllable-pitch supercavitating propeller with two, three, and...By means of several numerical examples, the feasibility of using a controllable-pitch supercavitating propeller is demonstrated. A practical application to a hydrofoil boat is also presented. (Author)

Potential Subjective Effectiveness of Active Interior Noise Control in Propeller Airplanes

NASA Technical Reports Server (NTRS)

Powell, Clemans A.; Sullivan, Brenda M.

2000-01-01

Active noise control technology offers the potential for weight-efficient aircraft interior noise reduction, particularly for propeller aircraft. However, there is little information on how passengers respond to this type of interior noise control. This paper presents results of two experiments that use sound quality engineering practices to determine the subjective effectiveness of hypothetical active noise control (ANC) systems in a range of propeller aircraft. The two experiments differed by the type of judgments made by the subjects: pair comparisons based on preference in the first and numerical category scaling of noisiness in the second. Although the results of the two experiments were in general agreement that the hypothetical active control measures improved the interior noise environments, the pair comparison method appears to be more sensitive to subtle changes in the characteristics of the sounds which are related to passenger preference. The reductions in subjective response due to the ANC conditions were predicted with reasonable accuracy by reductions in measured loudness level. Inclusion of corrections for the sound quality characteristics of tonality and fluctuation strength in multiple regression models improved the prediction of the ANC effects.

Motor Controllers

NASA Technical Reports Server (NTRS)

1984-01-01

Kollmorgen Corporation's Mermaid II two person submersible is propeller-driven by a system of five DC brushless motors with new electronic controllers that originated in work performed in a NASA/DOE project managed by Lewis Research Center. A key feature of the system is electric commutation rather than mechanical commutation for converting AC current to DC.

An evaluation of oxygen-hydrogen propulsion systems for the Space Station

NASA Technical Reports Server (NTRS)

Klemetson, R. W.; Garrison, P. W.; Hannum, N. P.

1985-01-01

Conceptual designs for O2/H2 chemical and resistojet propulsion systems for the space station was developed and evaluated. The evolution of propulsion requirements was considered as the space station configuration and its utilization as a space transportation node change over the first decade of operation. The characteristics of candidate O2/H2 auxiliary propulsion systems are determined, and opportunities for integration with the OTV tank farm and the space station life support, power and thermal control subsystems are investigated. OTV tank farm boiloff can provide a major portion of the growth station impulse requirements and CO2 from the life support system can be a significant propellant resource, provided it is not denied by closure of that subsystem. Waste heat from the thermal control system is sufficient for many propellant conditioning requirements. It is concluded that the optimum level of subsystem integration must be based on higher level space station studies.

Liquid Nitrogen Zero Boiloff Testing

NASA Technical Reports Server (NTRS)

Plachta, David; Feller, Jeffrey; Johnson, Wesley; Robinson, Craig

2017-01-01

Cryogenic propellants such as liquid hydrogen (LH2) and liquid oxygen (LO2) are a part of NASAs future space exploration due to their high specific impulse for rocket motors of upper stages suitable for transporting 10s to 100s of metric tons of payload mass to destinations outside of low earth orbit and for their return. However, the low storage temperatures of LH2 and LO2 cause substantial boil-off losses for missions with durations greater than several months. These losses can be eliminated by incorporating high performance cryocooler technology to intercept heat load to the propellant tanks and modulating the cryocooler to control tank pressure. The active thermal control technology being developed by NASA is the reverse turbo-Brayton cycle cryocooler and its integration to the propellant tank through a distributed cooling tubing network coupled to the tank wall. This configuration was recently tested at NASA Glenn Research Center, in a vacuum chamber and cryo-shroud that simulated the essential thermal aspects of low Earth orbit, its vacuum and temperature. Testing consisted of three passive tests with the active cryo-cooler system off, and 7 active tests, with the cryocooler powered up. The test matrix included zero boil-off tests performed at 90 full and 25 full, and several demonstrations at excess cooling capacity and reduced cooling capacity. From this, the tank pressure response with varied cryocooler power inputs was determined. This test series established that the active cooling system integrated with the propellant tank eliminated boil-off and robustly controlled tank pressure.

Ship electric propulsion simulator based on networking technology

NASA Astrophysics Data System (ADS)

Zheng, Huayao; Huang, Xuewu; Chen, Jutao; Lu, Binquan

2006-11-01

According the new ship building tense, a novel electric propulsion simulator (EPS) had been developed in Marine Simulation Center of SMU. The architecture, software function and FCS network technology of EPS and integrated power system (IPS) were described. In allusion to the POD propeller in ship, a special physical model was built. The POD power was supplied from the simulative 6.6 kV Medium Voltage Main Switchboard, its control could be realized in local or remote mode. Through LAN, the simulated feature information of EPS will pass to the physical POD model, which would reflect the real thruster working status in different sea conditions. The software includes vessel-propeller math module, thruster control system, distribution and emergency integrated management, double closed loop control system, vessel static water resistance and dynamic software; instructor main control software. The monitor and control system is realized by real time data collection system and CAN bus technology. During the construction, most devices such as monitor panels and intelligent meters, are developed in lab which were based on embedded microcomputer system with CAN interface to link the network. They had also successfully used in practice and would be suitable for the future demands of digitalization ship.

Controlling Gas-Flow Mass Ratios

NASA Technical Reports Server (NTRS)

Morris, Brian G.

1990-01-01

Proposed system automatically controls proportions of gases flowing in supply lines. Conceived for control of oxidizer-to-fuel ratio in new gaseous-propellant rocket engines. Gas-flow control system measures temperatures and pressures at various points. From data, calculates control voltages for electronic pressure regulators for oxygen and hydrogen. System includes commercially available components. Applicable to control of mass ratios in such gaseous industrial processes as chemical-vapor depostion of semiconductor materials and in automotive engines operating on compressed natural gas.

Low-thrust chemical orbit to orbit propulsion system propellant management study

NASA Technical Reports Server (NTRS)

Dergance, R. H.

1980-01-01

Propellant requirements, tankage configurations, preferred propellant management techniques, propulsion systems weights, and technology deficiencies for low thrust expendable propulsion systems are examined. A computer program was utilized which provided a complete propellant inventory (including boil-off for cryogenic cases), pressurant and propellant tank dimensions for a given ullage, pressurant requirements, insulation requirements, and miscellaneous masses. The output also includes the masses of all tanks; the mass of the insulation, engines and other components; total wet system and burnout mass; system mass fraction; total impulse and burn time.

A review of electron bombardment thruster systems/spacecraft field and particle interfaces

NASA Technical Reports Server (NTRS)

Byers, D. C.

1978-01-01

Information on the field and particle interfaces of electron bombardment ion thruster systems was summarized. Major areas discussed were the nonpropellant particles, neutral propellant, ion beam, low energy plasma, and fields. Spacecraft functions and subsystems reviewed were solar arrays, thermal control systems, optical sensors, communications, science, structures and materials, and potential control.

Propulsion System Testing for the Iodine Satellite (iSAT) Demonstration Mission

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Kamhawi, Hani

2015-01-01

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cm cu and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high delta v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ?Isp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (less than 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp -6) torr at -75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. An iodine-based system is not without its challenges. The primary challenge is that the entire feed system must be maintained at an elevated temperature to prevent the iodine from depositing (transitioning from the gas phase directly back into the solid phase), which will block the propellant feed lines. Furthermore, deposition will occur unless the temperature in the lines is not greater than the temperature of the propellant reservoir. The flow rate can be controlled by adjusting the heating applied to the reservoir, but as with any thermal control there is a relatively slow response to changes in the heating rate. In the present paper, we describe the propulsion and propellant feed system for the iodine satellite (iSAT) flight demonstration mission. The system is based around the Busek BHT-200 Hall thruster, which has been modified for chemical compatibility with iodine vapor. While the gross propellant flow rate is maintained by the heated propellant reservoir, the flow to the anode and cathode are adjusted using two heated Vacco proportional flow control valves (PFCV), which provide very fast response on the flow rate adjustment. The flight mission design layout will be presented, showing how the system will be packaged into the overall 12-U spacecraft and the techniques being employed to protect the remaining spacecraft hardware from the propulsion system (e.g., plasma impingement, iodine deposition, thermal loads). In addition to the flight system design, results of testing the thruster and cathode with both operating on iodine propellant are presented. The tests are conducted on a thrust stand (see Fig. 1) in a large vacuum chamber containing a beam dump chilled to below -100 C to 'cryopump' the propellant. The thruster performance during these tests is presented, with these data used to evaluate the feed system and guide further refinements. Results of relatively long duration testing are presented to demonstrate the capability to operate for the length of the iSAT mission and to perform a number of re-starts as will be required by the mission concept of operations.

Reduced Toxicity Fuel Satellite Propulsion System

NASA Technical Reports Server (NTRS)

Schneider, Steven J. (Inventor)

2001-01-01

A reduced toxicity fuel satellite propulsion system including a reduced toxicity propellant supply for consumption in an axial class thruster and an ACS class thruster. The system includes suitable valves and conduits for supplying the reduced toxicity propellant to the ACS decomposing element of an ACS thruster. The ACS decomposing element is operative to decompose the reduced toxicity propellant into hot propulsive gases. In addition the system includes suitable valves and conduits for supplying the reduced toxicity propellant to an axial decomposing element of the axial thruster. The axial decomposing element is operative to decompose the reduced toxicity propellant into hot gases. The system further includes suitable valves and conduits for supplying a second propellant to a combustion chamber of the axial thruster, whereby the hot gases and the second propellant auto-ignite and begin the combustion process for producing thrust.

Reduced Toxicity Fuel Satellite Propulsion System Including Plasmatron

NASA Technical Reports Server (NTRS)

Schneider, Steven J. (Inventor)

2003-01-01

A reduced toxicity fuel satellite propulsion system including a reduced toxicity propellant supply for consumption in an axial class thruster and an ACS class thruster. The system includes suitable valves and conduits for supplying the reduced toxicity propellant to the ACS decomposing element of an ACS thruster. The ACS decomposing element is operative to decompose the reduced toxicity propellant into hot propulsive gases. In addition the system includes suitable valves and conduits for supplying the reduced toxicity propellant to an axial decomposing element of the axial thruster. The axial decomposing element is operative to decompose the reduced toxicity propellant into hot gases. The system further includes suitable valves and conduits for supplying a second propellant to a combustion chamber of the axial thruster. whereby the hot gases and the second propellant auto-ignite and begin the combustion process for producing thrust.

Method for CFD Simulation of Propellant Slosh in a Spherical Tank

NASA Technical Reports Server (NTRS)

Benson, David J.; Mason, Paul A.

2011-01-01

Propellant sloshing can impart unwanted disturbances to spacecraft, especially if the spacecraft controller is driving the system at the slosh frequency. This paper describes the work performed by the authors in simulating propellant slosh in a spherical tank using computational fluid dynamics (CFD). ANSYS-CFX is the CFD package used to perform the analysis. A 42 in spherical tank is studied with various fill fractions. Results are provided for the forces on the walls and the frequency of the slosh. Snapshots of slosh animation give a qualitative understanding of the propellant slosh. The results show that maximum slosh forces occur at a tank fill fraction of 0.4 and 0.6 due to the amount of mass participating in the slosh and the room available for sloshing to occur. The slosh frequency increases as the tank fill fraction increases.

Validation of Slosh Modeling Approach Using STAR-CCM+

NASA Technical Reports Server (NTRS)

Benson, David J.; Ng, Wanyi

2018-01-01

Without an adequate understanding of propellant slosh, the spacecraft attitude control system may be inadequate to control the spacecraft or there may be an unexpected loss of science observation time due to higher slosh settling times. Computational fluid dynamics (CFD) is used to model propellant slosh. STAR-CCM+ is a commercially available CFD code. This paper seeks to validate the CFD modeling approach via a comparison between STAR-CCM+ liquid slosh modeling results and experimental, empirically, and analytically derived results. The geometries examined are a bare right cylinder tank and a right cylinder with a single ring baffle.

Optimal Navigation of Self-Propelled Colloids in Microstructured Mazes

NASA Astrophysics Data System (ADS)

Yang, Yuguang; Bevan, Michael

Controlling navigation of self-propelled microscopic `robots' subject to random Brownian motion in complex microstructured environments (e.g., porous media, tumor vasculature) is important to many emerging applications (e.g., enhanced oil recovery, drug delivery). In this work, we design an optimal feedback policy to navigate an active self-propelled colloidal rod in complex mazes with various obstacle types. Actuation of the rods is modelled based on a light-controlled osmotic flow mechanism, which produces different propulsion velocities along the rod's long axis. Actuator-parameterized Langevin equations, with soft rod-obstacle repulsive interactions, are developed to describe the system dynamics. A Markov decision process (MDP) framework is used for optimal policy calculations with design goals of colloidal rods reaching target end points in minimum time. Simulations show that optimal MDP-based policies are able to control rod trajectories to reach target regions order-of-magnitudes faster than uncontrolled rods, which diverges as maze complexity increases. An efficient multi-graph based implementation for MDP is also presented, which scales linearly with the maze dimension.

Potential biofouling of spacecraft propellant systems due to contaminated deionized water

NASA Astrophysics Data System (ADS)

Hogue, Patrick

2006-08-01

Deionized (DI) water, with a density close to hydrazine, is used to fill spacecraft propellant tanks for mechanical testing during ground operations, after which is it removed and the tanks dried for use with anhydrous hydrazine. Pure nitrogen is used as a pressurant during storage and during water fill and drain operations. Since DI water systems are notorious for contamination by slime-forming bacteria, DI water intended for use in New Horizons and STEREO hydrazine tanks at APL was assessed for microorganism content using the heterotrophic plate count (HPC) method. Results show that some growth occurred during storage of DI water in propellant tanks, however not at the logarithmic rate associated with well-nourished bacteria. Ralstonia and Burkholderia were present in DI water on-loaded however only Ralstonia was present in off-loaded water. One possible source of nutrients during water storage in propellant tanks is organic material originating from the EPDM (EPR per AF-E-332) expulsion diaphragm. This paper will demonstrate potential for bio-fouling of spacecraft propulsion systems due to growth of slime-forming bacteria and will suggest that specifications controlling microorganism content should be imposed on water used for spacecraft ground testing.

Safety Practices Followed in ISRO Launch Complex- An Overview

NASA Astrophysics Data System (ADS)

Krishnamurty, V.; Srivastava, V. K.; Ramesh, M.

2005-12-01

The spaceport of India, Satish Dhawan Space Centre (SDSC) SHAR of Indian Space Research Organisation (ISRO), is located at Sriharikota, a spindle shaped island on the east coast of southern India.SDSC SHAR has a unique combination of facilities, such as a solid propellant production plant, a rocket motor static test facility, launch complexes for different types of rockets, telemetry, telecommand, tracking, data acquisition and processing facilities and other support services.The Solid Propellant Space Booster Plant (SPROB) located at SDSC SHAR produces composite solid propellant for rocket motors of ISRO. The main ingredients of the propellant produced here are ammonium perchlorate (oxidizer), fine aluminium powder (fuel) and hydroxyl terminated polybutadiene (binder).SDSC SHAR has facilities for testing solid rocket motors, both at ambient conditions and at simulated high altitude conditions. Other test facilities for the environmental testing of rocket motors and their subsystems include Vibration, Shock, Constant Acceleration and Thermal / Humidity.SDSC SHAR has the necessary infrastructure for launching satellites into low earth orbit, polar orbit and geo-stationary transfer orbit. The launch complexes provide complete support for vehicle assembly, fuelling with both earth storable and cryogenic propellants, checkout and launch operations. Apart from these, it has facilities for launching sounding rockets for studying the Earth's upper atmosphere and for controlled reentry and recovery of ISRO's space capsule reentry missions.Safety plays a major role at SDSC SHAR right from the mission / facility design phase to post launch operations. This paper presents briefly the infrastructure available at SDSC SHAR of ISRO for launching sounding rockets, satellite launch vehicles, controlled reentry missions and the built in safety systems. The range safety methodology followed as a part of the real time mission monitoring is presented. The built in safety systems provided onboard the launch vehicle are automatic shut off the propulsion system based on real time mission performance and a passivation system incorporated in the orbit insertion stage are highlighted.

46 CFR 62.35-50 - Tabulated monitoring and safety control requirements for specific systems.

Code of Federal Regulations, 2012 CFR

2012-10-01

... lubrication Pressure Low Main propulsion, controllable pitch propeller Hydraulic oil Pressure High, Low... ......ditto (3) Trial for ignition Status Failure ......ditto Control power Available (pressure) Failure (low... Activated Starting power Pressure (voltage) Low Limit (2) Location in control Status Override (6) Shaft...

46 CFR 62.35-50 - Tabulated monitoring and safety control requirements for specific systems.

Code of Federal Regulations, 2013 CFR

2013-10-01

... lubrication Pressure Low Main propulsion, controllable pitch propeller Hydraulic oil Pressure High, Low... ......ditto (3) Trial for ignition Status Failure ......ditto Control power Available (pressure) Failure (low... Activated Starting power Pressure (voltage) Low Limit (2) Location in control Status Override (6) Shaft...

46 CFR 62.35-50 - Tabulated monitoring and safety control requirements for specific systems.

Code of Federal Regulations, 2014 CFR

2014-10-01

... lubrication Pressure Low Main propulsion, controllable pitch propeller Hydraulic oil Pressure High, Low... ......ditto (3) Trial for ignition Status Failure ......ditto Control power Available (pressure) Failure (low... Activated Starting power Pressure (voltage) Low Limit (2) Location in control Status Override (6) Shaft...

Low-order nonlinear dynamic model of IC engine-variable pitch propeller system for general aviation aircraft

NASA Technical Reports Server (NTRS)

Richard, Jacques C.

1995-01-01

This paper presents a dynamic model of an internal combustion engine coupled to a variable pitch propeller. The low-order, nonlinear time-dependent model is useful for simulating the propulsion system of general aviation single-engine light aircraft. This model is suitable for investigating engine diagnostics and monitoring and for control design and development. Furthermore, the model may be extended to provide a tool for the study of engine emissions, fuel economy, component effects, alternative fuels, alternative engine cycles, flight simulators, sensors, and actuators. Results show that the model provides a reasonable representation of the propulsion system dynamics from zero to 10 Hertz.

Space station needs, attributes, and architectural options study. Volume 2: Program options, architecture, and technology

NASA Technical Reports Server (NTRS)

1983-01-01

Mission scenarios and space station architectures are discussed. Electrical power subsystems (EPS), environmental control and life support, subsystems (ECLSS), and reaction control subsystem (RCS) architectures are addressed. Thermal control subsystems, (TCS), guidance/navigation and control (GN and C), information management systems IMS), communications and tracking (C and T), and propellant transfer and storage systems architectures are discussed.

14 CFR 33.95 - Engine-propeller systems tests.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Engine-propeller systems tests. 33.95... AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.95 Engine-propeller systems tests. If the engine is designed to operate with a propeller, the following tests must be made with a...

14 CFR 33.95 - Engine-propeller systems tests.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Engine-propeller systems tests. 33.95... AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.95 Engine-propeller systems tests. If the engine is designed to operate with a propeller, the following tests must be made with a...

14 CFR 33.95 - Engine-propeller systems tests.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Engine-propeller systems tests. 33.95... AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.95 Engine-propeller systems tests. If the engine is designed to operate with a propeller, the following tests must be made with a...

14 CFR 33.95 - Engine-propeller systems tests.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Engine-propeller systems tests. 33.95... AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.95 Engine-propeller systems tests. If the engine is designed to operate with a propeller, the following tests must be made with a...

14 CFR 33.95 - Engine-propeller systems tests.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Engine-propeller systems tests. 33.95... AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.95 Engine-propeller systems tests. If the engine is designed to operate with a propeller, the following tests must be made with a...

Liquid Propulsion: Propellant Feed System Design. Chapter 2.3.11

NASA Technical Reports Server (NTRS)

Cannon, James L.

2010-01-01

The propellant feed system of a liquid rocket engine determines how the propellants are delivered from the tanks to the thrust chamber. They are generally classified as either pressure fed or pump fed. The pressure-fed system is simple and relies on tank pressures to feed the propellants into the thrust chamber. This type of system is typically used for space propulsion applications and auxiliary propulsion applications requiring low system pressures and small quantities of propellants. In contrast, the pump-fed system is used for high pressure, high performance applications. The selection of one propellant feed system over another is determined based on design trade studies at both the engine and vehicle levels. This chapter first provides a brief overview of the basic configurations of pressure-fed systems. Pump-fed systems are then discussed with greater detail given to the turbomachinery design. Selected design requirements and configurations are provided.

Flight Validation of the Thermal Propellant Gauging Method used at EADS Astrium

NASA Astrophysics Data System (ADS)

Dandaleix, L.; Ounougha, L.; Jallade, S.

2004-10-01

EADS Astrium recently met a major milestone in the field of propellant gauging with the first reorbitation of an Eurostar tanks equipped satellite. It proved successful determining the remaining available propellant mass for spacecraft displacement beyond the customer specified graveyard orbit; thus demonstrating its expertness in Propellant Gauging in correlation with tank residual mass minimization. A critical parameter in satellite operational planning is indeed the accurate knowledge of the on-board remaining propellant mass; basically for the commercial telecommunication missions, where it is the major criterion for lifetime maximization. To provide an accurate and reliable process for measurement of this propellant mass throughout lifetime, EADS Astrium uses a Combination of two independent techniques: The Dead Reckoning Method (maximum accuracy at BOL), based on thrusters flow rate prediction &the Thermal Propellant Gauging Technique, deriving the propellant mass from the tank thermal capacity (Absolute gauging method, with increasing accuracy along lifetime). Then, the present article shows the recent flight validation of the Gauging method obtained for Eurostar E2000 propellant tanks including the validation of the different thermodynamic models. ABBREVIATIONS &ACRONYMS BOL, MOL, EOL: Beginning, Middle &End of Life Cempty: Empty tank thermal inertia [J/K] Chelium: Helium thermal inertia [J/K] Cpropellant: Propellant thermal inertia [J/K] Ct = C1+C2: Total tank thermal inertia (Subscript for upper node and for lower node) [J/K] CPS: Combined Propulsion System DR: Dead Reckoning FM: Flight Model LAE: Liquid Apogee Engine lsb: Least significant byte M0: TPGS Uncertainty component linked to Cempty mox, mfuel: Propellant mass of oxidiser &fuel [kg] Pox, Pfuel: Pressure of oxidiser &fuel [bar] PTA: Propellant Tank Assembly Q: Heater power [W] Qox, Qfuel: Mass flow rate of oxidiser &fuel [kg/s] RCT: Reaction Control Thrusters T0: Spacecraft platform equilibrium temperature TPGS: Thermal Propellant Gauging Software TPGT: Thermal Propellant Gauging Technique T1i: Internal thermal gradients [K] T2i: External thermal gradients [K] Ï 1: Internal thermal characteristic time [s] 2: External thermal characteristic time [s

Outboard Repair.

ERIC Educational Resources Information Center

Hardway, Jack

This consortium-developed instructor's manual for small engine repair (with focus on outboard motors) consists of the following nine instructional units: electrical remote control assembly, mechanical remote control assembly, tilt assemblies, exhaust housing, propeller and trim tabs, cooling system, mechanical gearcase, electrical gearcase, and…

TankSIM: A Cryogenic Tank Performance Prediction Program

NASA Technical Reports Server (NTRS)

Bolshinskiy, L. G.; Hedayat, A.; Hastings, L. J.; Moder, J. P.; Schnell, A. R.; Sutherlin, S. G.

2015-01-01

Accurate prediction of the thermodynamic state of the cryogenic propellants in launch vehicle tanks is necessary for mission planning and successful execution. Cryogenic propellant storage and transfer in space environments requires that tank pressure be controlled. The pressure rise rate is determined by the complex interaction of external heat leak, fluid temperature stratification, and interfacial heat and mass transfer. If the required storage duration of a space mission is longer than the period in which the tank pressure reaches its allowable maximum, an appropriate pressure control method must be applied. Therefore, predictions of the pressurization rate and performance of pressure control techniques in cryogenic tanks are required for development of cryogenic fluid long-duration storage technology and planning of future space exploration missions. This paper describes an analytical tool, Tank System Integrated Model (TankSIM), which can be used for modeling pressure control and predicting the behavior of cryogenic propellant for long-term storage for future space missions. It is written in the FORTRAN 90 language and can be compiled with any Visual FORTRAN compiler. A thermodynamic vent system (TVS) is used to achieve tank pressure control. Utilizing TankSIM, the following processes can be modeled: tank self-pressurization, boiloff, ullage venting, and mixing. Details of the TankSIM program and comparisons of its predictions with test data for liquid hydrogen and liquid methane will be presented in the final paper.

Reduced Toxicity Fuel Satellite Propulsion System Including Fuel Cell Reformer with Alcohols Such as Methanol

NASA Technical Reports Server (NTRS)

Schneider, Steven J. (Inventor)

2001-01-01

A reduced toxicity fuel satellite propulsion system including a reduced toxicity propellant supply for consumption in an axial class thruster and an ACS class thruster. The system includes suitable valves and conduits for supplying the reduced toxicity propellant to the ACS decomposing element of an ACS thruster. The ACS decomposing element is operative to decompose the reduced toxicity propellant into hot propulsive gases. In addition the system includes suitable valves and conduits for supplying the reduced toxicity propellant to an axial decomposing element of the axial thruster. The axial decomposing element is operative to decompose the reduced toxicity propellant into hot gases. The system further includes suitable valves and conduits for supplying a second propellant to a combustion chamber of the axial thruster, whereby the hot gases and the second propellant auto-ignite and begin the combustion process for producing thrust.

Fluidic emergency roll control system. [for emergency aircraft control following failure of primary roll control system

NASA Technical Reports Server (NTRS)

Haefner, K. B.; Honda, T. S.

1973-01-01

A fluidic emergency roll control system for aircraft stabilization in the event of primary flight control failure was evaluated. The fluidic roll control units were designed to provide roll torque proportional to an electrical command as operated by two diametrically opposed thrust nozzles located in the wing tips. The control package consists of a solid propellant gas generator, two diametrically opposed vortex valve modulated thrust nozzles, and an electromagnetic torque motor. The procedures for the design, development, and performance testing of the system are described.

Air Force Science and Technology Plan

DTIC Science & Technology

2011-01-01

charged particles and guide high- power microwaves and radiofrequency waves in the air • Bioenergy – developing renewable biosolar hydrogen...Aeronautical sciences, control sciences, structures and integration Directed Energy High- power microwaves , lasers, beam control, space situational...Propulsion Turbine and rocket engines, advanced propulsion systems , system -level thermal management, and propulsion fuels and propellants Sensors Air

Solar Torque Management for the Near Earth Asteroid Scout CubeSat Using Center of Mass Position Control

NASA Technical Reports Server (NTRS)

Orphee, Juan; Heaton, Andrew; Diedrich, Ben; Stiltner, Brandon C.

2018-01-01

A novel mechanism, the Active Mass Translator (AMT), has been developed for the NASA Near Earth Asteroid (NEA) Scout mission to autonomously manage the spacecraft momentum. The NEA Scout CubeSat will launch as a secondary payload onboard Exploration Mission 1 of the Space Launch System. To accomplish its mission, the CubeSat will be propelled by an 86 square-meter solar sail during its two-year journey to reach asteroid 1991VG. NEA Scout's primary attitude control system uses reaction wheels for holding attitude and performing slew maneuvers, while a cold gas reaction control system performs the initial detumble and early trajectory correction maneuvers. The AMT control system requirements, feedback architecture, and control performance will be presented. The AMT reduces the amount of reaction control propellant needed for momentum management and allows for smaller capacity reaction wheels suitable for the limited 6U spacecraft volume. The reduced spacecraft mass allows higher in-space solar sail acceleration, thus reducing time-of-flight. The reduced time-of-flight opens the range of possible missions, which is limited by the lifetime of typical non-radiation tolerant CubeSat avionics exposed to the deep-space environment.

Large-Scale Advanced Prop-Fan (LAP) pitch change actuator and control design report

NASA Technical Reports Server (NTRS)

Schwartz, R. A.; Carvalho, P.; Cutler, M. J.

1986-01-01

In recent years, considerable attention has been directed toward improving aircraft fuel consumption. Studies have shown that the high inherent efficiency previously demonstrated by low speed turboprop propulsion systems may now be extended to today's higher speed aircraft if advanced high-speed propeller blades having thin airfoils and aerodynamic sweep are utilized. Hamilton Standard has designed a 9-foot diameter single-rotation Large-Scale Advanced Prop-Fan (LAP) which will be tested on a static test stand, in a high speed wind tunnel and on a research aircraft. The major objective of this testing is to establish the structural integrity of large-scale Prop-Fans of advanced construction in addition to the evaluation of aerodynamic performance and aeroacoustic design. This report describes the operation, design features and actual hardware of the (LAP) Prop-Fan pitch control system. The pitch control system which controls blade angle and propeller speed consists of two separate assemblies. The first is the control unit which provides the hydraulic supply, speed governing and feather function for the system. The second unit is the hydro-mechanical pitch change actuator which directly changes blade angle (pitch) as scheduled by the control.

Utilizing Solar Power Technologies for On-Orbit Propellant Production

NASA Technical Reports Server (NTRS)

Fikes, John C.; Howell, Joe T.; Henley, Mark W.

2006-01-01

The cost of access to space beyond low Earth orbit may be reduced if vehicles can refuel in orbit. The cost of access to low Earth orbit may also be reduced by launching oxygen and hydrogen propellants in the form of water. To achieve this reduction in costs of access to low Earth orbit and beyond, a propellant depot is considered that electrolyzes water in orbit, then condenses and stores cryogenic oxygen and hydrogen. Power requirements for such a depot require Solar Power Satellite technologies. A propellant depot utilizing solar power technologies is discussed in this paper. The depot will be deployed in a 400 km circular equatorial orbit. It receives tanks of water launched into a lower orbit from Earth, converts the water to liquid hydrogen and oxygen, and stores up to 500 metric tons of cryogenic propellants. This requires a power system that is comparable to a large Solar Power Satellite capable of several 100 kW of energy. Power is supplied by a pair of solar arrays mounted perpendicular to the orbital plane, which rotates once per orbit to track the Sun. The majority of the power is used to run the electrolysis system. Thermal control is maintained by body-mounted radiators; these also provide some shielding against orbital debris. The propellant stored in the depot can support transportation from low Earth orbit to geostationary Earth orbit, the Moon, LaGrange points, Mars, etc. Emphasis is placed on the Water-Ice to Cryogen propellant production facility. A very high power system is required for cracking (electrolyzing) the water and condensing and refrigerating the resulting oxygen and hydrogen. For a propellant production rate of 500 metric tons (1,100,000 pounds) per year, an average electrical power supply of 100 s of kW is required. To make the most efficient use of space solar power, electrolysis is performed only during the portion of the orbit that the Depot is in sunlight, so roughly twice this power level is needed for operations in sunlight (slightly over half of the time). This power level mandates large solar arrays, using advanced Space Solar Power technology. A significant amount of the power has to be dissipated as heat, through large radiators. This paper briefly describes the propellant production facility and the requirements for a high power system capability. The Solar Power technologies required for such an endeavor are discussed.

Nozzle erosion characterization and minimization for high-pressure rocket motor applications

NASA Astrophysics Data System (ADS)

Evans, Brian

Understanding of the processes that cause nozzle throat erosion and developing methods for mitigation of erosion rate can allow higher operating pressures for advanced rocket motors. However, erosion of the nozzle throat region, which is a strong function of operating pressure, must be controlled to realize the performance gains of higher operating pressures. The objective of this work was the study the nozzle erosion rates at a broad range of pressures from 7 to 34.5 MPa (1,000 to 5,000 psia) using two different rocket motors. The first is an instrumented solidpropellant motor (ISPM), which uses two baseline solid propellants; one is a non-metallized propellant called Propellant S and the other is a metallized propellant called Propellant M. The second test rig is a non-metallized solid-propellant rocket motor simulator (RMS). The RMS is a gas rocket with the ability to vary the combustion-product species composition by systematically varying the flow rates of gaseous reactants. Several reactant mixtures were utilized in the study to determine the relative importance of different oxidizing species (such as H2O, OH, and CO2). Both test rigs are equipped with a windowed nozzle section for real-time X-ray radiography diagnostics of the instantaneous throat variations for deducing the instantaneous erosion rates. The nozzle test section for both motors can also incorporate a nozzle boundary-layer control system (NBLCS) as a means of nozzle erosion mitigation. The effectiveness of the NBLCS at preventing nozzle throat erosion was demonstrated for both the RMS and the ISPM motors at chamber pressures up to 34 MPa (4930 psia). All tests conducted with the NBLCS showed signs of coning of the propellant surface, leading to increased mass burning rate and resultant chamber pressure. Two correlations were developed for the nozzle erosion rates from solid propellant testing, one for metallized propellant and one for non-metallized propellants. The non-metallized propellant correlation also incorporates the RMS data, accounting for swirling flow of the products in the RMS combustor. These correlations are useful for rocket nozzle designs. The correlation for non-metallized propellant and RMS firings was developed in terms of the effective oxidizer mass fraction and effective Reynolds number. The results calculated from this correlation were compared with measured erosion rate data within +/-15% or 0.05 mm/s (2 mils/s). For metallized propellant, the nozzle erosion rate was found to be relatively independent of the concentration of oxidizing species due to the diffusion-controlled process and the partial surface coverage by the liquid Al/Al2O3 layer. The nozzle erosion rate was also found to be lower than those of non-metallized propellant cases. Agreement between predicted and measured erosion rates was found to be within +/-20% or 0.04 mm/s (2 mils/s).

The Effects of Propellant Slosh Dynamics on the Solar Dynamics Observatory

NASA Technical Reports Server (NTRS)

Mason, Paul; Starin, Scott R.

2011-01-01

The Solar Dynamics Observatory (SDO) mission, which is part of the Living With a Star program, was successfully launched and deployed from its Atlas V launch vehicle on February 11, 2010. SDO is an Explorer-class mission now operating in a geosynchronous orbit (GEO). The basic mission is to observe the Sun for a very high percentage of the 5-year mission (10-year goal) with long stretches of uninterrupted observations and with constant, high-data-rate transmission to a dedicated ground station located in White Sands, New Mexico. A significant portion of SDO's launch mass was propellant, contained in two large tanks. To ensure performance with this level of propellant, a slosh analysis was performed. This paper provides an overview of the SDO slosh analysis, the on-orbit experience, and the lessons learned. SDO is a three-axis controlled, single fault tolerant spacecraft. The attitude sensor complement includes sixteen coarse Sun sensors, a digital Sun sensor, three two-axis inertial reference units, two star trackers, and four guide telescopes. Attitude actuation is performed either using four reaction wheels or eight thrusters, depending on the control mode, along with single main engine which nominally provides velocity-change thrust. The attitude control software has five nominal control modes: three wheel-based modes and two thruster-based modes. A wheel-based Safehold running in the Attitude Control Electronics (ACE) box improves the robustness of the system as a whole. All six modes are designed on the same basic proportional-integral-derivative attitude error structure, with more robust modes setting their integral gains to zero. To achieve and maintain a geosynchronous orbit for a 2974-kilogram spacecraft in a cost effective manner, the SDO team designed a high-efficiency propulsive system. This bi-propellant design includes a 100-pound-force main engine and eight 5-pound-force attitude control thrusters. The main engine provides high specific impulse for the maneuvers to attain GEO, while the smaller Attitude Control System (ACS) thrusters manage the disturbance torques of the larger main engine and provide the capability for much smaller orbit adjustment burns. SDO's large solar profile produces a large solar torque disturbance and momentum buildup. This buildup drives the frequency of momentum unloads via ACS thrusters. SDO requires 1409 kilograms (which is approximately half the launch mass) of propellant to achieve and maintain the GEO orbit while performing the momentum unloads for 10 years.

Advanced Booster Composite Case/Polybenzimidazole Nitrile Butadiene Rubber Insulation Development

NASA Technical Reports Server (NTRS)

Gentz, Steve; Taylor, Robert; Nettles, Mindy

2015-01-01

The NASA Engineering and Safety Center (NESC) was requested to examine processing sensitivities (e.g., cure temperature control/variance, debonds, density variations) of polybenzimidazole nitrile butadiene rubber (PBI-NBR) insulation, case fiber, and resin systems and to evaluate nondestructive evaluation (NDE) and damage tolerance methods/models required to support human-rated composite motor cases. The proposed use of composite motor cases in Blocks IA and II was expected to increase performance capability through optimizing operating pressure and increasing propellant mass fraction. This assessment was to support the evaluation of risk reduction for large booster component development/fabrication, NDE of low mass-to-strength ratio material structures, and solid booster propellant formulation as requested in the Space Launch System NASA Research Announcement for Advanced Booster Engineering Demonstration and/or Risk Reduction. Composite case materials and high-energy propellants represent an enabling capability in the Agency's ability to provide affordable, high-performing advanced booster concepts. The NESC team was requested to provide an assessment of co- and multiple-cure processing of composite case and PBI-NBR insulation materials and evaluation of high-energy propellant formulations.

Simulation in production of open rotor propellers: from optimal surface geometry to automated control of mechanical treatment

NASA Astrophysics Data System (ADS)

Grinyok, A.; Boychuk, I.; Perelygin, D.; Dantsevich, I.

2018-03-01

A complex method of the simulation and production design of open rotor propellers was studied. An end-to-end diagram was proposed for the evaluating, designing and experimental testing the optimal geometry of the propeller surface, for the machine control path generation as well as for simulating the cutting zone force condition and its relationship with the treatment accuracy which was defined by the propeller elastic deformation. The simulation data provided the realization of the combined automated path control of the cutting tool.

Annoyance to Noise Produced by a Distributed Electric Propulsion High-Lift System

NASA Technical Reports Server (NTRS)

Rizzi, Stephen A.; Palumbo, Daniel L.; Rathsam, Jonathan; Christian, Andrew; Rafaelof, Menachem

2017-01-01

A psychoacoustic test was performed using simulated sounds from a distributed electric propulsion aircraft concept to help understand factors associated with human annoyance. A design space spanning the number of high-lift leading edge propellers and their relative operating speeds, inclusive of time varying effects associated with motor controller error and atmospheric turbulence, was considered. It was found that the mean annoyance response varies in a statistically significant manner with the number of propellers and with the inclusion of time varying effects, but does not differ significantly with the relative RPM between propellers. An annoyance model was developed, inclusive of confidence intervals, using the noise metrics of loudness, roughness, and tonality as predictors.

Structured Finite Volume Modeling of U.S. Navy Aircraft Engine Test Cells Task 2: Turboprop Engine Vol 1

DTIC Science & Technology

1993-06-01

propeller was modeled using an actuator disk model, where in all of the net results of the propeller (wash, swirl, turbulence) are modeled as source...8217)DUCTI* ~ hVAILJZL TV D0*C) ME.UUT FULLY’ LEGIBLE FREPROI)UCTIM 𔃾S \\JiA~v~r (r7, CL w4 bj4 APPENDIX B COPY AVAUAEL TO e MonC DO* IOT flR=1 VULL...dynamics HVAC systems, energy loss measurement, power ARMY FlAP generation) BDG Shore Faciltes 3B Controls and electrical conservation (electrical systems

Drive-By-Wire Technology

DTIC Science & Technology

2001-05-29

Symposium Intelligent Systems for the Objective Fleet uTransmission controls uSteering (both on-transmission and under-carriage) uBraking (service and...parking) uTransmission select uThrottle uOther Electromechanical Opportunities uTurret drives (elevation, traverse) uAutomatic propellant handling systems

Propulsion and control propellers with thruster nozzles primarily for aircraft applications

NASA Technical Reports Server (NTRS)

Pabst, W.

1986-01-01

A propulsion and control propeller with thruster nozzles, primarily for aircraft application is described. Adjustability of rotor blades at the hub and pressurized gas expulsion combined with an air propeller increase power. Both characteristics are combined in one simple device, and, furthermore, incorporate overall aircraft control so that mechanisms which govern lateral and horizontal movement become superfluous.

Summary of LO2/Ethanol OMS/RCS Technology and Advanced Development 99-2744

NASA Technical Reports Server (NTRS)

Curtis, Leslie A.; Hurlbert, Eric A.

1999-01-01

NASA is pursuing non-toxic propellant technologies applicable to RLV and Space Shuttle orbital maneuvering system (OMS) and reaction control system (RCS). The primary objectives of making advancements in an OMS/RCS system are improved safety, reliability, and reduced operations and maintenance cost, while meeting basic operational and performance requirements. An OMS/RCS has a high degree of direct interaction with the vehicle and crew and requires subsystem and components that are compatible with integration into the vehicle with regard to external mold-line, power, and thermal control. In July 1997, a Phase I effort for the technology and advanced development of an upgrade of the space shuttle was conducted to define the system architecture, propellant tank, feed system, RCS thrusters, and OMS engine. Phase I of the project ran from July 1997 to October 1998. Phase II is currently being planned for the development and test of full-scale prototype of the system in 1999 and 2000. The choice of pressure-fed liquid oxygen (LO2) and ethanol is the result of numerous trade studies conducted from 1980 to 1996. Liquid oxygen and ethanol are clean burning, high-density propellants that provide a high degree of commonality with other spacecraft subsystems including life support, power, and thermal control, and with future human exploration and development of space missions. The key to this pressure-fed system is the use of subcooled liquid oxygen at 350 psia. In this approach, there is 80 degrees R of subcooling, which means that boil-off will not occur until the temperature has risen 80 R. The sub-cooling results naturally from loading propellants at 163 R, which is the saturation temperature at 14.7 psia, and then pressurizing to 350 psia on the launch pad. Thermal insulation and conditioning techniques are then used to limit the LO2 temperature to 185 R maximum, and maintain the sub-cooling. The other key is the wide temperature range of ethanol, -173 F to +300 F, which can provide heat to gasify liquid oxygen or provide a good coolant.

Design of an FPGA-based electronic flow regulator (EFR) for spacecraft propulsion system

NASA Astrophysics Data System (ADS)

Manikandan, J.; Jayaraman, M.; Jayachandran, M.

2011-02-01

This paper describes a scheme for electronically regulating the flow of propellant to the thruster from a high-pressure storage tank used in spacecraft application. Precise flow delivery of propellant to thrusters ensures propulsion system operation at best efficiency by maximizing the propellant and power utilization for the mission. The proposed field programmable gate array (FPGA) based electronic flow regulator (EFR) is used to ensure precise flow of propellant to the thrusters from a high-pressure storage tank used in spacecraft application. This paper presents hardware and software design of electronic flow regulator and implementation of the regulation logic onto an FPGA.Motivation for proposed FPGA-based electronic flow regulation is on the disadvantages of conventional approach of using analog circuits. Digital flow regulation overcomes the analog equivalent as digital circuits are highly flexible, are not much affected due to noise, accurate performance is repeatable, interface is easier to computers, storing facilities are possible and finally failure rate of digital circuits is less. FPGA has certain advantages over ASIC and microprocessor/micro-controller that motivated us to opt for FPGA-based electronic flow regulator. Also the control algorithm being software, it is well modifiable without changing the hardware. This scheme is simple enough to adopt for a wide range of applications, where the flow is to be regulated for efficient operation.The proposed scheme is based on a space-qualified re-configurable field programmable gate arrays (FPGA) and hybrid micro circuit (HMC). A graphical user interface (GUI) based application software is also developed for debugging, monitoring and controlling the electronic flow regulator from PC COM port.

Space Transportation System (STS) propellant scavenging system study. Volume 1: Technical report

NASA Technical Reports Server (NTRS)

1985-01-01

The objectives are to define the most efficient and cost effective methods for scavenging cryogenic and storable propellants and then define the requirements for these scavenging systems. For cryogenic propellants, scavenging is the transfer of propellants from the Shuttle orbiter external tank (ET) and/or main propulsion subsystems (MPS) propellant lines into storage tanks located in the orbiter payload bay for delivery to the user station by a space based transfer stage or the Space Transportation System (STS) by direct insertion. For storable propellants, scavenging is the direct transfer from the orbital maneuvering subsystem (OMS) and/or tankage in the payload bay to users in LEO as well as users in the vicinity of the Space Station.

19 CFR 192.1 - Definitions.

Code of Federal Regulations, 2010 CFR

2010-04-01

... (CONTINUED) EXPORT CONTROL Exportation of Used Self-Propelled Vehicles, Vessels, and Aircraft § 192.1... entered into the commerce of a foreign country. Self-propelled vehicle. “Self-propelled vehicle” includes any automobile, truck, tractor, bus, motorcycle, motor home, self-propelled agricultural machinery...

19 CFR 192.1 - Definitions.

Code of Federal Regulations, 2011 CFR

2011-04-01

... (CONTINUED) EXPORT CONTROL Exportation of Used Self-Propelled Vehicles, Vessels, and Aircraft § 192.1... entered into the commerce of a foreign country. Self-propelled vehicle. “Self-propelled vehicle” includes any automobile, truck, tractor, bus, motorcycle, motor home, self-propelled agricultural machinery...

On spacecraft maneuvers control subject to propellant engine modes.

PubMed

Mazinan, A H

2015-09-01

The paper attempts to address a new control approach to spacecraft maneuvers based upon the modes of propellant engine. A realization of control strategy is now presented in engine on mode (high thrusts as well as further low thrusts), which is related to small angle maneuvers and engine off mode (specified low thrusts), which is also related to large angle maneuvers. There is currently a coarse-fine tuning in engine on mode. It is shown that the process of handling the angular velocities are finalized via rate feedback system in engine modes, where the angular rotations are controlled through quaternion based control (QBCL)strategy in engine off mode and these ones are also controlled through an optimum PID (OPIDH) strategy in engine on mode. Copyright © 2015 ISA. Published by Elsevier Ltd. All rights reserved.

14 CFR 21.3 - Reporting of failures, malfunctions, and defects.

Code of Federal Regulations, 2014 CFR

2014-01-01

... crew compartment or passenger cabin. (4) A malfunction, failure, or defect of a propeller control... structural or flight control system malfunction, defect, or failure which causes an interference with normal control of the aircraft for which derogates the flying qualities. (12) A complete loss of more than one...

14 CFR 21.3 - Reporting of failures, malfunctions, and defects.

Code of Federal Regulations, 2013 CFR

2013-01-01

... crew compartment or passenger cabin. (4) A malfunction, failure, or defect of a propeller control... structural or flight control system malfunction, defect, or failure which causes an interference with normal control of the aircraft for which derogates the flying qualities. (12) A complete loss of more than one...

14 CFR 21.3 - Reporting of failures, malfunctions, and defects.

Code of Federal Regulations, 2012 CFR

2012-01-01

... crew compartment or passenger cabin. (4) A malfunction, failure, or defect of a propeller control... structural or flight control system malfunction, defect, or failure which causes an interference with normal control of the aircraft for which derogates the flying qualities. (12) A complete loss of more than one...

Genetic Algorithm Optimizes Q-LAW Control Parameters

NASA Technical Reports Server (NTRS)

Lee, Seungwon; von Allmen, Paul; Petropoulos, Anastassios; Terrile, Richard

2008-01-01

A document discusses a multi-objective, genetic algorithm designed to optimize Lyapunov feedback control law (Q-law) parameters in order to efficiently find Pareto-optimal solutions for low-thrust trajectories for electronic propulsion systems. These would be propellant-optimal solutions for a given flight time, or flight time optimal solutions for a given propellant requirement. The approximate solutions are used as good initial solutions for high-fidelity optimization tools. When the good initial solutions are used, the high-fidelity optimization tools quickly converge to a locally optimal solution near the initial solution. Q-law control parameters are represented as real-valued genes in the genetic algorithm. The performances of the Q-law control parameters are evaluated in the multi-objective space (flight time vs. propellant mass) and sorted by the non-dominated sorting method that assigns a better fitness value to the solutions that are dominated by a fewer number of other solutions. With the ranking result, the genetic algorithm encourages the solutions with higher fitness values to participate in the reproduction process, improving the solutions in the evolution process. The population of solutions converges to the Pareto front that is permitted within the Q-law control parameter space.

Propellant management for low thrust chemical propulsion systems

NASA Technical Reports Server (NTRS)

Hamlyn, K. M.; Dergance, R. H.; Aydelott, J. C.

1981-01-01

Low-thrust chemical propulsion systems (LTPS) will be required for orbital transfer of large space systems (LSS). The work reported in this paper was conducted to determine the propellant requirements, preferred propellant management technique, and propulsion system sizes for the LTPS. Propellants were liquid oxygen (LO2) combined with liquid hydrogen (LH2), liquid methane or kerosene. Thrust levels of 100, 500, and 1000 lbf were combined with 1, 4, and 8 perigee burns for transfer from low earth orbit to geosynchronous earth orbit. This matrix of systems was evaluated with a multilayer insulation (MLI) or a spray-on-foam insulation. Vehicle sizing results indicate that a toroidal tank configuration is needed for the LO2/LH2 system. Multiple perigee burns and MLI allow far superior LSS payload capability. Propellant settling, combined with a single screen device, was found to be the lightest and least complex propellant management technique.

Low thrust chemical orbit to orbit propulsion system propellant management study

NASA Technical Reports Server (NTRS)

Dergance, R. H.; Hamlyn, K. M.; Tegart, J. R.

1981-01-01

Low thrust chemical propulsion systems were sized for transfer of large space systems from LEO to GEO. The influence of propellant combination, tankage and insulation requirements, and propellant management techniques on the LTPS mass and volume were studied. Liquid oxygen combined with hydrogen, methane or kerosene were the propellant combinations. Thrust levels of 445, 2230, and 4450 N were combined with 1, 4 and 8 perigee burn strategies. This matrix of systems was evaluated using multilayer insulation and spray-on-foam insulation systems. Various combinations of toroidal, cylindrical with ellipsoidal domes, and ellipsoidal tank shapes were investigated. Results indicate that low thrust (445 N) and single perigee burn approaches are considerably less efficient than the higher thrust level and multiple burn strategies. A modified propellant settling approach minimized propellant residuals and decreased system complexity, in addition, the toroid/ellipsoidal tank combination was predicted to be shortest.

75 FR 7934 - Airworthiness Directives; McCauley Propeller Systems 1A103/TCM Series Propellers

Federal Register 2010, 2011, 2012, 2013, 2014

2010-02-23

... with cracks that do not meet acceptable limits, and rework of propellers with cracks that meet..., replacement of propellers with cracks that do not meet acceptable limits, and rework of propellers with cracks... propeller hub, removal from service of propellers with cracks that do not meet acceptable limits, and rework...

14 CFR 35.40 - Functional test.

Code of Federal Regulations, 2012 CFR

2012-01-01

... representative engine on a test stand or on an airplane. The propeller must complete these tests without evidence... STANDARDS: PROPELLERS Tests and Inspections § 35.40 Functional test. The variable-pitch propeller system must be subjected to the applicable functional tests of this section. The same propeller system used in...

14 CFR 35.40 - Functional test.

Code of Federal Regulations, 2011 CFR

2011-01-01

... representative engine on a test stand or on an airplane. The propeller must complete these tests without evidence... STANDARDS: PROPELLERS Tests and Inspections § 35.40 Functional test. The variable-pitch propeller system must be subjected to the applicable functional tests of this section. The same propeller system used in...

46 CFR 154.540 - Quick-closing shut-off valves: Emergency shut-down system.

Code of Federal Regulations, 2011 CFR

2011-10-01

... BULK DANGEROUS CARGOES SAFETY STANDARDS FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design... emergency shut-down system that: (a) Closes all the valves; (b) Is actuated by a single control in at least two locations remote from the quick-closing valves; (c) Is actuated by a single control in each cargo...

46 CFR 154.540 - Quick-closing shut-off valves: Emergency shut-down system.

Code of Federal Regulations, 2010 CFR

2010-10-01

... BULK DANGEROUS CARGOES SAFETY STANDARDS FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design... emergency shut-down system that: (a) Closes all the valves; (b) Is actuated by a single control in at least two locations remote from the quick-closing valves; (c) Is actuated by a single control in each cargo...

46 CFR 154.910 - Inert gas piping: Location.

Code of Federal Regulations, 2011 CFR

2011-10-01

... STANDARDS FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design, Construction and Equipment Atmospheric Control in Cargo Containment Systems § 154.910 Inert gas piping: Location. Inert gas piping must not pass through or terminate in an accommodation, service, or control space. ...

46 CFR 154.702 - Refrigerated carriage.

Code of Federal Regulations, 2013 CFR

2013-10-01

... FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design, Construction and Equipment Cargo Pressure and Temperature Control § 154.702 Refrigerated carriage. (a) Each refrigeration system must: (1... the purpose of this section, a “refrigeration unit” includes a compressor and its motors and controls...

46 CFR 154.702 - Refrigerated carriage.

Code of Federal Regulations, 2012 CFR

2012-10-01

... FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design, Construction and Equipment Cargo Pressure and Temperature Control § 154.702 Refrigerated carriage. (a) Each refrigeration system must: (1... the purpose of this section, a “refrigeration unit” includes a compressor and its motors and controls...

46 CFR 154.702 - Refrigerated carriage.

Code of Federal Regulations, 2011 CFR

2011-10-01

... FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design, Construction and Equipment Cargo Pressure and Temperature Control § 154.702 Refrigerated carriage. (a) Each refrigeration system must: (1... the purpose of this section, a “refrigeration unit” includes a compressor and its motors and controls...

46 CFR 154.702 - Refrigerated carriage.

Code of Federal Regulations, 2010 CFR

2010-10-01

... FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design, Construction and Equipment Cargo Pressure and Temperature Control § 154.702 Refrigerated carriage. (a) Each refrigeration system must: (1... the purpose of this section, a “refrigeration unit” includes a compressor and its motors and controls...

46 CFR 154.702 - Refrigerated carriage.

Code of Federal Regulations, 2014 CFR

2014-10-01

... FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design, Construction and Equipment Cargo Pressure and Temperature Control § 154.702 Refrigerated carriage. (a) Each refrigeration system must: (1... the purpose of this section, a “refrigeration unit” includes a compressor and its motors and controls...

Acoustic prediction methods for the NASA generalized advanced propeller analysis system (GAPAS)

NASA Technical Reports Server (NTRS)

Padula, S. L.; Block, P. J. W.

1984-01-01

Classical methods of propeller performance analysis are coupled with state-of-the-art Aircraft Noise Prediction Program (ANOPP:) techniques to yield a versatile design tool, the NASA Generalized Advanced Propeller Analysis System (GAPAS) for the novel quiet and efficient propellers. ANOPP is a collection of modular specialized programs. GAPAS as a whole addresses blade geometry and aerodynamics, rotor performance and loading, and subsonic propeller noise.

Testing Fundamental Properties of Ionic Liquids for Colloid Microthruster Applications

NASA Technical Reports Server (NTRS)

Anderson, John R.; Plett, Gary; Anderson, Mark; Ziemer, John

2006-01-01

NASA's New Millennium Program is scheduled to test a Disturbance Reduction System (DRS) on Space Technology 7 (ST7) as part of the European Space Agency's (ESA's) LISA Pathfinder Mission in late 2009. Colloid Micronewton Thrusters (CMNTs) will be used to counteract forces, mainly solar photon pressure, that could disturb gravitational reference sensors as part of the DRS. The micronewton thrusters use an ionic liquid, a room temperature molten salt, as propellant. The ionic liquid has a number of unusual properties that have a direct impact on thruster design. One of the most important issues is bubble formation before and during operation, especially during rapid pressure transitions from atmospheric to vacuum conditions. Bubbles have been observed in the feed system causing variations in propellant flow rate that can adversely affect thruster control. Bubbles in the feed system can also increase the likelihood that propellant will spray onto surfaces that can eventually lead to shorting high voltage electrodes. Two approaches, reducing the probability of bubble formation and removing bubbles with a new bubble eliminator device in the flow system, were investigated at Busek Co., Inc. and the Jet Propulsion Laboratory (JPL) to determine the effectiveness of both approaches. Results show that bubble formation is mainly caused by operation at low pressure and volatile contaminants in the propellant coming out of solution. A specification for the maximum tolerable level of contamination has been developed, and procedures for providing system cleanliness have been tested and implemented. The bubble eliminator device has also been tested successfully and has been implemented in recent thruster designs at Busek. This paper focuses on the propellant testing work at JPL, including testing of a breadboard level bubble eliminator device.

Microfabricated Liquid Rocket Motors

NASA Technical Reports Server (NTRS)

Epstein, Alan H.; Joppin, C.; Kerrebrock, J. L.; Schneider, Steven J. (Technical Monitor)

2003-01-01

Under NASA Glenn Research Center sponsorship, MIT has developed the concept of micromachined, bipropellant, liquid rocket engines. This is potentially a breakthrough technology changing the cost-performance tradeoffs for small propulsion systems, enabling new applications, and redefining the meaning of the term low-cost-access-to-space. With this NASA support, a liquid-cooled, gaseous propellant version of the thrust chamber and nozzle was designed, built, and tested as a first step. DARPA is currently funding MIT to demonstrate turbopumps and controls. The work performed herein was the second year of a proposed three-year effort to develop the technology and demonstrate very high power density, regeneratively cooled, liquid bipropellant rocket engine thrust chamber and nozzles. When combined with the DARPA turbopumps and controls, this work would enable the design and demonstration of a complete rocket propulsion system. The original MIT-NASA concept used liquid oxygen-ethanol propellants. The military applications important to DARPA imply that storable liquid propellants are needed. Thus, MIT examined various storable propellant combinations including N2O4 and hydrazine, and H2O2 and various hydrocarbons. The latter are preferred since they do not have the toxicity of N2O4 and hydrazine. In reflection of the newfound interest in H2O2, it is once again in production and available commercially. A critical issue for the microrocket engine concept is cooling of the walls in a regenerative design. This is even more important at microscale than for large engines due to cube-square scaling considerations. Furthermore, the coolant behavior of rocket propellants has not been characterized at microscale. Therefore, MIT designed and constructed an apparatus expressly for this purpose. The report details measurements of two candidate microrocket fuels, JP-7 and JP-10.

An Application Of High-Speed Photography To The Real Ignition Course Of Composite Propellants

NASA Astrophysics Data System (ADS)

Fusheng, Zhang; Gongshan, Cheng; Yong, Zhang; Fengchun, Li; Fanpei, Lei

1989-06-01

That the actual solid rocket motor behavior and delay time of the ignition of Ap/HTPB composite propellant ignited by high energy pyrotechics contained condensed particles have been investigated is the key of this paper. In experiments, using high speed camera, the pressure transducer, the photodiode and synchro circuit control system designed by us synchronistically observe and record all course and details of the ignition. And pressure signal, photodiode signal and high speed photography frame are corresponded one by one.

High Accuracy Liquid Propellant Slosh Predictions Using an Integrated CFD and Controls Analysis Interface

NASA Technical Reports Server (NTRS)

Marsell, Brandon; Griffin, David; Schallhorn, Dr. Paul; Roth, Jacob

2012-01-01

Coupling computational fluid dynamics (CFD) with a controls analysis tool elegantly allows for high accuracy predictions of the interaction between sloshing liquid propellants and th e control system of a launch vehicle. Instead of relying on mechanical analogs which are not valid during aU stages of flight, this method allows for a direct link between the vehicle dynamic environments calculated by the solver in the controls analysis tool to the fluid flow equations solved by the CFD code. This paper describes such a coupling methodology, presents the results of a series of test cases, and compares said results against equivalent results from extensively validated tools. The coupling methodology, described herein, has proven to be highly accurate in a variety of different cases.

Integrated CFD and Controls Analysis Interface for High Accuracy Liquid Propellant Slosh Predictions

NASA Technical Reports Server (NTRS)

Marsell, Brandon; Griffin, David; Schallhorn, Paul; Roth, Jacob

2012-01-01

Coupling computational fluid dynamics (CFD) with a controls analysis tool elegantly allows for high accuracy predictions of the interaction between sloshing liquid propellants and the control system of a launch vehicle. Instead of relying on mechanical analogs which are n0t va lid during all stages of flight, this method allows for a direct link between the vehicle dynamic environments calculated by the solver in the controls analysis tool to the fluid now equations solved by the CFD code. This paper describes such a coupling methodology, presents the results of a series of test cases, and compares said results against equivalent results from extensively validated tools. The coupling methodology, described herein, has proven to be highly accurate in a variety of different cases.

Magnetically actuated propulsion at low Reynolds numbers: towards nanoscale control.

PubMed

Fischer, Peer; Ghosh, Ambarish

2011-02-01

Significant progress has been made in the fabrication of micron and sub-micron structures whose motion can be controlled in liquids under ambient conditions. The aim of many of these engineering endeavors is to be able to build and propel an artificial micro-structure that rivals the versatility of biological swimmers of similar size, e.g. motile bacterial cells. Applications for such artificial "micro-bots" are envisioned to range from microrheology to targeted drug delivery and microsurgery, and require full motion-control under ambient conditions. In this Mini-Review we discuss the construction, actuation, and operation of several devices that have recently been reported, especially systems that can be controlled by and propelled with homogenous magnetic fields. We describe the fabrication and associated experimental challenges and discuss potential applications.

Magnetically actuated propulsion at low Reynolds numbers: towards nanoscale control

NASA Astrophysics Data System (ADS)

Fischer, Peer; Ghosh, Ambarish

2011-02-01

Significant progress has been made in the fabrication of micron and sub-micron structures whose motion can be controlled in liquids under ambient conditions. The aim of many of these engineering endeavors is to be able to build and propel an artificial micro-structure that rivals the versatility of biological swimmers of similar size, e.g. motile bacterial cells. Applications for such artificial ``micro-bots'' are envisioned to range from microrheology to targeted drug delivery and microsurgery, and require full motion-control under ambient conditions. In this Mini-Review we discuss the construction, actuation, and operation of several devices that have recently been reported, especially systems that can be controlled by and propelled with homogenous magnetic fields. We describe the fabrication and associated experimental challenges and discuss potential applications.

Numerical simulation of the flow around a steerable propulsion unit

NASA Astrophysics Data System (ADS)

Pacuraru, F.; Lungu, A.; Ungureanu, C.; Marcu, O.

2010-08-01

Azimuth propulsion units have become during the last decade a more and more popular solution for all kinds of vessels. Azimuth thruster system, combining the propulsion and steering units of conventional ships replaces traditional propellers and lengthy drive shafts and rudders ensuring an excellent vessel steering. In many cases the interaction between the propeller and other components of the propulsion system strongly affects the inflow to the propeller and therefore its performance. The correct estimation of this influence is important for propulsion systems which consist of more than one element, such as pods (shaft, gondola and propeller), ducted propellers (duct, struts and propeller) or bow thrusters (ship form, tunnel, gondola and propeller). The paper proposes a numerical investigation based on RANS computation for solving the viscous flow around an azimuth thruster system to provide a detailed insight into the critical flow regions for determining the optimum inclination angle for struts, for studying the hydrodynamic interactions between various components of the system, for predicting the hydrodynamic performance of the propulsion system and to investigate regions with possible flow separations.

Modeling and simulation of combustion chamber and propellant dynamics and issues in active control of combustion instabilities

NASA Astrophysics Data System (ADS)

Isella, Giorgio Carlo

A method for a comprehensive approach to analysis of the dynamics of an actively controlled combustion chamber, with detailed analysis of the combustion models for the case of a solid rocket propellant, is presented here. The objective is to model the system as interconnected blocks describing the dynamics of the chamber, combustion and control. The analytical framework for the analysis of the dynamics of a combustion chamber is based on spatial averaging, as introduced by Culick. Combustion dynamics are analyzed for the case of a solid propellant. Quasi-steady theory is extended to include the dynamics of the gas-phase and also of a surface layer. The models are constructed so that they produce a combustion response function for the solid propellant that can be immediately introduced in the our analytical framework. The principal objective mechanisms responsible for the large sensitivity, observed experimentally, of propellant response to small variations. We show that velocity coupling, and not pressure coupling, has the potential to be the mechanism responsible for that high sensitivity. We also discuss the effect of particulate modeling on the global dynamics of the chamber and revisit the interpretation of the intrinsic stability limit for burning of solid propellants. Active control is also considered. Particular attention is devoted to the effect of time delay (between sensing and actuation); several methods to compensate for it are discussed, with numerical examples based on the approximate analysis produced by our framework. Experimental results are presented for the case of a Dump Combustor. The combustor exhibits an unstable burning mode, defined through the measurement of the pressure trace and shadowgraph imaging. The transition between stable and unstable modes of operation is characterized by the presence of hysteresis, also observed in other experimental works, and hence not a special characteristic of this combustor. Control is introduced in the form of pulsed secondary fuel. We show the capability of forcing the transition from unstable to stable burning, hence extending the stable operating regime of the combustor. The transition, characterized by the use of a shadowgraph movie sequence, is attributed to a combined fluid-mechanic and combustion mechanism.

Selection of a surface tension propellant management system for the Viking 75 Orbiter.

NASA Technical Reports Server (NTRS)

Dowdy, M. W.; Debrock, S. C.

1972-01-01

Discussion of the propellant management system requirements derived for the Viking 75 mission, and review of a series of surface tension propellant management system design concepts. The chosen concept is identified and its mission operation described. The ullage bubble and bulk liquid positioning characteristics are presented, along with propellant dynamic considerations entailed by thrust initiation/termination. Pressurization design considerations, required to assure minimum disturbance to the bulk propellant, are introduced as well as those of the tank ullage vent. Design provisions to assure liquid communication between tank ends are discussed. Results of a preliminary design study are presented, including mechanical testing requirements to assure structural integrity, propellant compatibility, and proper installation.

STS propellant scavenging systems study. Part 2, volume 1: Executive summary and study results

NASA Technical Reports Server (NTRS)

Williams, Frank L.

1987-01-01

The major objective of the STS Propellant Scavenging Study is to define the hardware, operations, and life cycle costs for recovery of unused Space Transportation System propellants. Earlier phases were concerned exclusively with the recovery of cryogenic propellants from the main propulsion system of the manned STS. The phase of the study covered by this report (Part II Extension) modified the objectives to include cryogenic propellants delivered to orbit by the unmanned cargo vehicle. The Part II Extension had the following objectives: (1) predict OTV propellant requirements from 1995 to 2010; investigate scavenging/transport tank reuse; determine optimum tank sizing and arrangement; and develop hardware concepts for tanks.

Propellant Feed Subsystem for the X-34 Main Propulsion System

NASA Technical Reports Server (NTRS)

McDonald, J. P.; Minor, R. B.; Knight, K. C.; Champion, R. H., Jr.; Russell, F. J., Jr.

1998-01-01

The Orbital Sciences Corporation X-34 vehicle demonstrates technologies and operations key to future reusable launch vehicles. The general flight performance goal of this unmanned rocket plane is Mach 8 flight at an altitude of 250,000 feet. The Main Propulsion System supplies liquid propellants to the main engine, which provides the primary thrust for attaining mission goals. Major NMS design and operational goals are aircraft-like ground operations, quick turnaround between missions, and low initial/operational costs. This paper reviews major design and analysis aspects of the X-34 propellant feed subsystem of the X-34 Main Propulsion System. Topics include system requirements, system design, the integration of flight and feed system performance, propellant acquisition at engine start, and propellant tank terminal drain.

Sources, control, and effects of noise from aircraft propellers and rotors

NASA Technical Reports Server (NTRS)

Mixson, J. S.; Greene, G. C.; Dempsey, T. K.

1981-01-01

Recent NASA and NASA sponsored research on the prediction and control of propeller and rotor source noise, on the analysis and design of fuselage sidewall noise control treatments, and on the measurement and quantification of the response of passengers to aircraft noise is described. Source noise predictions are compared with measurements for conventional low speed propellers, for new high speed propellers (propfans), and for a helicopter. Results from a light aircraft demonstration program are considered which indicates that about 5 dB reduction of flyover noise can be obtained without significant performance penalty. Sidewall design studies are examined for interior noise control in light general aviation aircraft and in large transports using propfan propulsion. The weight of the added acoustic treatment is estimated and tradeoffs between weight and noise reduction are discussed. A laboratory study of passenger response to combined broadband and tonal propeller-like noise is described. Subject discomfort ratings of combined tone broadband noises are compared with ratings of broadband (boundary layer) noise alone and the relative importance of the propeller tones is examined.

Fluid management in the optimization of space construction

NASA Technical Reports Server (NTRS)

Snyder, Howard

1990-01-01

Fluid management impacts strongly on the optimization of space construction. Large quantities of liquids are needed for propellants and life support. The mass of propellant liquids is comparable to that required for the structures. There may be a strong dynamic interaction between the stored liquids and the space structure unless the design minimizes the interaction. The constraints of cost and time required optimization of the supply/resupply strategy. The proper selection and design of the fluid management methods for: slosh control; stratification control; acquisition; transfer; gauging; venting; dumping; contamination control; selection of tank configuration and size; the storage state and the control system can improve the entire system performance substantially. Our effort consists of building mathematical/computer models of the various fluid management methods and testing them against the available experimental data. The results of the models are used as inputs to the system operations studies. During the past year, the emphasis has been on modeling: the transfer of cryogens; sloshing and the storage configuration. The work has been intermeshed with ongoing NASA design and development studies to leverage the funds provided by the Center.

14 CFR 91.527 - Operating in icing conditions.

Code of Federal Regulations, 2014 CFR

2014-01-01

..., windshield, stabilizing or control surface; to a powerplant installation; or to an airspeed, altimeter, rate of climb, or flight attitude instrument system or wing, except that takeoffs may be made with frost... each rotor blade, propeller, windshield, wing, stabilizing or control surface, and each airspeed...

14 CFR 91.527 - Operating in icing conditions.

Code of Federal Regulations, 2011 CFR

2011-01-01

..., windshield, stabilizing or control surface; to a powerplant installation; or to an airspeed, altimeter, rate of climb, or flight attitude instrument system or wing, except that takeoffs may be made with frost... each rotor blade, propeller, windshield, wing, stabilizing or control surface, and each airspeed...

14 CFR 91.527 - Operating in icing conditions.

Code of Federal Regulations, 2013 CFR

2013-01-01

..., windshield, stabilizing or control surface; to a powerplant installation; or to an airspeed, altimeter, rate of climb, or flight attitude instrument system or wing, except that takeoffs may be made with frost... each rotor blade, propeller, windshield, wing, stabilizing or control surface, and each airspeed...

30 CFR 57.4260 - Underground self-propelled equipment.

Code of Federal Regulations, 2011 CFR

2011-07-01

... Prevention and Control Firefighting Equipment § 57.4260 Underground self-propelled equipment. (a) Whenever self-propelled equipment is used underground, a fire extinguisher shall be on the equipment. This... 30 Mineral Resources 1 2011-07-01 2011-07-01 false Underground self-propelled equipment. 57.4260...

Pressure control and analysis report: Hydrogen Thermal Test Article (HTTA)

NASA Technical Reports Server (NTRS)

1971-01-01

Tasks accomplished during the HTTA Program study period included: (1) performance of a literature review to provide system guidelines; (2) development of analytical procedures needed to predict system performance; (3) design and analysis of the HTTA pressurization system considering (a) future utilization of results in the design of a spacecraft maneuvering system propellant package, (b) ease of control and operation, (c) system safety, and (d) hardware cost; and (4) making conclusions and recommendations for systems design.

Random sphere packing model of heterogeneous propellants

NASA Astrophysics Data System (ADS)

Kochevets, Sergei Victorovich

It is well recognized that combustion of heterogeneous propellants is strongly dependent on the propellant morphology. Recent developments in computing systems make it possible to start three-dimensional modeling of heterogeneous propellant combustion. A key component of such large scale computations is a realistic model of industrial propellants which retains the true morphology---a goal never achieved before. The research presented develops the Random Sphere Packing Model of heterogeneous propellants and generates numerical samples of actual industrial propellants. This is done by developing a sphere packing algorithm which randomly packs a large number of spheres with a polydisperse size distribution within a rectangular domain. First, the packing code is developed, optimized for performance, and parallelized using the OpenMP shared memory architecture. Second, the morphology and packing fraction of two simple cases of unimodal and bimodal packs are investigated computationally and analytically. It is shown that both the Loose Random Packing and Dense Random Packing limits are not well defined and the growth rate of the spheres is identified as the key parameter controlling the efficiency of the packing. For a properly chosen growth rate, computational results are found to be in excellent agreement with experimental data. Third, two strategies are developed to define numerical samples of polydisperse heterogeneous propellants: the Deterministic Strategy and the Random Selection Strategy. Using these strategies, numerical samples of industrial propellants are generated. The packing fraction is investigated and it is shown that the experimental values of the packing fraction can be achieved computationally. It is strongly believed that this Random Sphere Packing Model of propellants is a major step forward in the realistic computational modeling of heterogeneous propellant of combustion. In addition, a method of analysis of the morphology of heterogeneous propellants is developed which uses the concept of multi-point correlation functions. A set of intrinsic length scales of local density fluctuations in random heterogeneous propellants is identified by performing a Monte-Carlo study of the correlation functions. This method of analysis shows great promise for understanding the origins of the combustion instability of heterogeneous propellants, and is believed to become a valuable tool for the development of safe and reliable rocket engines.

Propellant and Purge System Contamination "2007: A Summer of Fun"

NASA Technical Reports Server (NTRS)

Galloway, Randy

2010-01-01

This slide presentation reviews the propellant and purge system contamination that occurred during the summer of 2007 at Stennis Space Center. During this period Multiple propellant/pressurant system contamination events prompted a thorough investigation, the results of which are reviewed.

Small Body Hopper Mobility Concepts

NASA Technical Reports Server (NTRS)

Howe, A. Scott; Gernhardt, Michael L.; Lee, Dave E.; Crues, E. Zack; Dexter, Dan E.; Abercromby, Andrew F. J.; Chappell, Steve P.; Nguyen, Hung T.

2015-01-01

A propellant-saving hopper mobility system was studied that could help facilitate the exploration of small bodies such as Phobos for long-duration human missions. The NASA Evolvable Mars Campaign (EMC) has proposed a mission to the moons of Mars as a transitional step for eventual Mars surface exploration. While a Mars transit habitat would be parked in High-Mars Orbit (HMO), crew members would visit the surface of Phobos multiple times for up to 14 days duration (up to 50 days at a time with logistics support). This paper describes a small body surface mobility concept that is capable of transporting a small, two-person Pressurized Exploration Vehicle (PEV) cabin to various sites of interest in the low-gravity environment. Using stored kinetic energy between bounces, a propellant-saving hopper mobility system can release the energy to vector the vehicle away from the surface in a specified direction. Alternatively, the stored energy can be retained for later use while the vehicle is stationary in respect to the surface. The hopper actuation was modeled using a variety of launch velocities, and the hopper mobility was evaluated using NASA Exploration Systems Simulations (NExSyS) for transit between surface sites of interest. A hopper system with linear electromagnetic motors and mechanical spring actuators coupled with Control Moment Gyroscope (CMG) for attitude control will use renewable electrical power, resulting in a significant propellant savings.

Channel-wing System for Thrust Deflection and Force/Moment Generation

NASA Technical Reports Server (NTRS)

Englar, Robert J. (Inventor); Bushnell, Dennis M. (Inventor)

2006-01-01

An aircraft comprising a Channel Wing having blown c h - ne1 circulation control wings (CCW) for various functions. The blown channel CCW includes a channel that has a rounded or near-round trailing edge. The channel further has a trailing-edge slot that is adjacent to the rounded trailing edge of the channel. The trailing-edge slot has an inlet connected to a source of pressurized air and is capable of tangentially discharging pressurized air over the rounded trailing edge. The aircraft further has a propeller that is located in the channel and ahead of the rounded trailing edge of the channel. The propeller provides a propeller thrust exhaust stream across the channel wing to propel the aircraft through the air and to provide high lift. The pressurized air being discharged over the rounded trailing edge provides a high lift that is obtained independent of an aircraft angle of attack, thus preventing the asymmetry. separated flow, and stall experienced by the CC wing at the high angle of attack it required for high lift generation. The aircraft can further include blown outboard circulation control wings (CCW) that are synergistically connected to the blown channel CCWs. The blown outboard CCWs provide additional high lift, control thrust/drag interchange, and can provide all three aerodynamic moments when differential blowing is applied front-to-rear or left-to-right. Both the blown channel CCW and the outboard CCW also have leading-edge blowing slots to prevent flow separation or to provide aerodynamic moments for control.

Models of an In-Situ Propellant Production Plant for Mars Exploration

NASA Technical Reports Server (NTRS)

Goodrich, Charlie; Kurien, James; Millar, Bill; Sweet, Adam; Waterman, Sue; Clancy, Daniel (Technical Monitor)

2001-01-01

An in-situ propellant production system (ISPP) is designed to make rocket fuel from chemicals in the Martian atmosphere in order to reduce the amount of materials that would need to be brought from Earth to support Mars missions. We have developed a description of a hypothetical ISPP system that we would like to make available to researchers who are interested in the problem of automatically diagnosing failures in complex NASA systems. This problem description will help researchers to investigate problems of interest to NASA. We would like to make the following material publicly available: (1) a 'common sense' model of an ISPP system; (2) low- and medium-fidelity simulations of the ISPP system written in Microsoft Excel and HCC; and (3) previously published data and diagrams concerning ISPP components. We do not believe there are any export considerations on these materials for the following reasons: (1) These models are not useful for guidance and real time control of vehicles, encrpytion, or any other software purpose categorized under the Export Control Classification Numbers; and (2) The models are very high level and would not by themselves enable real-time control of a real hardware system. The models are at the level of common sense. They capture, for example, that if a heater is turned on an increase in temperature should result(see the attached excerpt). We do not believe there is any commercial value to this material, given the low commercial demand for propellant plants on mars. We have spoken to acting Code IC Division Chief Dan Clancy, and he concurs with our desire to make these materials publicly available via a technical report.

Ion engine propelled Earth-Mars cycler with nuclear thermal propelled transfer vehicle, volume 2

NASA Technical Reports Server (NTRS)

Meyer, Rudolf X.; Baker, Myles; Melko, Joseph

1994-01-01

The goal of this project was to perform a preliminary design of a long term, reusable transportation system between earth and Mars which would be capable of providing both artificial gravity and shelter from solar flare radiation. The heart of this system was assumed to be a Cycler spacecraft propelled by an ion propulsion system. The crew transfer vehicle was designed to be propelled by a nuclear-thermal propulsion system. Several Mars transportation system architectures and their associated space vehicles were designed.

Long-Term Cryogenic Propellant Storage for the Titan Orbiter Polar Surveyor (TOPS) Mission

NASA Technical Reports Server (NTRS)

Mustafi, Shuvo; Francis, John; Li, Xiaoyi; DeLee, Hudson; Purves, Lloyd; Willis, Dewey; Nixon, Conor; Mcguinness, Dan; Riall, Sara; Devine, Matt;

Long-Term Cryogenic Propellant Storage for the TOPS Mission

NASA Technical Reports Server (NTRS)

Mustafi, Shuvo; Francis, John; Li, Xiaoyi; Purves, Lloyd; DeLee, Hudson; Riall, Sara; McGuinness, Dan; Willis, Dewey; Nixon, Conor; Devine Matt;

High-Lift Propeller System Configuration Selection for NASA's SCEPTOR Distributed Electric Propulsion Flight Demonstrator

NASA Technical Reports Server (NTRS)

Patterson, Michael D.; Derlaga, Joseph M.; Borer, Nicholas K.

2016-01-01

Although the primary function of propellers is typically to produce thrust, aircraft equipped with distributed electric propulsion (DEP) may utilize propellers whose main purpose is to act as a form of high-lift device. These \\high-lift propellers" can be placed upstream of wing such that, when the higher-velocity ow in the propellers' slipstreams interacts with the wing, the lift is increased. This technique is a main design feature of a new NASA advanced design project called Scalable Convergent Electric Propulsion Technology Operations Research (SCEPTOR). The goal of the SCEPTOR project is design, build, and y a DEP aircraft to demonstrate that such an aircraft can be much more ecient than conventional designs. This paper provides details into the high-lift propeller system con guration selection for the SCEPTOR ight demonstrator. The methods used in the high-lift propeller system conceptual design and the tradeo s considered in selecting the number of propellers are discussed.

30 CFR 57.4230 - Surface self-propelled equipment.

Code of Federal Regulations, 2011 CFR

2011-07-01

... Prevention and Control Firefighting Equipment § 57.4230 Surface self-propelled equipment. (a)(1) Whenever a fire or its effects could impede escape from self-propelled equipment, a fire extinguisher shall be on... 30 Mineral Resources 1 2011-07-01 2011-07-01 false Surface self-propelled equipment. 57.4230...

Minimization of Roll Firings for Optimal Propellant Maneuvers

NASA Astrophysics Data System (ADS)

Leach, Parker C.

Attitude control of the International Space Station (ISS) is critical for operations, impacting power, communications, and thermal systems. The station uses gyroscopes and thrusters for attitude control, and reorientations are normally assisted by thrusters on docked vehicles. When the docked vehicles are unavailable, the reduction in control authority in the roll axis results in frequent jet firings and massive fuel consumption. To improve this situation, new guidance and control schemes are desired that provide control with fewer roll firings. Optimal control software was utilized to solve for potential candidates that satisfied desired conditions with the goal of minimizing total propellant. An ISS simulation too was then used to test these solutions for feasibility. After several problem reformulations, multiple candidate solutions minimizing or completely eliminating roll firings were found. Flight implementation would not only save massive amounts of fuel and thus money, but also reduce ISS wear and tear, thereby extending its lifetime.

Propeller propulsion system integration: State of technology survey

NASA Technical Reports Server (NTRS)

Miley, S. J.; Vonlavante, E.

1985-01-01

A literature survey was performed to identify and review technical material applicable to the problem area of propeller propulsion system integration. The survey covered only aerodynamic interference aspects of the problem, and was restricted primarily to propeller effects on the airframe. The subject of airframe aerodynamic interference on the propeller was limited to the problem of vibration due to nonuniform inflow. The problem of airframe effects on propeller performance was not included. A total of 1121 references are given. The references are grouped into the subject areas of Aircraft Stability, Propulsive Efficiency, Aerodynamic Interference, Aerodynamic Interference-Propeller Vibration, and Miscellaneous.

Performance of a capillary propellant management device with hydrazine

NASA Technical Reports Server (NTRS)

Tegart, J. R.

1979-01-01

The propellant management device that was successfully used in the Viking Orbiter spacecraft was selected for the main propulsion system of the Teleoperator Retrieval System (TRS). Due to differences in the missions and different propellants, the operation of this sheet metal vane device required reverification for the TRS application. An analytical investigation was performed considering the adverse acceleration environment and the high contract angle of the hydrazine propellant. Drop tower tests demonstrated that the device would provide propellant acquisition while the TRS was docked with Skylab, but its operation would have to be supplemented through propellant settling when free-flying.

The Shuttle Orbital Maneuvering System P-V-T Propellant Quantity Gaging Accuracy and Leak Detection Allowance for Four Instrumentation Conditions

NASA Technical Reports Server (NTRS)

Duhon, D. D.

1975-01-01

The shuttle orbital maneuvering system (OMS) pressure-volume-temperature (P-V-T) propellant gaging module computes the quantity of usable OMS propellant remaining based on the real gas P-V-T relationship for the propellant tank pressurant, helium. The OMS P-V-T propellant quantity gaging error was determined for four sets of instrumentation configurations and accuracies with the propellant tank operating in the normal constant pressure mode and in the blowdown mode. The instrumentation inaccuracy allowance for propellant leak detection was also computed for these same four sets of instrumentation. These gaging errors and leak detection allowances are presented in tables designed to permit a direct comparison of the effectiveness of the four instrumentation sets. The results show the magnitudes of the improvements in propellant quantity gaging accuracies and propellant leak detection allowances which can be achieved by employing more accurate pressure and temperature instrumentation.

A review and update of the NASA aircraft noise prediction program propeller analysis system

NASA Technical Reports Server (NTRS)

Golub, Robert A.; Nguyen, L. Cathy

1989-01-01

The National Aeronautics and Space Administration (NASA) Aircraft Noise Prediction Program (ANOPP) Propeller Analysis System (PAS) is a set of computational modules for predicting the aerodynamics, performance, and noise of propellers. The ANOPP PAS has the capability to predict noise levels for propeller aircraft certification and produce parametric scaling laws for the adjustment of measured data to reference conditions. A technical overview of the prediction techniques incorporated into the system is presented. The prediction system has been applied to predict the noise signature of a variety of propeller configurations including the effects of propeller angle of attack. A summary of these validation studies is discussed with emphasis being placed on the wind tunnel and flight test programs sponsored by the Federal Aviation Administration (FAA) for the Piper Cherokee Lance aircraft. A number of modifications and improvements have been made to the system and both DEC VAX and IBM-PC versions of the system have been added to the original CDC NOS version.

Survey of Constellation-Era LOX/Methane Development Activities and Future Development Needs

NASA Technical Reports Server (NTRS)

Marshall, William M.; Stiegemeier, Benjamin; Greene, Sandra Elam; Hurlbert, Eric A.

2017-01-01

NASA formed the Constellation Program in 2005 to achieve the objectives of maintaining American presence in low-Earth orbit, returning to the moon for purposes of establishing an outpost, and laying the foundation to explore Mars and beyond in the first half of the 21st century. The Exploration Technology Development Program (ETDP) was formulated to address the technology needs to address Constellation architecture decisions. The Propellants and Cryogenic Advanced Development (PCAD) project was tasked with risk mitigation of specific propulsion related technologies to support ETDP. Propulsion systems were identified as critical technologies owing to the high gear-ratio of lunar Mars landers Cryogenic propellants offer performance advantage over storables (NTOMMH) Mass savings translate to greater payload capacity In-situ production of propellant an attractive feature; methane and oxygen identified as possible Martian in-situ propellants New technologies were required to meet more difficult missions High performance LOX/LH2 deep throttle descent engines High performance LOX/LCH4 ascent main and reaction control system (RCS) engines The PCAD project sought to provide those technologies through Reliable ignition pulse RCS Fast start High efficiency engines Stable deep throttling.

Design of a new VTOL UAV by combining cycloidal blades and FanWing propellers

NASA Astrophysics Data System (ADS)

Li, Daizong

Though the propelling principles of Cycloidal Blades and FanWing propellers are totally different, their structures are similar. Therefore, it is possible to develop an aircraft which combines both types of the propulsion modes of Cyclogyro and FanWing aircrafts. For this kind of aircraft, Cycloidal Blades Mode provides capabilities of Vertical Take-Off and Landing, Instantly Alterable Vector Thrusting, and Low Noise. The FanWing Mode provides capabilities of High Efficiency, Energy-Saving, and Cannot-Stall Low-Speed Cruising. Besides, because both of these propellers are observably better than conventional screw propeller in terms of efficiency, so this type of VTOL UAV could fly with Long Endurance. Furthermore, the usage of flying-wing takes advantage of high structure utilization and high aerodynamic efficiency, eliminates the interference of fuselage and tail, and overcomes flying wing's shortcomings of pitching direction instability and difficulty of control. A new magnetic suspension track-type cycloidal propulsion system is also presented in the paper to solve problems of heavy structure, high mechanical resistance, and low reliability in the traditional cycloidal propellers. The further purpose of this design is to trying to make long-endurance VTOL aircraft and Practical Flying Cars possible in reality, and to bring a new era to the aviation industry.

46 CFR 154.908 - Inert gas generator: Location.

Code of Federal Regulations, 2011 CFR

2011-10-01

... STANDARDS FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design, Construction and Equipment Atmospheric Control in Cargo Containment Systems § 154.908 Inert gas generator: Location. (a) Except as..., service, or control space. (b) An inert gas generator that does not use flame burning equipment may be...

46 CFR 154.1340 - Temperature measuring devices.

Code of Federal Regulations, 2014 CFR

2014-10-01

...) Each device must actuate an audible and visual alarm at the cargo control station and a remote group... STANDARDS FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design, Construction and Equipment... cargo control station. (c) Except for independent tanks type C, each cargo containment system for a...

46 CFR 154.1340 - Temperature measuring devices.

Code of Federal Regulations, 2013 CFR

2013-10-01

...) Each device must actuate an audible and visual alarm at the cargo control station and a remote group... STANDARDS FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design, Construction and Equipment... cargo control station. (c) Except for independent tanks type C, each cargo containment system for a...

46 CFR 154.1340 - Temperature measuring devices.

Code of Federal Regulations, 2011 CFR

2011-10-01

...) Each device must actuate an audible and visual alarm at the cargo control station and a remote group... STANDARDS FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design, Construction and Equipment... cargo control station. (c) Except for independent tanks type C, each cargo containment system for a...

46 CFR 154.1340 - Temperature measuring devices.

Code of Federal Regulations, 2010 CFR

2010-10-01

...) Each device must actuate an audible and visual alarm at the cargo control station and a remote group... STANDARDS FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design, Construction and Equipment... cargo control station. (c) Except for independent tanks type C, each cargo containment system for a...

46 CFR 154.1340 - Temperature measuring devices.

Code of Federal Regulations, 2012 CFR

2012-10-01

...) Each device must actuate an audible and visual alarm at the cargo control station and a remote group... STANDARDS FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Design, Construction and Equipment... cargo control station. (c) Except for independent tanks type C, each cargo containment system for a...

Development, Demonstration, and Analysis of an Integrated Iodine Hall Thruster Feed System

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Peeples, Steven R.; Burt, Adam O.; Martin, Adam K.; Martinez, Armando; Seixal, Joao F.; Mauro, Stephanie

2016-01-01

The design of an in-space iodine-vapor-fed Hall effect thruster propellant management system is described. The solid-iodine propellant tank has unique issues associated with the microgravity environment, requiring a solution where the iodine is maintained in intimate thermal contact with the heated tank walls. The flow control valves required alterations from earlier iterations to survive for extended periods of time in the corrosive iodine-vapor environment. Materials have been selected for the entire feed system that can chemically resist the iodine vapor, with the design now featuring Hastelloy or Inconel for almost all the wetted components. An integrated iodine feed system/Hall thruster demonstration unit was fabricated and tested, with all control being handled by an onboard electronics card specifically designed to operate the feed system. Structural analysis shows that the feed system can survive launch loads after the implementation of some minor reinforcement. Flow modeling, while still requiring significant additional validation, is presented to show its potential in capturing the behavior of components in this low-flow, low-pressure system.

In-Space Chemical Propulsion System Model

NASA Technical Reports Server (NTRS)

Byers, David C.; Woodcock, Gordon; Benfield, Michael P. J.

2004-01-01

Multiple, new technologies for chemical systems are becoming available and include high temperature rockets, very light propellant tanks and structures, new bipropellant and monopropellant options, lower mass propellant control components, and zero boil off subsystems. Such technologies offer promise of increasing the performance of in-space chemical propulsion for energetic space missions. A mass model for pressure-fed, Earth and space-storable, advanced chemical propulsion systems (ACPS) was developed in support of the NASA MSFC In-Space Propulsion Program. Data from flight systems and studies defined baseline system architectures and subsystems and analyses were formulated for parametric scaling relationships for all ACPS subsystem. The paper will first provide summary descriptions of the approaches used for the systems and the subsystems and then present selected analyses to illustrate use of the model for missions with characteristics of current interest.

In-Space Chemical Propulsion System Model

NASA Technical Reports Server (NTRS)

Byers, David C.; Woodcock, Gordon; Benfield, M. P. J.

2004-01-01

Multiple, new technologies for chemical systems are becoming available and include high temperature rockets, very light propellant tanks and structures, new bipropellant and monopropellant options, lower mass propellant control components, and zero boil off subsystems. Such technologies offer promise of increasing the performance of in-space chemical propulsion for energetic space missions. A mass model for pressure-fed, Earth and space-storable, advanced chemical propulsion systems (ACPS) was developed in support of the NASA MSFC In-Space Propulsion Program. Data from flight systems and studies defined baseline system architectures and subsystems and analyses were formulated for parametric scaling relationships for all ACPS subsystems. The paper will first provide summary descriptions of the approaches used for the systems and the subsystems and then present selected analyses to illustrate use of the model for missions with characteristics of current interest.

Installation effects on propeller wake/vortex induced structure-borne noise transmission

NASA Technical Reports Server (NTRS)

Unruh, J. F.

1989-01-01

A laboratory-based test apparatus was employed to investigate the effects of power-plant placement, engine/nacelle mass installation, and wing-to-fuselage attachment methods on propeller-induced structure-borne noise (SBN) transmission levels and their effects on noise-control measures. Data are presented showing SBN transmission is insensitive to propeller spanwise placement, however some sensitivity is seen in propeller-to-wing spacing. Installation of an engine/nacelle mass and variation in wing-to-fuselage attachments have measurable influences on SBN transmission and control measures.

Aerodynamic Design of a Propeller for High-Altitude Balloon Trajectory Control

NASA Technical Reports Server (NTRS)

Eppler, Richard; Somers, Dan M.

2012-01-01

The aerodynamic design of a propeller for the trajectory control of a high-altitude, scientific balloon has been performed using theoretical methods developed especially for such applications. The methods are described. Optimum, nonlinear chord and twist distributions have been developed in conjunction with the design of a family of airfoils, the SE403, SE404, and SE405, for the propeller. The very low Reynolds numbers along the propeller blade fall in a range that has yet to be rigorously investigated, either experimentally or theoretically.

Development of control strategies for safe microburst penetration: A progress report

NASA Technical Reports Server (NTRS)

Psiaki, Mark L.

1987-01-01

A single-engine, propeller-driven, general-aviation model was incorporated into the nonlinear simulation and into the linear analysis of root loci and frequency response. Full-scale wind tunnel data provided its aerodynamic model, and the thrust model included the airspeed dependent effects of power and propeller efficiency. Also, the parameters of the Jet Transport model were changed to correspond more closely to the Boeing 727. In order to study their effects on steady-state repsonse to vertical wind inputs, altitude and total specific energy (air-relative and inertial) feedback capabilities were added to the nonlinear and linear models. Multiloop system design goals were defined. Attempts were made to develop controllers which achieved these goals.

Combined high and low-thrust geostationary orbit insertion with radiation constraint

NASA Astrophysics Data System (ADS)

Macdonald, Malcolm; Owens, Steven Robert

2018-01-01

The sequential use of an electric propulsion system is considered in combination with a high-thrust propulsion system for application to the propellant-optimal Geostationary Orbit insertion problem, whilst considering both temporal and radiation flux constraints. Such usage is found to offer a combined propellant mass saving when compared with an equivalent high-thrust only transfer. This propellant mass saving is seen to increase as the allowable transfer duration is increased, and as the thrust from the low-thrust system is increased, assuming constant specific impulse. It was found that the required plane change maneuver is most propellant-efficiently performed by the high-thrust system. The propellant optimal trajectory incurs a significantly increased electron flux when compared to an equivalent high-thrust only transfer. However, the electron flux can be reduced to a similar order of magnitude by increasing the high-thrust propellant consumption, whilst still delivering an improved mass fraction.

The E-3 Test Facility at Stennis Space Center: Research and Development Testing for Cryogenic and Storable Propellant Combustion Systems

NASA Technical Reports Server (NTRS)

Pazos, John T.; Chandler, Craig A.; Raines, Nickey G.

2009-01-01

This paper will provide the reader a broad overview of the current upgraded capabilities of NASA's John C. Stennis Space Center E-3 Test Facility to perform testing for rocket engine combustion systems and components using liquid and gaseous oxygen, gaseous and liquid methane, gaseous hydrogen, hydrocarbon based fuels, hydrogen peroxide, high pressure water and various inert fluids. Details of propellant system capabilities will be highlighted as well as their application to recent test programs and accomplishments. Data acquisition and control, test monitoring, systems engineering and test processes will be discussed as part of the total capability of E-3 to provide affordable alternatives for subscale to full scale testing for many different requirements in the propulsion community.

The starting transient of solid propellant rocket motors with high internal gas velocities. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Peretz, A.; Caveny, L. H.; Kuo, K. K.; Summerfield, M.

1973-01-01

A comprehensive analytical model which considers time and space development of the flow field in solid propellant rocket motors with high volumetric loading density is described. The gas dynamics in the motor chamber is governed by a set of hyperbolic partial differential equations, that are coupled with the ignition and flame spreading events, and with the axial variation of mass addition. The flame spreading rate is calculated by successive heating-to-ignition along the propellant surface. Experimental diagnostic studies have been performed with a rectangular window motor (50 cm grain length, 5 cm burning perimeter and 1 cm hydraulic port diameter), using a controllable head-end gaseous igniter. Tests were conducted with AP composite propellant at port-to-throat area ratios of 2.0, 1.5, 1.2, and 1.06, and head-end pressures from 35 to 70 atm. Calculated pressure transients and flame spreading rates are in very good agreement with those measured in the experimental system.

Implementation Options for the PROPEL Electrodynamic Tether Demonstration Mission

NASA Technical Reports Server (NTRS)

Bilen, Sven G.; Johnson, Les; Gilchrist, Brian E.; Hoyt, Robert P.; Elder, Craig H.; Fuhrhop, Keith P.; Scadera, Michael P.; Stone, Nobie H.

2014-01-01

The PROPEL flight mission concept will demonstrate the safe use of an electrodynamic tether for generating thrust. PROPEL is being designed to be a versatile electrodynamic-tether system for multiple end users and to be flexible with respect to platform. As such, several implementation options are being explored, including a comprehensive mission design for PROPEL with a mission duration of six months; a space demonstration mission concept design with configuration of a pair of tethered satellites, one of which is the Japanese H-II Transfer Vehicle; and an ESPA-based system. We report here on these possible implementation options for PROPEL. electrodynamic tether; PROPEL demonstration mission; propellantless propulsion

Wave Amplitude Dependent Engineering Model of Propellant Slosh in Spherical Tanks

NASA Technical Reports Server (NTRS)

Brodnick, Jacob; Westra, Douglas G.; Eberhart, Chad J.; Yang, Hong Q.; West, Jeffrey S.

2016-01-01

Liquid propellant slosh is often a concern for the controllability of flight vehicles. Anti-slosh devices are traditionally included in propellant tank designs to limit the amount of sloshing allowed during flight. These devices and any necessary supports can be quite heavy to meet various structural requirements. Some of the burden on anti-slosh devices can be relieved by exploiting the nonlinear behavior of slosh waves in bare smooth wall tanks. A nonlinear regime slosh model for bare spherical tanks was developed through a joint analytical and experimental effort by NASA/MSFC. The developed slosh model accounts for the large damping inherent in nonlinear slosh waves which is more accurate and drives conservatism from vehicle stability analyses that use traditional bare tank slosh models. A more accurate slosh model will result in more realistic predicted slosh forces during flight reducing or removing the need for active controls during a maneuver or baffles in the tank design. Lower control gains and smaller or fewer tank baffles can reduce cost and system complexity while increasing vehicle performance. Both Computational Fluid Dynamics (CFD) simulation and slosh testing of three different spherical tank geometries were performed to develop the proposed slosh model. Several important findings were made during this effort in addition to determining the parameters to the nonlinear regime slosh model. The linear regime slosh damping trend for spherical tanks reported in NASA SP-106 was shown to be inaccurate for certain regions of a tank. Additionally, transition to the nonlinear regime for spherical tanks was only found to occur at very large wave amplitudes in the lower hemisphere and was a strong function of the propellant fill level in the upper hemisphere. The nonlinear regime damping trend was also found to be a function of the propellant fill level.

Shuttle cryogenic supply system optimization study. Volume 6: Appendixes

NASA Technical Reports Server (NTRS)

1973-01-01

The optimization of the cryogenic supply system for space shuttles is discussed. The subjects considered are: (1) auxiliary power unit parametric data, (2) propellant acquisition, (3) thermal protection and thermodynamic properties, (4) instrumentation and controls, and (5) initial component redundancy evaluations. Diagrams of the systems are provided. Graphs of the performance capabilities are included.

Potential propellant storage and feed systems for space station resistojet propulsion options

NASA Technical Reports Server (NTRS)

Bader, Clayton H.

1987-01-01

The resistojet system has been defined as part of the baseline propulsion system for the initial Operating Capability Space Station. The resistojet propulsion module will perform a reboost function using a wide variety of fluids as propellants. There are many optional propellants and propellant combinations for use in the resistojet including (but not limited to): hydrazine, hydrogen, oxygen, nitrogen, water, carbon dioxide, and methane. Many different types of propulsion systems have flown or have been conceptualized that may have application for use with resistojets. This paper describes and compares representative examples of these systems that may provide a basis for space station resistojet system design.

Assessment of Cost Impacts of Using Non-Toxic Propulsion in Satellites

NASA Astrophysics Data System (ADS)

Schiebener, P. J.; Gies, O.; Stuhlberger, J.; Schmitz, H.-D.

2002-01-01

The growing costs of space missions, the need for increased mission performance, and concerns associated with environmental issues deeply influence propulsion system design and propellant selection criteria. A propellant's performance was defined in the past exclusively in terms of specific impulse and density, but now high-performance, non-toxic, non-sophisticated mono- propellant systems are key drivers, and are considered for development to replace the traditional hydrazine (N2H4) mono-propellant thrusters. The mono-propellants under consideration are propellant formulations, which should be environmentally friendly, should have a high density, equal or better performance and better thermal characteristics than hydrazine. These considerations raised interest specially in the candidates of Hydroxylammonium Nitrate (HAN)-based propellants, Ammoniumdinitramide (ADN)-based propellants, Tri-ethanol (TEAN)-based propellants, Hydrazinium Nitroformate (HNF)-based propellants, Hydrogen Peroxide (H2O2)-based propellants. A near-term objective in consideration of satellite related process optimisation is to significantly reduce on-ground operations costs and at the same time improve mission performance. A far-term objective is to obtain a system presenting a very high performance, illustrated by a high specific impulse. Moving to a "non-toxic" propulsion system seems to be a solution to these two goals. The sought after benefits for non-toxic spacecraft mono-propellant propulsion are under investigation taking into account the four main parameters which are mandatory for customer satisfaction while meeting the price constraints: - Reliability, availability, maintainability and safety, - Manufacturing, assembly, integration and test, - Launch preparation and support, - Ground support equipment. These benefits of non-toxic mono-propellants can be proven by various examples, like an expected reduction of development costs due the non-toxicity of propellants which might allow "easier" design, reducing some inhibits for ground safety, leading to a shorter development time, and consequently to reduced program costs. Operational costs could be reduced due to the use of non-toxic propellant. Their non-toxicity, in comparison to the traditional propellants, will avoid special safety procedures and also parallelisation of processes during all phases of AIT and launch preparations. The costs directly associated with propellant handling, transport and storage should be lower, also follow-on costs risk is minimised because of the elimination or significant reduction of toxic and carcinogenic characteristics of the propellants. The physical characteristic and properties of some of the propellants formulations mentioned, like a higher density than hydrazine, support the beneficial aspects: a global S/C weight reduction could be achieved due to smaller tanks.

Reusable Reentry Satellite (RRS): Thermal control trade study

NASA Technical Reports Server (NTRS)

Wallace, Clark

1990-01-01

The design and assessment work performed in defining the on-orbit Thermal Control Subsystem (TCS) requirements for the Reusable Reentry Satellite (RRS) is discussed. Specifically, it describes the hardware and design measures necessary for maintaining the Payload Module (PM) Environmental Control Life Support System (ECLSS) heat exchanger, the hydrazine propellant, and PM water supply within their required temperature limits.

Analysis of 100-lb(sub f) (445-N) LO2-LCH4 Reaction Control Engine Impulse Bit Performance

NASA Technical Reports Server (NTRS)

Marshall, William M.; Klenhenz, Julie E.

2012-01-01

Recently, liquid oxygen-liquid methane (LO2-LCH4) has been considered as a potential green propellant alternative for future exploration missions. The Propulsion and Cryogenic Advanced Development (PCAD) project was tasked by NASA to develop this propulsion combination to enable safe and cost-effective exploration missions. To date, limited experience with such combinations exist, and as a result a comprehensive test program is critical to demonstrating with the viability of implementing such a system. The NASA Glenn Research Center conducted a test program of a 100-lbf (445-N) reaction control engine (RCE) at the Center s Altitude Combustion Stand (ACS), focusing on altitude testing over a wide variety of operational conditions. The ACS facility includes unique propellant conditioning feed systems (PCFS), which allow precise control of propellant inlet conditions to the engine. Engine performance as a result of these inlet conditions was examined extensively during the test program. This paper is a companion to the previous specific impulse testing paper, and discusses the pulsed-mode operation portion of testing, with a focus on minimum impulse bit (MIB) and repeatable pulse performance. The engine successfully demonstrated target MIB performance at all conditions, as well as successful demonstration of repeatable pulse widths. Some anomalous conditions experienced during testing are also discussed, including a double pulse phenomenon, which was not noted in previous test programs for this engine.

Summary of Air Force Research Laboratory Support for the NASA Green Propellant Infusion Mission

DTIC Science & Technology

2015-07-01

system to transfer propellant from a bulk propellant tank into a spacecraft tank. It also called for the transfer of propellant from a large transport...launch pressurized propellant tanks on a spacecraft or satellite, a fracture mechanics analysis is required to verify the safe design life of the...a bulk propellant tank into a spacecraft tank. It also called for the transfer of propellant from a large transport container into a specialized

Self-propelled in-tube shuttle and control system for automated measurements of magnetic field alignment

DOE Office of Scientific and Technical Information (OSTI.GOV)

Boroski, W.N.; Nicol, T.H.; Pidcoe, S.V.

1990-03-01

A magnetic field alignment gauge is used to measure the field angle as a function of axial position in each of the magnets for the Superconducting Super Collider (SSC). Present measurements are made by manually pushing the through the magnet bore tube and stopping at intervals to record field measurements. Gauge location is controlled through graduation marks and alignment pins on the push rods. Field measurements are recorded on a logging multimeter with tape output. Described is a computerized control system being developed to replace the manual procedure for field alignment measurements. The automated system employs a pneumatic walking devicemore » to move the measurement gauge through the bore tube. Movement of the device, called the Self-Propelled In-Tube Shuttle (SPITS), is accomplished through an integral, gas driven, double-acting cylinder. The motion of the SPITS is transferred to the bore tube by means of a pair of controlled, retractable support feet. Control of the SPITS is accomplished through an RS-422 interface from an IBM-compatible computer to a series of solenoid-actuated air valves. Direction of SPITS travel is determined by the air-valve sequence, and is managed through the control software. Precise axial position of the gauge within the magnet is returned to the control system through an optically-encoded digital position transducer attached to the shuttle. Discussed is the performance of the transport device and control system during preliminary testing of the first prototype shuttle. 1 ref., 7 figs.« less

The Iodine Satellite (iSAT) Propellant Feed System - Design and Development

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Seixal, Joao F.; Mauro, Stephanie L.; Burt, Adam O.; Martinez, Armando; Martin, Adam K.

2017-01-01

The development, modeling, and testing of components and subsystems required to feed iodine propellant to a 200-W Hall thruster and cathode are described. This work aims to address design deficiencies and issues associated with the propellant feed system that were revealed by an integrated thruster-cathode-feed system test. The feed system design is modified to use materials that are more resistant to the highly-reactive nature of iodine propellant. Dynamic modeling indicates that the inclusion of additional constraints on feed system tubing will reduce the vibrationally-induced stresses that occur during launch. Full spacecraft thermal modeling show that the feed system heater power levels are sufficient to heat the tank and propellant lines to operating temperatures, where iodine in the tank is sublimed to supply propellant for operation and the tubing is elevated in temperature to keep propellant from redepositing to block the flow. Experiments are conducted to demonstrate that is it possible through the application of heating to clear an iodine deposit blocking the flow. Deposits in the low-pressure portion of the system near the exit to vacuum are shown to be relatively easy to remove in this manner while blockages forming upstream nearer to the higher-pressure propellant tank require significantly more effort to remove. Fluid flow modeling of the feed system is performed, exhibiting some qualitative agreement with experimental data. However, the highly viscous nature of the fluid flow and the dependence of the component flow coefficients on the Reynolds number are likely causes of the generally-poor quantitative agreement between the modeling results and experimentally-measured fluid flow properties.

Propeller Study. Part 2: the Design of Propellers for Minimum Noise

NASA Technical Reports Server (NTRS)

Ormsbee, A. I.; Woan, C. J.

1977-01-01

The design of propellers which are efficient and yet produce minimum noise requires accurate determinations of both the flow over the propeller. Topics discussed in relating aerodynamic propeller design and propeller acoustics include the necessary approximations and assumptions involved, the coordinate systems and their transformations, the geometry of the propeller blade, and the problem formulations including the induced velocity, required in the determination of mean lines of blade sections, and the optimization of propeller noise. The numerical formulation for the lifting-line model are given. Some applications and numerical results are included.

46 CFR 151.12-10 - Operation of oceangoing non-self-propelled ships Carrying Category D NLS.

Code of Federal Regulations, 2011 CFR

2011-10-01

... 46 Shipping 5 2011-10-01 2011-10-01 false Operation of oceangoing non-self-propelled ships... and Operating Requirements for Control of Pollution From Category D NLS Cargoes § 151.12-10 Operation of oceangoing non-self-propelled ships Carrying Category D NLS. (a) An oceangoing non-self-propelled...

14 CFR 23.771 - Pilot compartment.

Code of Federal Regulations, 2014 CFR

2014-01-01

... the propellers so that no part of the pilot or the controls lies in the region between the plane of rotation of any inboard propeller and the surface generated by a line passing through the center of the propeller hub making an angle of 5 degrees forward or aft of the plane of rotation of the propeller. [Doc...

14 CFR 23.771 - Pilot compartment.

Code of Federal Regulations, 2012 CFR

2012-01-01

... the propellers so that no part of the pilot or the controls lies in the region between the plane of rotation of any inboard propeller and the surface generated by a line passing through the center of the propeller hub making an angle of 5 degrees forward or aft of the plane of rotation of the propeller. [Doc...

14 CFR 23.771 - Pilot compartment.

Code of Federal Regulations, 2013 CFR

2013-01-01

... the propellers so that no part of the pilot or the controls lies in the region between the plane of rotation of any inboard propeller and the surface generated by a line passing through the center of the propeller hub making an angle of 5 degrees forward or aft of the plane of rotation of the propeller. [Doc...

Development of 4.5 kW SPT T-160 for Joint Tests with NPP TOPAZ-2

NASA Astrophysics Data System (ADS)

Koroteev, Anatoly; Petrosov, Valery A.; Vasin, Anatoly L.; Baranov, Vladimir L.; Rebrov, Serdei G.; Wetch, Joseph R.; Wong, See-Pok

1994-07-01

The electric thruster T-160 is related to a low thrust rocket engines group which can be used on spacecrafts for their orbit correction and parameters change. T-160 flight tests are supposed to be conducted together with the TOPAZ-2 Nuclear Power Plant (NPP) as a part of the NEPSTP program. The thruster T-160 is a Hall-type plasma accelerator with the closed electron drift. Plasma is accelerated in crossed magnetic and electric fields in an accelerating channel made of insulating material. Gaseous xenon is used as a propellant for the thruster, electric power supply of 4-4.5 kW is required for its operation. The cathode-neutralizer emission element is made of porous tungsten saturated with barium salts. The thruster T-160 is equipped with a flow control unit ensuring propellant coming to the cathode and gas distributing anode, and a power processing unit and can be delivered with a propellant storage and supply system, thus forming a complete propulsion system set.

Space Propulsion Research Facility (B-2): An Innovative, Multi-Purpose Test Facility

NASA Technical Reports Server (NTRS)

Hill, Gerald M.; Weaver, Harold F.; Kudlac, Maureen T.; Maloney, Christian T.; Evans, Richard K.

2011-01-01

The Space Propulsion Research Facility, commonly referred to as B-2, is designed to hot fire rocket engines or upper stage launch vehicles with up to 890,000 N force (200,000 lb force), after environmental conditioning of the test article in simulated thermal vacuum space environment. As NASA s third largest thermal vacuum facility, and the largest designed to store and transfer large quantities of propellant, it is uniquely suited to support developmental testing associated with large lightweight structures and Cryogenic Fluid Management (CFM) systems, as well as non-traditional propulsion test programs such as Electric and In-Space propulsion. B-2 has undergone refurbishment of key subsystems to support the NASA s future test needs, including data acquisition and controls, vacuum, and propellant systems. This paper details the modernization efforts at B-2 to support the Nation s thermal vacuum/propellant test capabilities, the unique design considerations implemented for efficient operations and maintenance, and ultimately to reduce test costs.

Space Technology for Medical Aids

NASA Technical Reports Server (NTRS)

1982-01-01

Under one of the earliest contracts awarded in the Apollo lunar landing program, Parker Hannifin Corporation developed and produced equipment for controlling the flow of propellants into the mammoth engines of the Saturn moonbooster. Today, Parker is supplying the huge valves that control propellant flow from the Space Shuttle's external fuel tank to the engines of the Shuttle Orbiter as well as the "peanut valve," named for its small size. In 1977, NASA, recognizing the company's special expertise in miniature systems, asked Parker to participate in the development of an implantable artificial sphincter for control of urinary incontinence. The company's peanut valve experience provided an ideal base for a new biomedical project, the Programmable Implantable Medication System (PIMS) for continuous, computer-directed delivery of precisely metered medication -- insulin, for example -- within a patient's body. The work on PIMS also inspired development of Micromed, a related programmable medication device for external, rather than implantable use. The Biomedical Products Division has also applied its fluid handling expertise to a drugless therapy system called Cryomax for the treatment of such disorders as rheumatoid arthritis and lupus.

Advanced Liquid Feed Experiment

NASA Astrophysics Data System (ADS)

Distefano, E.; Noll, C.

1993-06-01

The Advanced Liquid Feed Experiment (ALFE) is a Hitchhiker experiment flown on board the Shuttle of STS-39 as part of the Space Test Payload-1 (STP-1). The purpose of ALFE is to evaluate new propellant management components and operations under the low gravity flight environment of the Space Shuttle for eventual use in an advanced spacecraft feed system. These components and operations include an electronic pressure regulator, an ultrasonic flowmeter, an ultrasonic point sensor gage, and on-orbit refill of an auxiliary propellant tank. The tests are performed with two transparent tanks with dyed Freon 113, observed by a camera and controlled by ground commands and an on-board computer. Results show that the electronic pressure regulator provides smooth pressure ramp-up, sustained pressure control, and the flexibility to change pressure settings in flight. The ultrasonic flowmeter accurately measures flow and detects gas ingestion. The ultrasonic point sensors function well in space, but not as a gage during sustained low-gravity conditions, as they, like other point gages, are subject to the uncertainties of propellant geometry in a given tank. Propellant transfer operations can be performed with liquid-free ullage equalization at a 20 percent fill level, gas-free liquid transfer from 20-65 percent fill level, minimal slosh, and can be automated.

Noise of a simulated installed model counterrotation propeller at angle-of-attack and takeoff/approach conditions

NASA Technical Reports Server (NTRS)

Woodward, Richard P.

1990-01-01

Acoustic results for two model counterrotation propellers are presented. The propellers were tested over a range of rotational speeds and propeller axis angles of attack in both the baseline configuration and the installed configuration consisting of a simulated upstream nacelle support pylon and fuselage section. Acoustic data were taken with a polar microphone probe attached to the downstream propeller housing, capable of surveying directivities at several azimuthal locations. The forward and aft rotor power coefficients and fundamental rotor-alone tone levels are found to be directly controlled by propeller axis angle of attack. The second-order rotor-alone tones are strongly influenced by the upstream pylon wake at 80 percent speed; however, rotor-alone mechanisms control the tone level at 90 percent speed, while rotor-rotor interaction tones are essentially unaffected by the presence of the simulated installation.

Inflight Characterization of the Cassini Spacecraft Propellant Slosh and Structural Frequencies

NASA Technical Reports Server (NTRS)

Lee, Allan Y.; Stupik, Joan

2015-01-01

While there has been extensive theoretical and analytical research regarding the characterization of spacecraft propellant slosh and structural frequencies, there have been limited studies to compare the analytical predictions with measured flight data. This paper uses flight telemetry from the Cassini spacecraft to get estimates of high-g propellant slosh frequencies and the magnetometer boom frequency characteristics, and compares these values with those predicted by theoretical works. Most Cassini attitude control data are available at a telemetry frequency of 0.5 Hz. Moreover, liquid sloshing is attenuated by propellant management device and attitude controllers. Identification of slosh and structural frequency are made on a best-effort basis. This paper reviews the analytical approaches that were used to predict the Cassini propellant slosh frequencies. The predicted frequencies are then compared with those estimated using telemetry from selected Cassini burns where propellant sloshing was observed (such as the Saturn Orbit Insertion burn).

Electromagnetic Pumps for Conductive-Propellant Feed Systems

NASA Technical Reports Server (NTRS)

Markusic, T. E.; Polzin, K. A.

2005-01-01

There has been a recent, renewed interest in high-power electric thrusters for application in nuclear-electric propulsion systems. Two of the most promising thrusters utilize liquid metal propellants: the lithium-fed magnetoplasmadynamic thruster and the bismuth-fed Hall thruster. An important element of part of the maturation of these thrusters will be the development of compact, reliable conductive-propellant feed system components. In the present paper we provide design considerations and experimental calibration data for electromagnetic (EM) pumps. The role of an electromagnetic pump in a liquid metal feed system is to establish a pressure gradient between the propellant reservoir and the thruster - to establish the requisite mass flow rate. While EM pumps have previously been used to a limited extent in nuclear reactor cooling loops, they have never been implemented in electric propulsion (EP) systems. The potential benefit of using EM pumps for EP are reliability (no moving parts) and the ability to precisely meter the propellant flow rate. We have constructed and tested EM pumps that use gallium, lithium, and bismuth propellants. Design details, test results (pressure developed versus current), and material compatibility issues are reported. It is concluded that EM pumps are a viable technology for application in both laboratory and flight EP conductive-propellant feed systems.

100-lbf LO2/CH4 RCS Thruster Testing and Validation

NASA Technical Reports Server (NTRS)

Barnes, Frank; Cannella, Matthew; Gomez, Carlos; Hand, Jeffrey; Rosenberg, David

2009-01-01

100 pound thrust liquid Oxygen-Methane thruster sized for RCS (Reaction Control System) applications. Innovative Design Characteristics include: a) Simple compact design with minimal part count; b) Gaseous or Liquid propellant operation; c) Affordable and Reusable; d) Greater flexibility than existing systems; e) Part of NASA'S study of "Green Propellants." Hot-fire testing validated performance and functionality of thruster. Thruster's dependence on mixture ratio has been evaluated. Data has been used to calculate performance parameters such as thrust and Isp. Data has been compared with previous test results to verify reliability and repeatability. Thruster was found to have an Isp of 131 s and 82 lbf thrust at a mixture ratio of 1.62.

Propellant Readiness Level: A Methodological Approach to Propellant Characterization

NASA Technical Reports Server (NTRS)

Bossard, John A.; Rhys, Noah O.

2010-01-01

A methodological approach to defining propellant characterization is presented. The method is based on the well-established Technology Readiness Level nomenclature. This approach establishes the Propellant Readiness Level as a metric for ascertaining the readiness of a propellant or a propellant combination by evaluating the following set of propellant characteristics: thermodynamic data, toxicity, applications, combustion data, heat transfer data, material compatibility, analytical prediction modeling, injector/chamber geometry, pressurization, ignition, combustion stability, system storability, qualification testing, and flight capability. The methodology is meant to be applicable to all propellants or propellant combinations; liquid, solid, and gaseous propellants as well as monopropellants and propellant combinations are equally served. The functionality of the proposed approach is tested through the evaluation and comparison of an example set of hydrocarbon fuels.

Producing propellants from water in lunar soil using solar lasers

NASA Astrophysics Data System (ADS)

de Morais Mendonca Teles, Antonio

The exploration of the Solar System is directly related to the efficiency of engines designed to explore it, and consequently, to the propulsion techniques, materials and propellants for those engines. With the present day propulsion techniques it is necessary great quantities of propellants to impulse a manned spacecraft to Mars and beyond in the Solar System, which makes these operations financially very expensive because of the costs involved in launching it from planet Earth, due to its high gravity field strength. To solve this problem, it is needed a planetary place with smaller gravity field strength, near to the Earth and with great quantities of substances at the surface necessary for the in-situ production of propellants for spacecrafts. The only place available is Earth's natural satellite the Moon. So, here in this paper, I propose the creation of a Lunar Propellant Manufacturer. It is a robot-spacecraft which can be launched from Earth using an Energia Rocket, and to land on the Moon in an area (principally near to the north pole where it was discovered water molecules ice recently) with great quantities of oxygen and hydrogen (propellants) in the silicate soil, previously observed and mapped by spacecrafts in lunar orbit, for the extraction of those molecules from the soil and the in-situ production of the necessary propellants. The Lunar Propellant Manufacturer (LPM) spacecraft consists of: 1) a landing system with four legs (extendable) and rovers -when the spacecraft touches down, the legs retract in order that two apparatuses, analogue to tractor's wheeled belts parallel sided and below the spacecraft, can touch firmly the ground -it will be necessary for the displacement of the spacecraft to new areas with richer propellants content, when the early place has already exhausted in propellants; 2) a digging machine -a long, resistant extendable arm with an excavator hand, in the outer part of the spacecraft -it will extend itself to the ground, collect soil and retract itself to put the material on the top of the spacecraft inside a hole which will be opened; 3) an infrared laser based on solar electrical energy -a "solar laser" -when the soil be inside the chamber inside the spacecraft, the solar laser will be turned on and it will strike against the soil, heating it up, and release all oxygen and hydrogen from it. The oxygen and hydrogen molecules will be separated from the rest of the material by a mass spectrometer and they will be liquefied by thermal and pressure internal control sub-systems of the spacecraft, and pumped to vessels in a way similar to a micro-industrial line production process; the vessels with the propellants will be then ready to be taken by astronauts, from a small door outside the LPM. The shape of this spacecraft must be conical in order to not unbalance it during the landing and roving maneuvers and soil cargoes, and it will be shielded externally from heat and radiation from the Sun, and micrometeoroids, to prevent the internal thermal conduction and electronic operations from damaging. A solar array externally deployed can produce 44 KW of electric soil energy for the production process. This miniature chemical-processing plant can possibly have an output of 100 Kg of liquid oxygen and 200 Kg of liquid hydrogen per day. Telecommunications with Earth will provide the onboard computer courses for roving to new mapped areas with richer propellants content in the soil. The spacecraft can weight approximately 6,000 Kg (at launch time from Earth). It will be necessary two LPMs for providing all the liquid oxygen and hydrogen needed to supply spacecrafts next to a semi-permanent small manned lunar base. With the Lunar Propellant Manufacturer it will solve the problem of not-expensively producing great quantities of propellants for a manned spacecraft to explore Mars and beyond In the Solar System.

Cryogenic propulsion for the Titan Orbiter Polar Surveyor (TOPS) mission

NASA Astrophysics Data System (ADS)

Mustafi, S.; DeLee, C.; Francis, J.; Li, X.; McGuinness, D.; Nixon, C. A.; Purves, L.; Willis, W.; Riall, S.; Devine, M.; Hedayat, A.

2016-03-01

Liquid hydrogen (LH2) and liquid oxygen (LO2) cryogenic propellants can dramatically enhance NASA's ability to explore the solar system due to their superior specific impulse (Isp) capability. Although these cryogenic propellants can be challenging to manage and store, they allow significant mass advantages over traditional hypergolic propulsion systems and are therefore enabling for many planetary science missions. New cryogenic storage techniques such as subcooling and the use of advanced insulation and low thermal conductivity support structures will allow for the long term storage and use of cryogenic propellants for solar system exploration and hence allow NASA to deliver more payloads to targets of interest, launch on smaller and less expensive launch vehicles, or both. These new cryogenic storage technologies were implemented in a design study for the Titan Orbiter Polar Surveyor (TOPS) mission, with LH2 and LO2 as propellants, and the resulting spacecraft design was able to achieve a 43% launch mass reduction over a TOPS mission, that utilized a traditional hypergolic propulsion system with mono-methyl hydrazine (MMH) and nitrogen tetroxide (NTO) propellants. This paper describes the cryogenic propellant storage design for the TOPS mission and demonstrates how these cryogenic propellants are stored passively for a decade-long Titan mission that requires the cryogenics propellants to be stored for 8.5 years.

[Progress in the protective medicine against [correction of aganist] rocket propellents].

PubMed

Hu, W X; Tan, C Y; Tan, S J; Jiang, J

1999-12-01

To review the progress in the major assignment, the organization and implementation of protection against liquid rocket propellent. The safety detection methods of the rocket [correction of rocked] propellent in the launching field were also discussed. Three steps of the sanitation and protection of the liquid propellent, the toxicity and the toxicology of hydrazine on central nervous system, blood circulatory system, assimilation system, respiratory system, immune system, liver, kidney, eye, skin and its hereditary toxicology were described. In addition, the clinical types of poisoning, the current principle and the common ways of the prevention and treatment of hydrazine and nitrogen oxides poisoning were summarized.

Application of X-ray television image system to observation in solid rocket motor

NASA Astrophysics Data System (ADS)

Fujiwara, T.; Ito, K.; Tanemura, T.; Shimizu, M.; Godai, T.

The X-ray television image system is used to observe the solid propellant burning surface during rocket motor operation as well as to inspect defects in solid rocket motors in a real time manner. This system can test 200 mm diameter dummy propellant rocket motors with under 2 percent discriminative capacity. Viewing of a 50 mm diameter internal-burning rocket motor, propellant burning surface time transition and propellant burning process of the surroundings of artificial defects were satisfactorily observed. The system was demonstrated to be effective for nondestructive testing and combustion research of solid rocket motors.

Air-fluidized grains as a model system: Self-propelling and jamming

NASA Astrophysics Data System (ADS)

Daniels, Lynn J.

This thesis examines two concepts -- self-propelling and jamming -- that have been employed to unify disparate non-equilibrium systems, in the context of a monolayer of grains fluidized by a temporally and spatially homogeneous upflow of air. The first experiment examines the single particle dynamics of air-fluidized rods. For Brownian rods, equipartition of energy holds and rotational motion sets a timescale after which directional memory is lost. Air-fluidized rods no longer obey equipartion; they self-propel, moving preferentially along their long axis. We show that self-propelling can be treated phenomenologically as an enhanced memory effect causing directional memory to persist for times longer than expected for thermal systems. The second experiment studies dense collections of self-propelling air-fluidized rods. We observe collective propagating modes that give rise to anomalously large fluctuations in the local number density. We quantify these compression waves by calculating the dynamic structure factor and show that the wavespeed is weakly linear with increasing density. It has been suggested that the observed behavior might be explained using the framework put forth by Baskaran et al. [12]. The third experiment seeks to determine whether a force analogous to the critical Casimir force in fluids exists for a large sphere fluidized in the presence of a background of smaller spheres. The behavior of such a large sphere is fully characterized showing that, rather than behaving like a sphere driven by turbulence, the large ball self-propels. We also show that the background is responsible for the purely attractive, intermediate-ranged interaction force between two simultaneously-fluidized large balls. The final experiment seeks to determine what parameters control the diverging relaxation timescale associated with the jamming transition. By tilting our apparatus, we quantify pressure, packing fraction, and temperature simultaneously with dynamics as we approach jamming. We obtain an equation of state that agrees well with simulation and free volume theory. We collapse the relaxation time by defining a time- and energy-scale using pressure, consistent with recent simulation [82]. These experiments are further confirmation of the universality of the concepts of self-propelling and jamming.

Wall mounted heat exchanger characterization. [cryogenic propellant tanks

NASA Technical Reports Server (NTRS)

Bullard, B. R.

1975-01-01

Analytical models are presented for describing the heat and mass transfer and the energy distribution in the contents of a cryogenic propellant tank, under varying gravity levels. These models are used to analytically evaluate the effectiveness of a wall heat exchanger as a means of controlling the pressure in the tank during flight and during fill operations. Pressure and temperature histories are presented for tanks varying in size from 4 to 22.5 feet in diameter and gravity levels from 0-1. Results from the subscale test program, utilizing both non-cryogenic and cryogenic fluid, designed to evaluate a tank wall heat exchanger are described and compared with the analytical models. Both the model and test results indicate that a passive tank wall heat exchanger can effectively control tank pressure. However, the weight of such a system is considerably higher than that of an active mixer system.

14 CFR 23.771 - Pilot compartment.

Code of Federal Regulations, 2010 CFR

2010-01-01

... the propellers so that no part of the pilot or the controls lies in the region between the plane of... propeller hub making an angle of 5 degrees forward or aft of the plane of rotation of the propeller. [Doc...

Propellants and Life Support SCAPE Suit and ECU Capability

NASA Technical Reports Server (NTRS)

Goetzfried, Andreas

2011-01-01

This presentation outlines the details for a conference booth that exhibits the Propellant Handlers Ensemble (PHE) and the Environmental Control Unit (ECU) for personnel loading propellants. A demonstration of the ECU Loading will be performed at the conference.

Restoring Redundancy to the Wilkinson Microwave Anisotrophy Probe Propulsion System

NASA Technical Reports Server (NTRS)

O'Donnell, James R., Jr.; Davis, Gary T.; Ward, David K.

2004-01-01

The Wilkinson Microwave Anisotropy Probe is a follow-on to the Differential Microwave Radiometer instrument on the Cosmic Background Explorer. Attitude control system engineers discovered sixteen months before launch that configuration changes after the critical design review had resulted in a significant migration of the spacecraft's center of mass. As a result, the spacecraft no longer had a viable backup control mode in the event of a failure of the negative pitch-axis thruster. A tiger team was formed and identified potential solutions to this problem, such as adding thruster-plume shields to redirect thruster torque, adding or removing mass from the spacecraft, adding an additional thruster, moving thrusters, bending thruster nozzles or propellant tubing, or accepting the loss of redundancy. The project considered the impacts on mass, cost, fuel budget, and schedule for each solution, and decided to bend the propellant tubing of the two roll-control thrusters to allow the pair to be used for backup control in the negative pitch axis. This paper discusses the problem and the potential solutions, and documents the hardware and software changes and verification performed. Flight data are presented to show the on-orbit performance of the propulsion system and lessons learned are described.

LOX/hydrocarbon auxiliary propulsion system study

NASA Technical Reports Server (NTRS)

Orton, G. F.; Mark, T. D.; Weber, D. D.

1982-01-01

Liquid oxygen/hydrocarbon propulsion systems applicable to a second generation orbiter OMS/RCS were compared, and major system/component options were evaluated. A large number of propellant combinations and system concepts were evaluated. The ground rules were defined in terms of candidate propellants, system/component design options, and design requirements. System and engine component math models were incorporated into existing computer codes for system evaluations. The detailed system evaluations and comparisons were performed to identify the recommended propellant combination and system approach.

Mission demonstration concept for the long-duration storage and transfer of cryogenic propellants

NASA Astrophysics Data System (ADS)

McLean, C.; Deininger, W.; Ingram, K.; Schweickart, R.; Unruh, B.

This paper describes an experimental platform that will demonstrate the major technologies required for the handling and storage of cryogenic propellants in a low-to-zero-g environment. In order to develop a cost-effective, high value-added demonstration mission, a review of the complete mission concept of operations (CONOPS) was performed. The overall cost of such a mission is driven not only by the spacecraft platform and on-orbit experiments themselves, but also by the complexities of handling cryogenic propellants during ground-processing operations. On-orbit storage methodologies were looked at for both passive and active systems. Passive systems rely purely on isolation of the stored propellant from environmental thermal loads, while active cooling employs cryocooler technologies. The benefit trade between active and passive systems is mission-dependent due to the mass, power, and system-level penalties associated with active cooling systems. The experimental platform described in this paper is capable of demonstrating multiple advanced micro-g cryogenic propellant management technologies. In addition to the requirements of demonstrating these technologies, the methodology of propellant transfer must be evaluated. The handling of multiphase liquids in micro-g is discussed using flight-heritage micro-g propellant management device technologies as well as accelerated tank stratification for access to vapor-free or liquid-free propellants. The mission concept presented shows the extensibility of the experimental platform to demonstrate advanced cryogenic components and technologies, propellant transfer methodologies, as well as the validation of thermal and fluidic models, from subscale tankage to an operational architecture.

Reduced Boil-Off System Sizing

NASA Technical Reports Server (NTRS)

Guzik, Monica C.; Plachta, David W.; Feller, Jeffrey R.

2015-01-01

NASA is currently developing cryogenic propellant storage and transfer systems for future space exploration and scientific discovery missions by addressing the need to raise the technology readiness level of cryogenic fluid management technologies. Cryogenic propellants are baselined in many propulsion systems due to their inherently high specific impulse; however, their low boiling points can cause substantial boil-off losses over time. Recent efforts such as the Reduced Boil-off Testing and the Active Thermal Control Scaling Study provide important information on the benefit of an active cooling system applied to LH2 propellant storage. Findings show that zero-boil off technologies can reduce overall mass in LH2 storage systems when low Earth orbit loiter periods extend beyond two months. A significant part of this mass reduction is realized by integrating two stages of cooling: a 20 K stage to intercept heat at the tank surface, and a 90 K stage to reduce the heat entering the less efficient 20 K stage. A missing element in previous studies, which is addressed in this paper, is the development of a direct method for sizing the 90 K cooling stage. Such a method requires calculation of the heat entering both the 90 K and 20 K stages as compared to the overall system masses, and is reliant upon the temperature distribution, performance, and unique design characteristics of the system in question. By utilizing the known conductance of a system without active thermal control, the heat being intercepted by a 90 K stage can be calculated to find the resultant lift and mass of each active thermal control stage. Integral to this is the thermal conductance of the cooling straps and the broad area cooling shield, key parts of the 90 K stage. Additionally, a trade study is performed to show the ability of the 90 K cooling stage to reduce the lift on the 20 K cryocooler stage, which is considerably less developed and efficient than 90 K cryocoolers.

Cold Helium Gas Pressurization For Spacecraft Cryogenic Propulsion Systems

NASA Technical Reports Server (NTRS)

Morehead, Robert L.; Atwell. Matthew J.; Hurlbert, Eric A.; Melcher, J. C.

2017-01-01

To reduce the dry mass of a spacecraft pressurization system, helium pressurant may be stored at low temperature and high pressure to increase mass in a given tank volume. Warming this gas through an engine heat exchanger prior to tank pressurization both increases the system efficiency and simplifies the designs of intermediate hardware such as regulators, valves, etc. since the gas is no longer cryogenic. If this type of cold helium pressurization system is used in conjunction with a cryogenic propellant, though, a loss in overall system efficiency can be expected due to heat transfer from the warm ullage gas to the cryogenic propellant which results in a specific volume loss for the pressurant, interpreted as the Collapse Factor. Future spacecraft with cryogenic propellants will likely have a cold helium system, with increasing collapse factor effects as vehicle sizes decrease. To determine the collapse factor effects and overall implementation strategies for a representative design point, a cold helium system was hotfire tested on the Integrated Cryogenic Propulsion Test Article (ICPTA) in a thermal vacuum environment at the NASA Glenn Research Center Plum Brook Station. The ICPTA vehicle is a small lander-sized spacecraft prototype built at NASA Johnson Space Center utilizing cryogenic liquid oxygen/liquid methane propellants and cryogenic helium gas as a pressurant to operate one 2,800lbf 5:1 throttling main engine, two 28lbf Reaction Control Engines (RCE), and two 7lbf RCEs (Figure 1). This vehicle was hotfire tested at a variety of environmental conditions at NASA Plum Brook, ranging from ambient temperature/simulated high altitude, deep thermal/high altitude, and deep thermal/high vacuum conditions. A detailed summary of the vehicle design and testing campaign may be found in Integrated Cryogenic Propulsion Test Article Thermal Vacuum Hotfire Testing, AIAA JPC 2017.

Low-thrust chemical propulsion system propellant expulsion and thermal conditioning study. Executive summary

NASA Technical Reports Server (NTRS)

Merino, F.; Wakabayashi, I.; Pleasant, R. L.; Hill, M.

1982-01-01

Preferred techniques for providing abort pressurization and engine feed system net positive suction pressure (NPSP) for low thrust chemical propulsion systems (LTPS) were determined. A representative LTPS vehicle configuration is presented. Analysis tasks include: propellant heating analysis; pressurant requirements for abort propellant dump; and comparative analysis of pressurization techniques and thermal subcoolers.

Determination of solid-propellant transient regression rates using a microwave Doppler shift technique

NASA Technical Reports Server (NTRS)

Strand, L. D.; Schultz, A. L.; Reedy, G. K.

1972-01-01

A microwave Doppler shift system, with increased resolution over earlier microwave techniques, was developed for the purpose of measuring the regression rates of solid propellants during rapid pressure transients. A continuous microwave beam is transmitted to the base of a burning propellant sample cast in a metal waveguide tube. A portion of the wave is reflected from the regressing propellant-flame zone interface. The phase angle difference between the incident and reflected signals and its time differential are continuously measured using a high resolution microwave network analyzer and related instrumentation. The apparent propellant regression rate is directly proportional to this latter differential measurement. Experiments were conducted to verify the (1) spatial and time resolution of the system, (2) effect of propellant surface irregularities and compressibility on the measurements, and (3) accuracy of the system for quasi-steady-state regression rate measurements. The microwave system was also used in two different transient combustion experiments: in a rapid depressurization bomb, and in the high-frequency acoustic pressure environment of a T-burner.

In-Space Propellant Production Using Water

NASA Technical Reports Server (NTRS)

Notardonato, William; Johnson, Wesley; Swanger, Adam; McQuade, William

2012-01-01

A new era of space exploration is being planned. Manned exploration architectures under consideration require the long term storage of cryogenic propellants in space, and larger science mission directorate payloads can be delivered using cryogenic propulsion stages. Several architecture studies have shown that in-space cryogenic propulsion depots offer benefits including lower launch costs, smaller launch vehicles, and enhanced mission flexibility. NASA is currently planning a Cryogenic Propellant Storage and Transfer (CPST) technology demonstration mission that will use existing technology to demonstrate long duration storage, acquisition, mass gauging, and transfer of liquid hydrogen in low Earth orbit. This mission will demonstrate key technologies, but the CPST architecture is not designed for optimal mission operations for a true propellant depot. This paper will consider cryogenic propellant depots that are designed for operability. The operability principles considered are reusability, commonality, designing for the unique environment of space, and use of active control systems, both thermal and fluid. After considering these operability principles, a proposed depot architecture will be presented that uses water launch and on orbit electrolysis and liquefaction. This could serve as the first true space factory. Critical technologies needed for this depot architecture, including on orbit electrolysis, zero-g liquefaction and storage, rendezvous and docking, and propellant transfer, will be discussed and a developmental path forward will be presented. Finally, use of the depot to support the NASA Science Mission Directorate exploration goals will be presented.

Earth-to-orbit propellant transportation overview

NASA Technical Reports Server (NTRS)

Fester, D.

1984-01-01

The transportation of large quantities of cryogenic propellants which are needed to support Space Station/OTV operation is discussed. Two ways to send propellants into space are: transporting them in dedicated tankers or scavenging unused STS propellant. Scavenging propellant, both with and without an aft cargo carrier system is examined. An average of two to four flights per year can be saved by scavenging and manifesting propellant as payload. Addition of an aft cargo carrier permits loading closer to maximum, reduces the required number of flights, and reduces the propellant available for scavenging. Sufficient propellant remains, however, for OTV needs.

Liquid oscillations in the tanks' level sensors of aerospace objects

NASA Astrophysics Data System (ADS)

Davydov, I. E.; Davydov, E. I.

2017-01-01

Here are considered reasons of longitudinal oscillations appearance in the propellant consumption control system and ways of reducing them. Estimation of the most significant disturbing factors, which influence on the accuracy of the whole system's work, has been made. There is set the formation of geometric layout of sensitive elements in the system's level sensors.

Credit PSR. This interior view of the building equipment room ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Credit PSR. This interior view of the building equipment room displays heat exchangers and fan units with insulated piping for hot and cold water at left. Environmental controls and fire fighting system controls appear at right - Jet Propulsion Laboratory Edwards Facility, Propellant Curing Building, Edwards Air Force Base, Boron, Kern County, CA

Pulsed plasma solid propellant microthruster for the synchronous meteorological satellite. Task 4: Engineering model fabrication and test report

NASA Technical Reports Server (NTRS)

Guman, W. J. (Editor)

1972-01-01

Two flight prototype solid propellant pulsed plasma microthruster propulsion systems for the SMS satellite were fabricated, assembled and tested. The propulsion system is a completely self contained system requiring only three electrical inputs to operate: a 29.4 volt power source, a 28 volt enable signal and a 50 millsec long command fire signal that can be applied at any rate from 50 ppm to 110 ppm. The thrust level can be varied over a range 2.2 to 1 at constant impulse bit amplitude. By controlling the duration of the 28 volt enable either steady state thrust or a series of discrete impulse bits can be generated. A new technique of capacitor charging was implemented to reduce high voltage stress on energy storage capacitors.

Orbital Maneuvering system design evolution

NASA Technical Reports Server (NTRS)

Gibson, C.; Humphries, C.

1985-01-01

Preliminary design considerations and changes made in the baseline space shuttle orbital maneuvering system (OMS) to reduce cost and weight are detailed. The definition of initial subsystem requirements, trade studies, and design approaches are considered. Design features of the engine, its injector, combustion chamber, nozzle extension and bipropellant valve are illustrated and discussed. The current OMS consists of two identical pods that use nitrogen tetroxide (NTO) and monomethylhydrazine (MMH) propellants to provide 1000 ft/sec of delta velocity for a payload of 65,000 pounds. Major systems are pressurant gas storage and control, propellant storage supply and quantity measurement, and the rocket engine, which includes a bipropellant valve, an injector/thrust chamber, and a nozzle. The subsystem provides orbit insertion, circularization, and on orbit and deorbit capability for the shuttle orbiter.

Analyzing the Use of Gaseous Helium as a Pressurant with Cryogenic Propellants with Thermodynamic Venting System Modelling and Test Data

NASA Technical Reports Server (NTRS)

Hedayat, A.; Nelson, S.L.; Hastings, L.J.; Flachbart, R.H.; Vermillion, D.J.; Tucker, S.P.

2007-01-01

Cryogens are viable candidate propellants for NASA's Lunar and Mars exploration programs. To provide adequate mass flow to the system's engines and/or to prevent feed system cavitation, gaseous helium (GHe) is frequently considered as a pressurant. During low gravity operations, a Thermodynamic Venting System (TVS) is designed to maintain tank pressure during low gravity operations without propellant resettling. Therefore, a series of tests were conducted in the Multi-purpose Hydrogen Test Bed (MHTB) of Marshall Space Flight Center (MSFC) in order to evaluate the effects of GHe pressurant on pressure control performance of a TVS with liquid hydrogen (LH2) and nitrogen (LN2) as the test liquids. The TVS used in these test series consists of a recirculation pump, Joule-Thomson (J-T) expansion valve, and a parallel flow concentric tube heat exchanger combined with a longitudinal spray bar. Using a small amount of liquid extracted from the tank recirculation line, passing it through the J-T valve, and then through the heat exchanger, thermal energy is extracted from the bulk liquid and ullage thereby enabling pressure control. The LH2/GHe tests were performed at fill levels of 90%, 50%, and 25% and LN2/GHe tests were conducted at fill levels of 50% and 25%. Moreover, each test was conducted with a specified tank ullage pressure control band. A one-dimensional TVS performance program was used to analyze and correlate the test data. Predictions and comparisons with test data of ullage pressure and temperature and bulk liquid saturation pressure and temperature with test data are presented.

Cryogenic Boil-Off Reduction System

NASA Astrophysics Data System (ADS)

Plachta, David W.; Guzik, Monica C.

2014-03-01

A computational model of the cryogenic boil-off reduction system being developed by NASA as part of the Cryogenic Propellant Storage and Transfer technology maturation project has been applied to a range of propellant storage tanks sizes for high-performing in-space cryogenic propulsion applications. This effort focuses on the scaling of multi-layer insulation (MLI), cryocoolers, broad area cooling shields, radiators, solar arrays, and tanks for liquid hydrogen propellant storage tanks ranging from 2 to 10 m in diameter. Component scaling equations were incorporated into the Cryogenic Analysis Tool, a spreadsheet-based tool used to perform system-level parametric studies. The primary addition to the evolution of this updated tool is the integration of a scaling method for reverse turbo-Brayton cycle cryocoolers, as well as the development and inclusion of Self-Supporting Multi-Layer Insulation. Mass, power, and sizing relationships are traded parametrically to establish the appropriate loiter period beyond which this boil-off reduction system application reduces mass. The projected benefit compares passive thermal control to active thermal control, where active thermal control is evaluated for reduced boil-off with a 90 K shield, zero boil-off with a single heat interception stage at the tank wall, and zero boil-off with a second interception stage at a 90 K shield. Parametric studies show a benefit over passive storage at loiter durations under one month, in addition to showing a benefit for two-stage zero boil-off in terms of reducing power and mass as compared to single stage zero boil-off. Furthermore, active cooling reduces the effect of varied multi-layer insulation performance, which, historically, has been shown to be significant.

A translational velocity command system for VTOL low speed flight

NASA Technical Reports Server (NTRS)

Merrick, V. K.

1982-01-01

A translational velocity flight controller, suitable for very low speed maneuvering, is described and its application to a large class of VTOL aircraft from jet lift to propeller driven types is analyzed. Estimates for the more critical lateral axis lead to the conclusion that the controller would provide a jet lift (high disk loading) VTOL aircraft with satisfactory "hands off" station keeping in operational conditions more stringent than any specified in current or projected requirements. It also seems likely that ducted fan or propeller driven (low disk loading) VTOL aircraft would have acceptable hovering handling qualities even in high turbulence, although in these conditions pilot intervention to maintain satisfactory station keeping would probably be required for landing in restricted areas.

Aircraft control system

NASA Technical Reports Server (NTRS)

Kendall, Greg T. (Inventor); Lisoski, Derek L. (Inventor)

2007-01-01

A solar rechargeable, long-duration, span-loaded flying wing, having no fuselage or rudder. Having a two-hundred foot wingspan that mounts photovoltaic cells on most all of the wing's top surface, the aircraft uses only differential thrust of its eight propellers to turn, pitch and yaw. The wing is configured to deform under flight loads to position the propellers such that the control can be achieved. Each of five segments of the wing has one or more motors and photovoltaic arrays, and produces its own lift independent of the other segments, to avoid loading them. Five two-sided photovoltaic arrays, in all, are mounted on the wing, and receive photovoltaic energy both incident on top of the wing, and which is incident also from below, through a bottom, transparent surface.

Solar Thermal Upper Stage Liquid Hydrogen Pressure Control Testing

NASA Technical Reports Server (NTRS)

Moore, J. D.; Otto, J. M.; Cody, J. C.; Hastings, L. J.; Bryant, C. B.; Gautney, T. T.

2015-01-01

High-energy cryogenic propellant is an essential element in future space exploration programs. Therefore, NASA and its industrial partners are committed to an advanced development/technology program that will broaden the experience base for the entire cryogenic fluid management community. Furthermore, the high cost of microgravity experiments has motivated NASA to establish government/aerospace industry teams to aggressively explore combinations of ground testing and analytical modeling to the greatest extent possible, thereby benefitting both industry and government entities. One such team consisting of ManTech SRS, Inc., Edwards Air Force Base, and Marshall Space Flight Center (MSFC) was formed to pursue a technology project designed to demonstrate technology readiness for an SRS liquid hydrogen (LH2) in-space propellant management concept. The subject testing was cooperatively performed June 21-30, 2000, through a partially reimbursable Space Act Agreement between SRS, MSFC, and the Air Force Research Laboratory. The joint statement of work used to guide the technical activity is presented in appendix A. The key elements of the SRS concept consisted of an LH2 storage and supply system that used all of the vented H2 for solar engine thrusting, accommodated pressure control without a thermodynamic vent system (TVS), and minimized or eliminated the need for a capillary liquid acquisition device (LAD). The strategy was to balance the LH2 storage tank pressure control requirements with the engine thrusting requirements to selectively provide either liquid or vapor H2 at a controlled rate to a solar thermal engine in the low-gravity environment of space operations. The overall test objective was to verify that the proposed concept could enable simultaneous control of LH2 tank pressure and feed system flow to the thruster without necessitating a TVS and a capillary LAD. The primary program objectives were designed to demonstrate technology readiness of the SRS concept at a system level as a first step toward actual flight vehicle demonstrations. More specific objectives included testing the pressure and feed control system concept hardware for functionality, operability, and performance. Valuable LH2 thermodynamic and fluid dynamics data were obtained for application to both the SRS concept and to future missions requiring space-based cryogen propellant management.

14 CFR 23.937 - Turbopropeller-drag limiting systems.

Code of Federal Regulations, 2010 CFR

2010-01-01

... actuated after engine power loss, can move the propeller blades toward the feather position to reduce... General § 23.937 Turbopropeller-drag limiting systems. (a) Turbopropeller-powered airplane propeller-drag... normal or emergency operation results in propeller drag in excess of that for which the airplane was...

14 CFR 23.937 - Turbopropeller-drag limiting systems.

Code of Federal Regulations, 2012 CFR

2012-01-01

... actuated after engine power loss, can move the propeller blades toward the feather position to reduce... General § 23.937 Turbopropeller-drag limiting systems. (a) Turbopropeller-powered airplane propeller-drag... normal or emergency operation results in propeller drag in excess of that for which the airplane was...

14 CFR 23.937 - Turbopropeller-drag limiting systems.

Code of Federal Regulations, 2014 CFR

2014-01-01

... actuated after engine power loss, can move the propeller blades toward the feather position to reduce... General § 23.937 Turbopropeller-drag limiting systems. (a) Turbopropeller-powered airplane propeller-drag... normal or emergency operation results in propeller drag in excess of that for which the airplane was...

14 CFR 23.937 - Turbopropeller-drag limiting systems.

Code of Federal Regulations, 2013 CFR

2013-01-01

... actuated after engine power loss, can move the propeller blades toward the feather position to reduce... General § 23.937 Turbopropeller-drag limiting systems. (a) Turbopropeller-powered airplane propeller-drag... normal or emergency operation results in propeller drag in excess of that for which the airplane was...

14 CFR 23.937 - Turbopropeller-drag limiting systems.

Code of Federal Regulations, 2011 CFR

2011-01-01

... actuated after engine power loss, can move the propeller blades toward the feather position to reduce... General § 23.937 Turbopropeller-drag limiting systems. (a) Turbopropeller-powered airplane propeller-drag... normal or emergency operation results in propeller drag in excess of that for which the airplane was...

System design of the Pioneer Venus spacecraft. Volume 10: Propulsion/orbit insertion subsystem studies

NASA Technical Reports Server (NTRS)

Rosenstein, B. J.

1973-01-01

The Pioneer Venus orbiter and multiprobe missions require spacecraft maneuvers for successful accomplishment. This report presents the results of studies performed to define the propulsion subsystems required to perform those maneuvers. Primary goals were to define low mass subsystems capable of performing the required missions with a high degree of reliability for low cost. A review was performed of all applicable propellants and thruster types, as well as propellant management techniques. Based on this review, a liquid monopropellant hydrazine propulsion subsystem was selected for all multiprobe mission maneuvers, and for all orbiter mission maneuvers except orbit insertion. A pressure blowdown operating mode was selected using helium as the pressurizing gas. The forces associated with spacecraft rotations were used to control the liquid-gas interface and resulting propellant orientation within the tank.

Reusable module for the storage, transportation, and supply of multiple propellants in a space environment

NASA Technical Reports Server (NTRS)

Mazanek, Daniel D. (Inventor); Mankins, John C. (Inventor)

2004-01-01

A space module has an outer structure designed for traveling in space, a docking mechanism for facilitating a docking operation therewith in space, a first storage system storing a first propellant that burns as a result of a chemical reaction therein, a second storage system storing a second propellant that burns as a result of electrical energy being added thereto, and a bi-directional transfer interface coupled to each of the first and second storage systems to transfer the first and second propellants into and out thereof. The space module can be part of a propellant supply architecture that includes at least two of the space modules placed in an orbit in space.

The PROPEL Electrodynamic Tether Demonstration Mission

NASA Technical Reports Server (NTRS)

Bilen, Sven G.; Johnson, C. Les; Wiegmann, Bruce M.; Alexander, Leslie; Gilchrist, Brian E.; Hoyt, Robert P.; Elder, Craig H.; Fuhrhop, Keith P.; Scadera, Michael

2012-01-01

The PROPEL ("Propulsion using Electrodynamics") mission will demonstrate the operation of an electrodynamic tether propulsion system in low Earth orbit and advance its technology readiness level for multiple applications. The PROPEL mission has two primary objectives: first, to demonstrate the capability of electrodynamic tether technology to provide robust and safe, near-propellantless propulsion for orbit-raising, de-orbit, plane change, and station keeping, as well as to perform orbital power harvesting and formation flight; and, second, to fully characterize and validate the performance of an integrated electrodynamic tether propulsion system, qualifying it for infusion into future multiple satellite platforms and missions with minimal modification. This paper provides an overview of the PROPEL system and design reference missions; mission goals and required measurements; and ongoing PROPEL mission design efforts.

75 FR 18774 - Airworthiness Directives; McCauley Propeller Systems Model 4HFR34C653/L106FA Propellers

Federal Register 2010, 2011, 2012, 2013, 2014

2010-04-13

... (ECI) of the propeller hub for cracks. That proposed AD resulted from reports of 3 hubs found cracked... NPRM results from reports of 7 additional hubs found cracked during propeller overhaul, totaling 10 cracked hubs. We are proposing this supplemental AD to prevent failure of the propeller hub, which could...

Space storable propellant acquisition system

NASA Technical Reports Server (NTRS)

Tegart, J. R.; Uney, P. E.; Anderson, J. E.; Fester, D. A.

1972-01-01

Surface tension propellant acquisition concepts for an advanced spacecraft propulsion system having a 10-year mission capability were investigated. Surface tension systems were specified because they were shown to be the best propellant acquisition technique for various interplanetery spacecraft in a prior study. A variety of surface tension concepts for accomplishing propellant acquisition were formulated for the baseline space storable propulsion module and Jupiter Orbiter mission. Analyses and evaluations were then conducted on each candidate concept to assess fabricability, performance capability, and spacecraft compatibility. A comparative evaluation of the results showed the Fruhof-class of low-g surface tension systems to be preferred for these interplanetary applications.

14 CFR 35.40 - Functional test.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Functional test. 35.40 Section 35.40... STANDARDS: PROPELLERS Tests and Inspections § 35.40 Functional test. The variable-pitch propeller system must be subjected to the applicable functional tests of this section. The same propeller system used in...

14 CFR 35.40 - Functional test.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Functional test. 35.40 Section 35.40... STANDARDS: PROPELLERS Tests and Inspections § 35.40 Functional test. The variable-pitch propeller system must be subjected to the applicable functional tests of this section. The same propeller system used in...

14 CFR 35.40 - Functional test.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Functional test. 35.40 Section 35.40... STANDARDS: PROPELLERS Tests and Inspections § 35.40 Functional test. The variable-pitch propeller system must be subjected to the applicable functional tests of this section. The same propeller system used in...

JPRS Report, Science & Technology, USSR: Electronics & Electrical Engineering.

DTIC Science & Technology

1988-02-05

Sirena -1 Self-propelled Flaw Detector [PRIBORYI SISTEMY UPRAVLENIYA, Jan 87] 14 Crane Strain-measurement Scales With Data Processing by a Microprocessor...was 3-5 m. 06415/06662 UDC 620.179.1:620.165.29 Algorithimization of Control of Electric Motor Drive of Sirena -1 Self-propelled Flaw Detector...The article describes one of the most optimum algorithms of control of the electric motor drive of the Sirena -1 self-propelled flaw detector

Reciprocating Feed System Development Status

NASA Technical Reports Server (NTRS)

Trewek, Mary (Technical Monitor); Blackmon, James B.; Eddleman, David E.

2005-01-01

The reciprocating feed system (RFS) is an alternative means of providing high pressure propellant flow at low cost and system mass, with high fail-operational reliability. The RFS functions by storing the liquid propellants in large, low-pressure tanks and then expelling each propellant through two or three small, high-pressure tanks. Each RFS tank is sequentially filled, pressurized, expelled, vented, and refilled so as to provide a constant, or variable, mass flow rate to the engine. This type of system is much lighter than a conventional pressure fed system in part due to the greatly reduced amount of inert tank weight. The delivered payload for an RFS is superior to that of conventional pressure fed systems for conditions of high total impulse and it is competitive with turbopump systems, up to approximately 2000 psi. An advanced version of the RFS uses autogenous pressurization and thrust augmentation to achieve higher performance. In this version, the pressurization gases are combusted in a small engine, thus making the pressurization system, in effect, part of the propulsion system. The RFS appears to be much less expensive than a turbopump system, due to reduced research and development cost and hardware cost, since it is basically composed of small high- pressure tanks, a pressurization system, and control valves. A major benefit is the high reliability fail-operational mode; in the event of a failure in one of the three tank-systems, it can operate on the two remaining tanks. Other benefits include variable pressure and flow rates, ease of engine restart in micro-gravity, and enhanced propellant acquisition and control under adverse acceleration conditions. We present a system mass analysis tool that accepts user inputs for various design and mission parameters and calculates such output values payload and vehicle weights for the conventional pressure fed system, the RFS, the Autogenous Pressurization Thrust Augmentation (APTA) RFS, and turbopump systems. Using this tool, a preliminary design of a representative crew exploration vehicle (CEV) has been considered. The design parameters selected for a representative system were modeled after the orbital maneuvering system (OMS) on the Shuttle Orbiter, with an increase of roughly a factor of ten in the delta- V capability and a greater thrust (30,000 lbs, vs. 12,000 lbs). Both storable and cryogenic propellants were considered. Results show that a RFS is a low mass alternative to conventional pressure fed systems, with a substantial increase in payload capability and that it is weight-competitive with turbopump systems at low engine pressure (a few hundred psi); at high engine pressures, the APTA RFS appears to offer the highest payload. We also present the status of the RFS test bed fabrication, assembly, and checkout. This test bed is designed to provide flow rates appropriate for engines in the roughly 10,000 to 30,000 lb thrust range.

Solar Powered Aircraft, Photovoltaic Array/Battery System Tabletop Demonstration: Design and Operation Manual

NASA Technical Reports Server (NTRS)

Colozza, Anthony J.; Scheiman, David A.; Bailey, Sheila (Technical Monitor)

2000-01-01

A system was constructed to demonstrate the power system operation of a solar powered aircraft. The system consists of a photovoltaic (PV) array, a charge controller, a battery, an electric motor and propeller. The system collects energy from the PV array and either utilizes this energy to operate an electric motor or stores it in a rechargeable battery for future use. The system has a control panel which displays the output of the array and battery as well as the total current going to the electric motor. The control panel also has a means for adjusting the output to the motor to control its speed. The entire system is regulated around 12 VDC.

Zero-Propellant Maneuver[TM] Flight Results for 180 deg ISS Rotation

NASA Technical Reports Server (NTRS)

Bedrossian, Nazareth; Bhatt, Sagar; Lammers, Mike; Nguyen, Louis

2007-01-01

This paper presents results for the Zero Propellant Maneuver (ZPM) TradeMark attitude control concept flight demonstration. On March 3, 2007, a ZPM was used to reorient the International Space Station 180 degrees without using any propellant. The identical reorientation performed with thrusters would have burned 110lbs of propellant. The ZPM was a pre-planned trajectory used to command the CMG attitude hold controller to perform the maneuver between specified initial and final states while maintaining the CMGs within their operational limits. The trajectory was obtained from a PseudoSpectral solution to a new optimal attitude control problem. The flight test established the breakthrough capability to simultaneously perform a large angle attitude maneuver and momentum desaturation without the need to use thrusters. The flight implementation did not require any modifications to flight software. This approach is applicable to any spacecraft that are controlled by momentum storage devices.

Molecular System for the Division of Self-Propelled Oil Droplets by Component Feeding.

PubMed

Banno, Taisuke; Toyota, Taro

2015-06-30

Unique dynamics using inanimate molecular assemblies have drawn a great amount of attention for demonstrating prebiomimetic molecular systems. For the construction of an organized logic combining two fundamental dynamics of life, we demonstrate here a molecular system that exhibits both division and self-propelled motion using oil droplets. The key molecule of this molecular system is a novel cationic surfactant containing a five-membered acetal moiety, and the molecular system can feed the self-propelled oil droplet composed of a benzaldehyde derivative and an alkanol. The division dynamics of the self-propelled oil droplets were observed through the hydrolysis of the cationic surfactant in bulk solution. The mechanism of the current dynamics is argued to be based on the supply of "fresh" oil components in the moving oil droplets, which is induced by the Marangoni instability. We consider this molecular system to be a prototype of self-reproducing inanimate molecular assembly exhibiting self-propelled motion.

Space Transportatioin System (STS) propellant scavenging system study. Volume 3: Cost and work breakdown structure-dictionary

NASA Technical Reports Server (NTRS)

1985-01-01

Fundamentally, the volumes of the oxidizer and fuel propellant scavenged from the orbiter and external tank determine the size and weight of the scavenging system. The optimization of system dimensions and weights is stimulated by the requirement to minimize the use of partial length of the orbiter payload bay. Thus, the cost estimates begin with weights established for the optimum design. Both the design, development, test, and evaluation and theoretical first unit hardware production costs are estimated from parametric cost weight scaling relations for four subsystems. For cryogenic propellants, the widely differing characteristics of the oxidizer and the fuel lead to two separate tank subsystems, in addition to the electrical and instrumentation subsystems. Hardwares costs also involve quantity, as an independent variable, since the number of production scavenging systems is not firm. For storable propellants, since the tankage volume of the oxidizer and fuel are equal, the hardware production costs for developing these systems are lower than for cryogenic propellants.

Aerosciences, Aero-Propulsion and Flight Mechanics Technology Development for NASA's Next Generation Launch Technology Program

NASA Technical Reports Server (NTRS)

Cockrell, Charles E., Jr.

2003-01-01

The Next Generation Launch Technology (NGLT) program, Vehicle Systems Research and Technology (VSR&T) project is pursuing technology advancements in aerothermodynamics, aeropropulsion and flight mechanics to enable development of future reusable launch vehicle (RLV) systems. The current design trade space includes rocket-propelled, hypersonic airbreathing and hybrid systems in two-stage and single-stage configurations. Aerothermodynamics technologies include experimental and computational databases to evaluate stage separation of two-stage vehicles as well as computational and trajectory simulation tools for this problem. Additionally, advancements in high-fidelity computational tools and measurement techniques are being pursued along with the study of flow physics phenomena, such as boundary-layer transition. Aero-propulsion technology development includes scramjet flowpath development and integration, with a current emphasis on hypervelocity (Mach 10 and above) operation, as well as the study of aero-propulsive interactions and the impact on overall vehicle performance. Flight mechanics technology development is focused on advanced guidance, navigation and control (GN&C) algorithms and adaptive flight control systems for both rocket-propelled and airbreathing vehicles.

Proceedings of the Ship Control Systems Symposium (5th), Held at U. S. Naval Academy, Annapolis, Maryland on October 30 - November 3, 1978. Volume 4,

DTIC Science & Technology

1978-11-03

Flow The flow characteristics of the ahead and astern valves are given in Figures 2 and 3. These data are used to convert the valve signal from the...velocity and positive propeller rotational speed are given in the David Taylor Model Basin Test Report (Reference 3). These data were put in terms of a...modified advanced coefficient,Qa , and are plotted in Figures 4 and 5. No data for this propeller were available for negative advance velocity and/or

Inflight Characterization of the Cassini Spacecraft Propellant Slosh and Structural Frequencies

NASA Technical Reports Server (NTRS)

Lee, Allan Y.; Stupik, Joan

2015-01-01

While there has been extensive theoretical and analytical research regarding the characterization of spacecraft propellant slosh and structural frequencies, there have been limited studies to compare the analytical predictions with measured flight data. This paper uses flight telemetry from the Cassini spacecraft to get estimates of high-g propellant slosh frequencies and the magnetometer boom frequency characteristics, and compares these values with those predicted by theoretical works. Most Cassini attitude control data are available at a telemetry frequency of 0.5 Hz. Moreover, liquid sloshing is attenuated by propellant management device and attitude controllers. Identification of slosh and structural frequency are made on a best-effort basis. This paper reviews the analytical approaches that were used to predict the Cassini propellant slosh frequencies. The predicted frequencies are then compared with those estimated using telemetry from selected Cassini burns where propellant sloshing was observed (such as the Saturn Orbit Insertion burn). Determination of the magnetometer boom structural frequency is also discussed.

Design and control of an embedded vision guided robotic fish with multiple control surfaces.

PubMed

Yu, Junzhi; Wang, Kai; Tan, Min; Zhang, Jianwei

2014-01-01

This paper focuses on the development and control issues of a self-propelled robotic fish with multiple artificial control surfaces and an embedded vision system. By virtue of the hybrid propulsion capability in the body plus the caudal fin and the complementary maneuverability in accessory fins, a synthesized propulsion scheme including a caudal fin, a pair of pectoral fins, and a pelvic fin is proposed. To achieve flexible yet stable motions in aquatic environments, a central pattern generator- (CPG-) based control method is employed. Meanwhile, a monocular underwater vision serves as sensory feedback that modifies the control parameters. The integration of the CPG-based motion control and the visual processing in an embedded microcontroller allows the robotic fish to navigate online. Aquatic tests demonstrate the efficacy of the proposed mechatronic design and swimming control methods. Particularly, a pelvic fin actuated sideward swimming gait was first implemented. It is also found that the speeds and maneuverability of the robotic fish with coordinated control surfaces were largely superior to that of the swimming robot propelled by a single control surface.

Design and Control of an Embedded Vision Guided Robotic Fish with Multiple Control Surfaces

PubMed Central

Wang, Kai; Tan, Min; Zhang, Jianwei

2014-01-01

This paper focuses on the development and control issues of a self-propelled robotic fish with multiple artificial control surfaces and an embedded vision system. By virtue of the hybrid propulsion capability in the body plus the caudal fin and the complementary maneuverability in accessory fins, a synthesized propulsion scheme including a caudal fin, a pair of pectoral fins, and a pelvic fin is proposed. To achieve flexible yet stable motions in aquatic environments, a central pattern generator- (CPG-) based control method is employed. Meanwhile, a monocular underwater vision serves as sensory feedback that modifies the control parameters. The integration of the CPG-based motion control and the visual processing in an embedded microcontroller allows the robotic fish to navigate online. Aquatic tests demonstrate the efficacy of the proposed mechatronic design and swimming control methods. Particularly, a pelvic fin actuated sideward swimming gait was first implemented. It is also found that the speeds and maneuverability of the robotic fish with coordinated control surfaces were largely superior to that of the swimming robot propelled by a single control surface. PMID:24688413

Integrated Cryogenic Propulsion Test Article Thermal Vacuum Hotfire Testing

NASA Technical Reports Server (NTRS)

Morehead, Robert L.; Melcher, J. C.; Atwell, Matthew J.; Hurlbert, Eric A.

2017-01-01

In support of a facility characterization test, the Integrated Cryogenic Propulsion Test Article (ICPTA) was hotfire tested at a variety of simulated altitude and thermal conditions in the NASA Glenn Research Center Plum Brook Station In-Space Propulsion Thermal Vacuum Chamber (formerly B2). The ICPTA utilizes liquid oxygen and liquid methane propellants for its main engine and four reaction control engines, and uses a cold helium system for tank pressurization. The hotfire test series included high altitude, high vacuum, ambient temperature, and deep cryogenic environments, and several hundred sensors on the vehicle collected a range of system level data useful to characterize the operation of an integrated LOX/Methane spacecraft in the space environment - a unique data set for this propellant combination.

Preliminary Characterization of the Altair Lunar Lander Slosh Dynamics and Some Implications for the Thrust Vector Control Design

NASA Technical Reports Server (NTRS)

Lee, Allan Y.; Strahan, Alan; Tanimoto, Rebekah; Casillas, Arturo

2010-01-01

This paper describes a conceptual design of the Thrust Vector Control (TVC) system and preliminary modeling of propellant slosh, for the Altair Lunar Lander. Altair is a vehicle element of the NASA Constellation Program aimed at returning humans to the moon. Guidance, Navigation, and Control (GN&C) is the measurement and control of spacecraft position, velocity, and attitude in support of mission objectives. One key GN&C function is the commanding of effectors that control attitude and impart delta V on the vehicle, utilizing both reaction control system (RCS) thrusters and throttling and TVC gimbaling of the vehicle main engine. Both the Altair descent and ascent modules carry fuel tanks. During thrusting maneuvers, the sloshing of liquid fuels in partially filled tanks can interact with the controlled system in such a way as to cause the overall system to be unstable. These fuel tanks must be properly placed, relative to the spacecraft's c.m., to avoid any unstable interactions. Following this will be a discussion of propellant slosh modeling work performed for the present vehicle configuration, including slosh frequency and participatory fluid mass predictions. Knowing the range of slosh mode frequencies over mission phases, the TVC bandwidth must be carefully selected so as not to excite the slosh modes at those frequencies. The likely need to increase the damping factor of slosh modes via baffles will also be discussed. To conclude, a discussion of operations procedures aimed at minimizing TVC-slosh interactions will be given.

40 CFR Appendix III to Part 1042 - Not-to-Exceed Zones

Code of Federal Regulations, 2011 CFR

2011-07-01

... marine engines used with controllable-pitch propellers or with electrically coupled propellers, as... CONTROLS CONTROL OF EMISSIONS FROM NEW AND IN-USE MARINE COMPRESSION-IGNITION ENGINES AND VESSELS Pt. 1042... (or at Maximum Test Torque for constant-speed engines). (2) Percent speed means the percentage of...

40 CFR Appendix III to Part 1042 - Not-to-Exceed Zones

Code of Federal Regulations, 2010 CFR

2010-07-01

... marine engines used with controllable-pitch propellers or with electrically coupled propellers, as... CONTROLS CONTROL OF EMISSIONS FROM NEW AND IN-USE MARINE COMPRESSION-IGNITION ENGINES AND VESSELS Pt. 1042... (or at Maximum Test Torque for constant-speed engines). (2) Percent speed means the percentage of...

40 CFR Appendix III to Part 1042 - Not-to-Exceed Zones

Code of Federal Regulations, 2014 CFR

2014-07-01

... marine engines used with controllable-pitch propellers or with electrically coupled propellers, as... CONTROLS CONTROL OF EMISSIONS FROM NEW AND IN-USE MARINE COMPRESSION-IGNITION ENGINES AND VESSELS Pt. 1042... (or at Maximum Test Torque for constant-speed engines). (2) Percent speed means the percentage of...

40 CFR Appendix III to Part 1042 - Not-to-Exceed Zones

Code of Federal Regulations, 2013 CFR

2013-07-01

... marine engines used with controllable-pitch propellers or with electrically coupled propellers, as... CONTROLS CONTROL OF EMISSIONS FROM NEW AND IN-USE MARINE COMPRESSION-IGNITION ENGINES AND VESSELS Pt. 1042... (or at Maximum Test Torque for constant-speed engines). (2) Percent speed means the percentage of...

40 CFR Appendix III to Part 1042 - Not-to-Exceed Zones

Code of Federal Regulations, 2012 CFR

2012-07-01

... marine engines used with controllable-pitch propellers or with electrically coupled propellers, as... CONTROLS CONTROL OF EMISSIONS FROM NEW AND IN-USE MARINE COMPRESSION-IGNITION ENGINES AND VESSELS Pt. 1042... (or at Maximum Test Torque for constant-speed engines). (2) Percent speed means the percentage of...

Thermal Optimization of an On-Orbit Long Duration Cryogenic Propellant Depot

NASA Technical Reports Server (NTRS)

Honour, Ryan; Kwas, Robert; O'Neil, Gary; Kutter, Gary

2012-01-01

A Cryogenic Propellant Depot (CPD) operating in Low Earth Orbit (LEO) could provide many near term benefits to NASA's space exploration efforts. These benefits include elongation/extension of spacecraft missions and requirement reduction of launch vehicle up-mass. Some of the challenges include controlling cryogenic propellant evaporation and managing the high costs and long schedules associated with the new development of spacecraft hardware. This paper describes a conceptual CPD design that is thermally optimized to achieve extremely low propellant boil-off rates. The CPD design is based on existing launch vehicle architecture, and its thermal optimization is achieved using current passive thermal control technology. Results from an integrated thermal model are presented showing that this conceptual CPD design can achieve propellant boil-off rates well under 0.05% per day, even when subjected to the LEO thermal environment.

Thermal Optimization and Assessment of a Long Duration Cryogenic Propellant Depot

NASA Technical Reports Server (NTRS)

Honour, Ryan; Kwas, Robert; O'Neil, Gary; Kutter, Bernard

2012-01-01

A Cryogenic Propellant Depot (CPD) operating in Low Earth Orbit (LEO) could provide many near term benefits to NASA space exploration efforts. These benefits include elongation/extension of spacecraft missions and reduction of launch vehicle up-mass requirements. Some of the challenges include controlling cryogenic propellant evaporation and managing the high costs and long schedules associated with new spacecraft hardware development. This paper describes a conceptual CPD design that is thermally optimized to achieve extremely low propellant boil-off rates. The CPD design is based on existing launch vehicle architecture, and its thermal optimization is achieved using current passive thermal control technology. Results from an integrated thermal model are presented showing that this conceptual CPD design can achieve propellant boil-off rates well under 0.05% per day, even when subjected to the LEO thermal environment.

Sources, control, and effects of noise from aircraft propellers and rotors. [noise prediction (aircraft

NASA Technical Reports Server (NTRS)

Mixson, J. S.; Greene, G. C.; Dempsey, T. K.

1981-01-01

Source noise predictions are compared with measurements for conventional low-speed propellers, for new high speed propellers (propfans), and for a helicopter. Results from a light aircraft demonstration program are described, indicating that about 5-dB reduction of flyover noise can be obtained without significant performance penalty. Sidewall design studies are described for interior noise control in light general aviation aircraft and in large transports using propfan propulsion. The weight of the added acoustic treatment is estimated and tradeoffs between weight and noise reduction are discussed. A laboratory study of passenger response to combined broadband and tonal propeller like noise is described. Subject discomfort ratings of combined tone broadband noises are compared with ratings of broadband (boundary layer) noise alone, and the relative importance of the propeller tones is examined.

49 CFR 236.714 - Cab.

Code of Federal Regulations, 2010 CFR

2010-10-01

... TRANSPORTATION RULES, STANDARDS, AND INSTRUCTIONS GOVERNING THE INSTALLATION, INSPECTION, MAINTENANCE, AND REPAIR OF SIGNAL AND TRAIN CONTROL SYSTEMS, DEVICES, AND APPLIANCES Definitions § 236.714 Cab. The compartment of a locomotive from which the propelling power and power brakes of the train are manually...

In-space propellant logistics. Volume 1: Executive summary

NASA Technical Reports Server (NTRS)

1972-01-01

The study addresses the systems and operational problems associated with the transport, transfer, and storage of cryogenic propellants in low earth orbits. The safety problems connected with in-space propellant logistics operations are also considered.Correlation between the two projects was maintained by including safety considerations, resulting from the system safety analysis, in the trade studies and evaluations of alternate operating concepts in the systems operations analysis.

Active Structural Acoustic Control in an Original A400M Aircraft Structure

NASA Astrophysics Data System (ADS)

Koehne, C.; Sachau, D.; Renger, K.

2016-09-01

Low frequency noise has always been a challenge in propeller driven aircraft. At low frequencies passive noise treatments are not as efficient as active noise reduction systems. The Helmut-Schmidt-University has built up a full-scale test rig with an original A400M aircraft structure. This provides a good opportunity to develop and test active noise reduction systems in a realistic environment. The currently installed system consists of mechanical actuators and acoustical sensors. The actuators are called TVAs (Tuneable Vibration Absorber) and contain two spring-mass systems whose natural frequencies are adjusted to the BPFs (Blade Passage Frequency) of the propellers. The TVAs are mounted to the frames and the force direction is normal to the skin. The sensors are condenser microphones which are attached to the primary structure of the airframe. The TVAs are equipped with signal processing devices. These components carry out Fourier transforms and signal amplification for the sensor data and actuator signals. The communication between the TVAs and the central control unit is implemented by the CAN Bus protocol and mainly consists of complex coefficients for the sensor and actuator data. This paper describes the basic structure of the system, the hardware set-up and function tests of the controller.

Summary of Altitude Pulse Testing of a 100-lbf L02/LCH4 Reaction Control Engine

NASA Technical Reports Server (NTRS)

Marshall, William M.; Kleinhenz, Julie E.

2011-01-01

Recently, liquid oxygen-liquid methane (LO2/LCH4) has been considered as a potential "green" propellant alternative for future exploration missions. The Propulsion and Cryogenic Advanced Development (PCAD) project has been tasked by NASA to develop this propulsion combination to enable safe and cost effective exploration missions. To date, limited experience with such combinations exist, and as a result a comprehensive test program is critical to demonstrating the viability of implementing such a system. The NASA Glenn Research Center has conducted a test program of a 100-lbf (445-N) reaction control engine (RCE) at the center s Altitude Combustion Stand (ACS), focusing on altitude testing over a wide variety of operational conditions. The ACS facility includes a unique propellant conditioning feed system (PCFS) which allows precise control of propellant inlet conditions to the engine. Engine performance as a result of these inlet conditions was examined extensively during the test program. This paper is a companion to the previous specific impulse testing paper, and discusses the pulsed mode operation portion of testing, with a focus on minimum impulse bit (I-bit) and repeatable pulse performance. The engine successfully demonstrated target minimum impulse bit performance at all conditions, as well as successful demonstration of repeatable pulse widths. Some anomalous conditions experienced during testing are also discussed, including a double pulse phenomenon which was not noted in previous test programs for this engine.

Catalytic ignitor for regenerative propellant gun

NASA Technical Reports Server (NTRS)

Voecks, Gerald E. (Inventor); Ferraro, Ned W. (Inventor)

1994-01-01

An ignitor initiates combustion of liquid propellant in a gun by utilizing a heated catalyst onto which the liquid propellant is sprayed in a manner which mitigates the occurrence of undesirable combustion chamber oscillations. The heater heats the catalyst sufficiently to provide the activation necessary to initiate combustion of the liquid propellant sprayed thereonto. Two embodiments of the ignitor and three alternative mountings thereof within the combustion chamber are disclosed. The ignitor may also be utilized to dispose of contaminated, excess, or waste liquid propellant in a safe, controlled, simple, and reliable manner.

Catalytic Ignitor for Regenerative Propellant Gun

NASA Technical Reports Server (NTRS)

Voecks, Gerald E. (Inventor); Ferraro, Ned W. (Inventor)

1997-01-01

An ignitor initiates combustion of liquid propellant in a gun by utilizing a heated catalyst onto which the liquid propellant is sprayed in a manner which mitigates the occurrence of undesirable combustion chamber oscillations. The heater heats the catalyst sufficiently to provide the activation necessary to initiate combustion of the liquid propellant sprayed thereonto. Two embodiments of the igniter and three alternative mountings thereof within the combustion chamber are disclosed. The ignitor may also be utilized to dispose of contaminated, excess, or waste liquid propellant in a safe, controlled, simple, and reliable manner.

Solid rocket propellant waste disposal/ingredient recovery study

NASA Technical Reports Server (NTRS)

Mcintosh, M. J.

1976-01-01

A comparison of facility and operating costs of alternate methods shows open burning to be the lowest cost incineration method of waste propellant disposal. The selection, development, and implementation of an acceptable alternate is recommended. The recovery of ingredients from waste propellant has the probability of being able to pay its way, and even show a profit, when large consistent quantities of composite propellant are available. Ingredients recovered from space shuttle waste propellant would be worth over $1.5 million. Open and controlled burning are both energy wasteful.

LH2 tank pressure control by thermodynamic vent system (TVS) at zero gravity

NASA Astrophysics Data System (ADS)

Wang, B.; Huang, Y. H.; Chen, Z. C.; Wu, J. Y.; Li, P.; Sun, P. J.

2017-02-01

Thermodynamic vent system (TVS) is employed for pressure control of propellant tanks at zero gravity. An analytical lumped parameter model is developed to predict pressure variation in an 18.09 m3 liquid hydrogen tank equipped with TVS. Mathematical simulations are carried out assuming tank is filled up to 75% volume (liquid mass equals to 945 kg) and is subjected to heat flux of 0.76 W/m2. Tank pressure controls at 165.5-172.4, 165.5-179.3 and 165.5-182.2 kPa are compared with reference to number of vent cycles, vent duration per cycle and loss of hydrogen. Analysis results indicate that the number of vent cycles significantly decreases from 62 to 21 when tank pressure control increases from 6.9 to 20.4 kPa. Also, duration of vent cycle increases from 63 to 152 and cycle duration decreases from 3920 to 3200 s. Further, the analysis result suggests that LH2 evaporation loss per day decreases from 0.17 to 0.14%. Based on the results of analysis, TVS is found effective in controlling the propellant tank pressure in zero gravity.

Lessons Learned Entry: Hypergolic Propellant Related Spills and Fires

NASA Technical Reports Server (NTRS)

Nufer, Brian

2009-01-01

The attached report is a compilation of all credible, unintentional hypergolic fluid related spills, fires, and explosions from the Apollo Program, the Space Shuttle Program, Titan Program, and a few other programs. Spill sites include the following government facilities: KSC, JSC, WSTF, VAFB, CCAFS, EAFB, Little Rock AFB, and McConnell AFB. The root causes and consequences of the incidents contained in this document vary drastically; however, certain "themes" can be deduced and utilized for future hypergolic propellant handling. Some of those common "themes" are summarized below: (1) Improper configuration control and complacency can lead to being falsely comfortable with a system (2) Communication breakdown can escalate an incident to a level where injuries occur and/or hardware is damaged (3) Improper propulsion system and ground support system designs can destine a system for failure (4) Improper training of technicians, engineers, and safety personnel can put lives in danger (5) Improper PPE, spill protection, and staging of fire extinguishing equipment can result in unnecessary injuries or hardware damage if an incident occurs (6) Improper procedural oversight, development, and adherence to the procedure can be detrimental and quickly lead to an undesirable incident (7) Improper local cleanliness or compatibility can result in fires or explosions The items listed above are only a short list of the issues that should be recognized prior to handling of hypergolic fluids or processing of vehicles containing hypergolic propellants. The summary of incidents in this report is intended to cover many more issues than those listed above that have been found during nearly the entire spectrum. of hypergolic propellant and/or vehicle processing.

Bismuth Propellant Feed System Development at NASA-MSFC

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.

2007-01-01

NASA-MSFC has been developing liquid metal propellant feed systems capable of delivering molten bismuth at a prescribed mass flow rate to the vaporizer of an electric thruster. The first such system was delivered to NASA-JPL as part of the Very High Isp Thruster with Anode Layer (VHITAL) program. In this system, the components pictured were placed in a vacuum chamber and heated while the control electronics were located outside the chamber. The system was successfully operated at JPL in conjunction with a propellant vaporizer, and data was obtained demonstrating a new liquid bismuth flow sensing technique developed at MSFC. The present effort is aimed at producing a feed-system for use in conjunction with a bismuth-fed Hall thruster developed by Busek Co. Developing this system is more ambitious, however, in that it is designed to self-contain all the control electronics inside the same vacuum chamber as an operating bismuth-fed thruster. Consequently, the entire system, including an on-board computer, DC-output power supplies, and a gas-pressurization electro-pneumatic regulator, must be designed to survive a vacuum environment and shielded to keep bismuth plasma from intruding on the electronics and causing a shortcircuit. In addition, the hot portions of the feed system must be thermally isolated from the electronics to avoid failure due to high heat loads. This is accomplished using a thermal protection system (TPS) consisting of multiple layers of aluminum foil. The only penetrations into the vacuum chamber are an electrically isolated (floating) 48 VDC line and a fiberoptic line. The 48 VDC provides power for operation of the power supplies and electronics co-located with the system in the vacuum chamber. The fiberoptic Ethernet connection is used to communicate user-input control commands to the on-board computer and transmit real-time data back to the external computer. The partially assembled second-generation system is shown. Before testing at Busek, a more detailed flow sensor calibration will be performed to accurately quantify the flow monitoring capabilities. This effort is funded under a Technology Innovation Program (TIP) award from NASA-MSFC's Technology Transfer office and performed under SAA8-061060.

A computer simulation of Skylab dynamics and attitude control for performance verification and operational support

NASA Technical Reports Server (NTRS)

Buchanan, H.; Nixon, D.; Joyce, R.

1974-01-01

A simulation of the Skylab attitude and pointing control system (APCS) is outlined and discussed. Implementation is via a large hybrid computer and includes those factors affecting system momentum management, propellant consumption, and overall vehicle performance. The important features of the flight system are discussed; the mathematical models necessary for this treatment are outlined; and the decisions involved in implementation are discussed. A brief summary of the goals and capabilities of this tool is also included.

14 CFR 25.149 - Minimum control speed.

Code of Federal Regulations, 2014 CFR

2014-01-01

... (7) If applicable, the propeller of the inoperative engine— (i) Windmilling; (ii) In the most..., the propeller of the inoperative engine in the position it achieves without pilot action, assuming the... propeller of the more critical inoperative engine in the position it achieves without pilot action, assuming...

14 CFR 25.149 - Minimum control speed.

Code of Federal Regulations, 2013 CFR

2013-01-01

... (7) If applicable, the propeller of the inoperative engine— (i) Windmilling; (ii) In the most..., the propeller of the inoperative engine in the position it achieves without pilot action, assuming the... propeller of the more critical inoperative engine in the position it achieves without pilot action, assuming...

14 CFR 25.149 - Minimum control speed.

Code of Federal Regulations, 2012 CFR

2012-01-01

... (7) If applicable, the propeller of the inoperative engine— (i) Windmilling; (ii) In the most..., the propeller of the inoperative engine in the position it achieves without pilot action, assuming the... propeller of the more critical inoperative engine in the position it achieves without pilot action, assuming...

The Potential of Aluminium Metal Powder as a Fuel for Space Propulsion Systems

NASA Astrophysics Data System (ADS)

Ismail, A. M.; Osborne, B.; Welch, C. S.

Metal powder propulsion systems have been addressed intermittently since the Second World War, initially in the field of underwater propulsion where research in the application of propelling torpedoes continues until this day. During the post war era, researchers attempted to utilise metal powders as a fuel for ram jet applications in missiles. The 1960's and 1970's saw additional interest in the use of `pure powder' propellants, i.e. fluidised metal fuel and oxidiser, both in solid particulate form. Again the application was for employment in space-constrained missiles where the idea was to maximise the performance of high energy density powder propellants in order to enhance the missile's flight duration. Metal powder as possible fuel was investigated for in-situ resource utilisation propulsion systems post-1980's where the emphasis was on the use of gaseous oxygen or liquid oxygen combined with aluminium metal powder for use as a ``lunar soil propellant'' or carbon dioxide and magnesium metal powder as a ``Martian propellant''.Albeit aluminium metal powder propellants are lower in performance than cryogenic and Earth storable propellants, the former does have an advantage inasmuch that the propulsion system is generic, i.e. it can be powered with chemicals mined and processed on Earth, the Moon and Mars. Thus, due to the potential refuelling capability, the lower performing aluminium metal powder propellant would effectively possess a much higher change in velocity (V) for multiple missions than the cryogenic or Earth storable propellant which is only suitable for one planet or one mission scenario, respectively.One of the principal limitations of long duration human spaceflight beyond cis-lunar orbit is the lack of refuelling capabilities on distant planets resulting in the reliance on con- ventional non-cryogenic, propellants produced on Earth. If one could develop a reliable propulsion system operating on pro- pellants derived entirely of ingredients found on nearby plan- etary bodies, then not only could mission duration be extended, larger amounts of payload could be ferried to and from the destination and eventually the cost of transporting propellant ingredients from Earth could be reduced, if not eliminated.

14 CFR Appendix A to Part 43 - Major Alterations, Major Repairs, and Preventive Maintenance

Code of Federal Regulations, 2013 CFR

2013-01-01

... hub design. (iii) Changes in the governor or control design. (iv) Installation of a propeller governor.... (iv) Engine mounts. (v) Control system. (vi) Landing gear. (vii) Hull or floats. (viii) Elements of an... of gravity limits of the aircraft. (xii) Changes to the basic design of the fuel, oil, cooling...

Application of Semi Active Control Techniques to the Damping Suppression Problem of Solar Sail Booms

NASA Technical Reports Server (NTRS)

Adetona, O.; Keel, L. H.; Whorton, M. S.

2007-01-01

Solar sails provide a propellant free form for space propulsion. These are large flat surfaces that generate thrust when they are impacted by light. When attached to a space vehicle, the thrust generated can propel the space vehicle to great distances at significant speeds. For optimal performance the sail must be kept from excessive vibration. Active control techniques can provide the best performance. However, they require an external power-source that may create significant parasitic mass to the solar sail. However, solar sails require low mass for optimal performance. Secondly, active control techniques typically require a good system model to ensure stability and performance. However, the accuracy of solar sail models validated on earth for a space environment is questionable. An alternative approach is passive vibration techniques. These do not require an external power supply, and do not destabilize the system. A third alternative is referred to as semi-active control. This approach tries to get the best of both active and passive control, while avoiding their pitfalls. In semi-active control, an active control law is designed for the system, and passive control techniques are used to implement it. As a result, no external power supply is needed so the system is not destabilize-able. Though it typically underperforms active control techniques, it has been shown to out-perform passive control approaches and can be unobtrusively installed on a solar sail boom. Motivated by this, the objective of this research is to study the suitability of a Piezoelectric (PZT) patch actuator/sensor based semi-active control system for the vibration suppression problem of solar sail booms. Accordingly, we develop a suitable mathematical and computer model for such studies and demonstrate the capabilities of the proposed approach with computer simulations.

Pulsed hydrojet

DOEpatents

Bohachevsky, I.O.; Torrey, M.D.

1986-06-10

An underwater pulsed hydrojet propulsion system is provided for accelerating and propelling a projectile or other vessel. A reactant, such as lithium, is fluidized and injected into a water volume. The resulting reaction produces an energy density in a time effective to form a steam pocket. Thrust flaps or baffles direct the pressure from the steam pocket toward an exit nozzle for accelerating a water volume to create thrust. A control system regulates the dispersion of reactant to control thrust characteristics.

Mesoscale pattern formation of self-propelled rods with velocity reversal

NASA Astrophysics Data System (ADS)

Großmann, Robert; Peruani, Fernando; Bär, Markus

2016-11-01

We study self-propelled particles with velocity reversal interacting by uniaxial (nematic) alignment within a coarse-grained hydrodynamic theory. Combining analytical and numerical continuation techniques, we show that the physics of this active system is essentially controlled by the reversal frequency. In particular, we find that elongated, high-density, ordered patterns, called bands, emerge via subcritical bifurcations from spatially homogeneous states. Our analysis reveals further that the interaction of bands is weakly attractive and, consequently, bands fuse upon collision in analogy with nonequilibrium nucleation processes. Moreover, we demonstrate that a renormalized positive line tension can be assigned to stable bands below a critical reversal rate, beyond which they are transversally unstable. In addition, we discuss the kinetic roughening of bands as well as their nonlinear dynamics close to the threshold of transversal instability. Altogether, the reduction of the multiparticle system onto the dynamics of bands provides a unified framework to understand the emergence and stability of nonequilibrium patterns in this self-propelled particle system. In this regard, our results constitute a proof of principle in favor of the hypothesis in microbiology that velocity reversal of gliding rod-shaped bacteria regulates the transitions between various self-organized patterns observed during the bacterial life cycle.

Hydrodynamic investigation of a self-propelled robotic fish based on a force-feedback control method.

PubMed

Wen, L; Wang, T M; Wu, G H; Liang, J H

2012-09-01

We implement a mackerel (Scomber scombrus) body-shaped robot, programmed to display the three most typical body/caudal fin undulatory kinematics (i.e. anguilliform, carangiform and thunniform), in order to biomimetically investigate hydrodynamic issues not easily tackled experimentally with live fish. The robotic mackerel, mounted on a servo towing system and initially at rest, can determine its self-propelled speed by measuring the external force acting upon it and allowing for the simultaneous measurement of power, flow field and self-propelled speed. Experimental results showed that the robotic swimmer with thunniform kinematics achieved a faster final swimming speed (St = 0.424) relative to those with carangiform (St = 0.43) and anguilliform kinematics (St = 0.55). The thrust efficiency, estimated from a digital particle image velocimetry (DPIV) flow field, showed that the robotic swimmer with thunniform kinematics is more efficient (47.3%) than those with carangiform (31.4%) and anguilliform kinematics (26.6%). Furthermore, the DPIV measurements illustrate that the large-scale characteristics of the flow pattern generated by the robotic swimmer with both anguilliform and carangiform kinematics were wedge-like, double-row wake structures. Additionally, a typical single-row reverse Karman vortex was produced by the robotic swimmer using thunniform kinematics. Finally, we discuss this novel force-feedback-controlled experimental method, and review the relative self-propelled hydrodynamic results of the robot when utilizing the three types of undulatory kinematics.

Development of unified propulsion system for geostationary satellite

NASA Astrophysics Data System (ADS)

Murayama, S.; Kobayashi, H.; Masuda, I.; Kameishi, M.; Miyoshi, K.; Takahashi, M.

Japan's first Liquid Apogee Propulsion System (LAPS) has been developed for ETS-VI (Engineering Test Satellite - VI) 2-ton class geostationary satellite. The next largest (2-ton class) geostationary satellite, COMETS (Communication and Broadcasting Engineering Test Satellite), requires a more compact apogee propulsion system in order to increase the space for mission instruments. The study for such a propulsion system concluded with a Unified Propulsion System (UPS), which uses a common N2H4 propellant tank for both bipropellant apogee engines and monopropellant Reaction Control System (RCS) thrusters. This type of propulsion system has several significant advantages compared with popular nitrogen tetroxide/monomethyl hydrazine (NTO/MMH) bipropellant satellite propulsion systems: The NTO/N2H4 apogee engine has a high specific impulse, and N2H4 thrusters have high reliability. Residual of N2H4 caused by propellant utilization of apogee engine firing (AEF) can be consumed by N2H4 monopropellant thrusters; that means a considerably prolonged satellite life.

Numerical Modeling of an Integrated Vehicle Fluids System Loop for Pressurizing a Cryogenic Tank

NASA Technical Reports Server (NTRS)

LeClair, A. C.; Hedayat, A.; Majumdar, A. K.

2017-01-01

This paper presents a numerical model of the pressurization loop of the Integrated Vehicle Fluids (IVF) system using the Generalized Fluid System Simulation Program (GFSSP). The IVF propulsion system, being developed by United Launch Alliance to reduce system weight and enhance reliability, uses boiloff propellants to drive thrusters for the reaction control system as well as to run internal combustion engines to develop power and drive compressors to pressurize propellant tanks. NASA Marshall Space Flight Center (MSFC) conducted tests to verify the functioning of the IVF system using a flight-like tank. GFSSP, a finite volume based flow network analysis software developed at MSFC, has been used to support the test program. This paper presents the simulation of three different test series, comparison of numerical prediction and test data and a novel method of presenting data in a dimensionless form. The paper also presents a methodology of implementing a compressor map in a system level code.

Generalized Advanced Propeller Analysis System (GAPAS). Volume 2: Computer program user manual

NASA Technical Reports Server (NTRS)

Glatt, L.; Crawford, D. R.; Kosmatka, J. B.; Swigart, R. J.; Wong, E. W.

1986-01-01

The Generalized Advanced Propeller Analysis System (GAPAS) computer code is described. GAPAS was developed to analyze advanced technology multi-bladed propellers which operate on aircraft with speeds up to Mach 0.8 and altitudes up to 40,000 feet. GAPAS includes technology for analyzing aerodynamic, structural, and acoustic performance of propellers. The computer code was developed for the CDC 7600 computer and is currently available for industrial use on the NASA Langley computer. A description of all the analytical models incorporated in GAPAS is included. Sample calculations are also described as well as users requirements for modifying the analysis system. Computer system core requirements and running times are also discussed.

Microelectrospray Thrusters

NASA Technical Reports Server (NTRS)

Dankanich, John; Demmons, Nate; Marrese-Reading, Colleen; Lozano, Paulo

2015-01-01

Propulsion technology is often a critical enabling technology for space missions. NASA is investing in technologies to enable high value missions with very small spacecraft, even CubeSats. However, these nanosatellites currently lack any appreciable propulsion capability. CubeSats are typically deployed and tumble or drift without any ability to transfer to higher value orbits, perform orbit maintenance, or perform de-orbit. Larger spacecraft can also benefit from high precision attitude control systems. Existing practices include reaction wheels with lifetime concerns and system level complexity. Microelectrospray thrusters will provide new propulsion capabilities to address these mission needs. Electric propulsion is an approach to accelerate propellant to very high exhaust velocities through the use of electrical power. Typical propulsion systems are limited to the combustion energy available in the chemical bonds of the fuel and then acceleration through a converging diverging nozzle. However, electric propulsion can accelerate propellant to ten times higher velocities and therefore increase momentum transfer efficiency, or essentially, increase the fuel economy. Fuel efficiency of thrusters is proportional to the exhaust velocity and referred to as specific impulse (Isp). The state-of-the-art (SOA) for CubeSats is cold gas propulsion with an Isp of 50-80 s. The Space Shuttle main engine demonstrated a specific impulse of 450 s. The target Isp for the Mars Exploration Program (MEP) systems is >1,500 s. This propellant efficiency can enable a 1-kg, 10-cm cube to transfer from low-Earth orbit to interplanetary space with only 200 g of propellant. In September 2013, NASA's Game Changing Development program competitively awarded three teams with contracts to develop MEP systems from Technology Readiness Level-3 (TRL-3), experimental concept, to TRL-5, system validation in a relevant environment. The project is planned for 18 months of system development. Due to the ambitious project goals, NASA has awarded contracts to mature three unique methods to achieve the desired goals. Some of the MEP concepts have been developed for more than a decade at the component level, but are now ready for system maturation. The three concepts include the high aspect ratio porous surface (HARPS) microthruster system, the scalable ion electrospray propulsion system (S-iEPS), and an indium microfluidic electrospray propulsion system. The HARPS system is under development by Busek Co. The HARPS thruster is an electrospray thruster that relies on surface emission of a porous metal with a passive capillary wicking system for propellant management. The HARPS thruster is expected to provide a simple, high ?V and low-cost solution. The HARPS thruster concept is shown in figure 1. Figure 1 includes the thruster, integrated power processing unit, and propellant reservoir.

Extrusion foaming of thermoplastic cellulose acetate from renewable resources using a two-component physical blowing agent system

NASA Astrophysics Data System (ADS)

Hopmann, Ch.; Windeck, C.; Hendriks, S.; Zepnik, S.; Wodke, T.

2014-05-01

Thermoplastic cellulose acetate (CA) is a bio-based polymer with optical, mechanical and thermal properties comparable to those of polystyrene (PS). The substitution of the predominant petrol-based PS in applications like foamed food trays can lead to a more sustainable economic practice. However, CA is also suitable for more durable applications as the biodegradability rate can be controlled by adjusting the degree of substitutions. The extrusion foaming of CA still has to overcome certain challenges. CA is highly hydrophilic and can suffer from hydrolytic degradation if not dried properly. Therefore, the influence of residual moisture on the melt viscosity is rather high. Beyond, the surface quality of foam CA sheets is below those of PS due to the particular foaming behaviour. This paper presents results of a recent study on extrusion foamed CA, using a two-component physical blowing agent system compromising HFO 1234ze as blowing agent and organic solvents as co-propellant. Samples with different co-propellants are processed on a laboratory single screw extruder at IKV. Morphology and surface topography are investigated with respect to the blowing agent composition and the die pressure. In addition, relationships between foam density, foam morphology and the propellants are analysed. The choice of the co-propellant has a significant influence on melt-strength, foaming behaviour and the possible blow-up ratio of the sheet. Furthermore, a positive influence of the co-propellant on the surface quality can be observed. In addition, the focus is laid on the effect of external contact cooling of the foamed sheets after the die exit.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Fearn, D.G.

The UK-10 ion thruster system is based closely on the 10 cm diameter T5 Kaufman-type thruster and its power conditioning and control system, which were developed during the 1970s for the North-South station-keeping application. The T5 device was designed to produce a thrust of 10 mN using mercury propellant. However, in the current work, mercury has been replaced by xenon to avoid any possibility of adverse chemical reactions with materials used in constructing spacecraft. In the previous phase of the program, it was shown that the system was fully suitable for its intended mission and that its integration into amore » spacecraft should present no difficulties. This paper reexamines that conclusion, bearing in mind the different physical characteristics of the new propellant. It is confirmed that the UK-10 system, using xenon, is compatible with the requirements of a wide range of applications. 31 references.« less

In-flight measurements of propeller blade deformation on a VUT100 cobra aeroplane using a co-rotating camera system

NASA Astrophysics Data System (ADS)

Boden, F.; Stasicki, B.; Szypuła, M.; Ružička, P.; Tvrdik, Z.; Ludwikowski, K.

2016-07-01

Knowledge of propeller or rotor blade behaviour under real operating conditions is crucial for optimizing the performance of a propeller or rotor system. A team of researchers, technicians and engineers from Avia Propeller, DLR, EVEKTOR and HARDsoft developed a rotating stereo camera system dedicated to in-flight blade deformation measurements. The whole system, co-rotating with the propeller at its full speed and hence exposed to high centrifugal forces and strong vibration, had been successfully tested on an EVEKTOR VUT 100 COBRA aeroplane in Kunovice (CZ) within the project AIM2—advanced in-flight measurement techniques funded by the European Commission (contract no. 266107). This paper will describe the work, starting from drawing the first sketch of the system up to performing the successful flight test. Apart from a description of the measurement hardware and the applied IPCT method, the paper will give some impressions of the flight test activities and discuss the results obtained from the measurements.

Two-step rocket engine bipropellant valve concept

NASA Technical Reports Server (NTRS)

Capps, J. E.; Ferguson, R. E.; Pohl, H. O.

1969-01-01

Initiating combustion of altitude control rocket engines in a precombustion chamber of ductile material reduces high pressure surges generated by hypergolic propellants. Two-step bipropellant valve concepts control initial propellant flow into precombustion chamber and subsequent full flow into main chamber.

Coated oxidizers for combustion stability in solid-propellant rockets

NASA Technical Reports Server (NTRS)

Helmy, A. M.; Ramohalli, K. N. R.

1985-01-01

Experiments are conducted in a laboratory-scale (6.25-cm diameter) end-burning rocket motor with state-of-the-art, ammonium perchlorate hydroxy-terminated polybutadiene (HTPB), nonmetallized propellants. The concept of tailoring the stability characteristics with a small amount (less than 1 percent by weight) of COATING on the oxidizer is explored. The thermal degradation characteristics of the coat chemical are deduced through theoretical arguments on thermal diffusivity of the composite material (propellant). Several candidate coats are selected and propellants are cast. These propellants (with coated oxidizers) are fired in a laboratory-scale end-burning rocket motor, and real-time pressure histories are recorded. The control propellant (with no coating) is also tested for comparison. The uniformity of the coating, confirmed by SEM pictures and BET adsorption measurements, is thought to be an advance in technology. The frequency of bulk mode instability (BMI), the pressure fluctuation amplitudes, and stability boundaries are correlated with parameters related to the characteristic length (L-asterisk) of the rocket motor. The coated oxidizer propellants, in general, display greater combustion stability than the control (state-of-the-art). The correlations of the various parameters are thought to be new to a field filled with much uncertainty.

19 CFR 192.3 - Penalties.

Code of Federal Regulations, 2011 CFR

2011-04-01

...) EXPORT CONTROL Exportation of Used Self-Propelled Vehicles, Vessels, and Aircraft § 192.3 Penalties. (a... importer or exporter who knowingly imports, exports or attempts to import or export: (1) Any stolen self-propelled vehicle, vessel, aircraft or part of a self-propelled vehicle, vessel or aircraft; or (2) Any self...

19 CFR 192.3 - Penalties.

Code of Federal Regulations, 2010 CFR

2010-04-01

...) EXPORT CONTROL Exportation of Used Self-Propelled Vehicles, Vessels, and Aircraft § 192.3 Penalties. (a... importer or exporter who knowingly imports, exports or attempts to import or export: (1) Any stolen self-propelled vehicle, vessel, aircraft or part of a self-propelled vehicle, vessel or aircraft; or (2) Any self...

14 CFR 91.611 - Authorization for ferry flight with one engine inoperative.

Code of Federal Regulations, 2013 CFR

2013-01-01

... (with its propeller removed or in a configuration desired by the operator and with all other engines... with— (i) The landing gear extended; (ii) The critical engine inoperative and its propeller removed or... controlled satisfactorily with the critical engine inoperative (with its propeller removed or in a...

14 CFR 91.611 - Authorization for ferry flight with one engine inoperative.

Code of Federal Regulations, 2011 CFR

2011-01-01

... (with its propeller removed or in a configuration desired by the operator and with all other engines... with— (i) The landing gear extended; (ii) The critical engine inoperative and its propeller removed or... controlled satisfactorily with the critical engine inoperative (with its propeller removed or in a...

14 CFR 91.611 - Authorization for ferry flight with one engine inoperative.

Code of Federal Regulations, 2012 CFR

2012-01-01

... (with its propeller removed or in a configuration desired by the operator and with all other engines... with— (i) The landing gear extended; (ii) The critical engine inoperative and its propeller removed or... controlled satisfactorily with the critical engine inoperative (with its propeller removed or in a...

14 CFR 91.611 - Authorization for ferry flight with one engine inoperative.

Code of Federal Regulations, 2014 CFR

2014-01-01

... (with its propeller removed or in a configuration desired by the operator and with all other engines... with— (i) The landing gear extended; (ii) The critical engine inoperative and its propeller removed or... controlled satisfactorily with the critical engine inoperative (with its propeller removed or in a...

49 CFR 240.7 - Definitions.

Code of Federal Regulations, 2014 CFR

2014-10-01

... accordance with § 240.104(a)(2)): (1) With one or more propelling motors designed for moving other equipment; (2) With one or more propelling motors designed to carry freight or passenger traffic or both; or (3) Without propelling motors but with one or more control stands. Locomotive engineer means any person who...

49 CFR 240.7 - Definitions.

Code of Federal Regulations, 2013 CFR

2013-10-01

... accordance with § 240.104(a)(2)): (1) With one or more propelling motors designed for moving other equipment; (2) With one or more propelling motors designed to carry freight or passenger traffic or both; or (3) Without propelling motors but with one or more control stands. Locomotive engineer means any person who...

Development of a novel hydroxyl ammonium nitrate based liquid propellant for air-independent propulsion

NASA Astrophysics Data System (ADS)

Fontaine, Joseph Henry

The focus of this dissertation is the development of an Unmanned Undersea Vehicle (UUV) liquid propellant employing Hydroxyl Ammonium Nitrate (HAN) as the oxidizer. Hydroxyl Ammonium Nitrate is a highly acidic aqueous based liquid oxidizer. Therefore, in order to achieve efficient combustion of a propellant using this oxidizer, the fuel must be highly water soluble and compatible with the oxidizer to prevent a premature ignition prior to being heated within the combustion chamber. An extensive search of the fuel to be used with this oxidizer was conducted. Propylene glycol was chosen as the fuel for this propellant, and the propellant given the name RF-402. The propellant development process will first evaluate the propellants thermal stability and kinetic parameters using a Differential Scanning Calorimeter (DSC). The purpose of the thermal stability analysis is to determine the temperature at which the propellant decomposition begins for the future safe handling of the propellant and the optimization of the combustion chamber. Additionally, the thermogram results will provide information regarding any undesirable endotherms prior to the decomposition and whether or not the decomposition process is a multi-step process. The Arrhenius type kinetic parameters will be determined using the ASTM method for thermally unstable materials. The activation energy and pre-exponential factor of the propellant will be determined by evaluating the decomposition peak temperature over a temperature scan rate ranging from 1°C per minute to 10°C per minute. The kinetic parameters of the propellant will be compared to those of 81 wt% HAN to determine if the HAN decomposition is controlling the overall decomposition of the propellant RF-402. The lifetime of individual droplets will be analyzed using both experimental and theoretical techniques. The theoretical technique will involve modeling the lifetime of an individual droplet in a combustion chamber like operating environment. The experimental technique will consist of subjecting droplets suspended from a fine gauge thermocouple to an instantaneous hot gas source and recording its temperature response while imaging it using a high power video microscope to determine the physical response of the droplet. This analysis will be the foundation for all future efforts in developing a propulsion system employing the use of RF-402.

Development of a green bipropellant hydrogen peroxide thruster for attitude control on satellites

NASA Astrophysics Data System (ADS)

Woschnak, A.; Krejci, D.; Schiebl, M.; Scharlemann, C.

2013-03-01

This document describes the selection assessment of propellants for a 1-newton green bipropellant thruster for attitude control on satellites. The development of this thruster was conducted as a part of the project GRASP (Green Advanced Space Propellants) within the European FP7 research program. The green propellant combinations hydrogen peroxide (highly concentrated with 87.5 %(wt.)) with kerosene or hydrogen peroxide (87.5 %(wt.)) with ethanol were identified as interesting candidates and were investigated in detail with the help of an experimental combustion chamber in the chemical propulsion laboratory at the Forschungsund Technologietransfer GmbH ― Fotec. Based on the test results, a final selection of propellants was performed.

A knowledge based expert system for propellant system monitoring at the Kennedy Space Center

NASA Technical Reports Server (NTRS)

Jamieson, J. R.; Delaune, C.; Scarl, E.

1985-01-01

The Lox Expert System (LES) is the first attempt to build a realtime expert system capable of simulating the thought processes of NASA system engineers, with regard to fluids systems analysis and troubleshooting. An overview of the hardware and software describes the techniques used, and possible applications to other process control systems. LES is now in the advanced development stage, with a full implementation planned for late 1985.

Polar Flagellar Motility of the Vibrionaceae

PubMed Central

McCarter, Linda L.

2001-01-01

Polar flagella of Vibrio species can rotate at speeds as high as 100,000 rpm and effectively propel the bacteria in liquid as fast as 60 μm/s. The sodium motive force powers rotation of the filament, which acts as a propeller. The filament is complex, composed of multiple subunits, and sheathed by an extension of the cell outer membrane. The regulatory circuitry controlling expression of the polar flagellar genes of members of the Vibrionaceae is different from the peritrichous system of enteric bacteria or the polar system of Caulobacter crescentus. The scheme of gene control is also pertinent to other members of the gamma purple bacteria, in particular to Pseudomonas species. This review uses the framework of the polar flagellar system of Vibrio parahaemolyticus to provide a synthesis of what is known about polar motility systems of the Vibrionaceae. In addition to its propulsive role, the single polar flagellum of V. parahaemolyticus is believed to act as a tactile sensor controlling surface-induced gene expression. Under conditions that impede rotation of the polar flagellum, an alternate, lateral flagellar motility system is induced that enables movement through viscous environments and over surfaces. Although the dual flagellar systems possess no shared structural components and although distinct type III secretion systems direct the simultaneous placement and assembly of polar and lateral organelles, movement is coordinated by shared chemotaxis machinery. PMID:11528005

Optimal spacecraft formation establishment and reconfiguration propelled by the geomagnetic Lorentz force

NASA Astrophysics Data System (ADS)

Huang, Xu; Yan, Ye; Zhou, Yang

2014-12-01

The Lorentz force acting on an electrostatically charged spacecraft as it moves through the planetary magnetic field could be utilized as propellantless electromagnetic propulsion for orbital maneuvering, such as spacecraft formation establishment and formation reconfiguration. By assuming that the Earth's magnetic field could be modeled as a tilted dipole located at the center of Earth that corotates with Earth, a dynamical model that describes the relative orbital motion of Lorentz spacecraft is developed. Based on the proposed dynamical model, the energy-optimal open-loop trajectories of control inputs, namely, the required specific charges of Lorentz spacecraft, for Lorentz-propelled spacecraft formation establishment or reconfiguration problems with both fixed and free final conditions constraints are derived via Gauss pseudospectral method. The effect of the magnetic dipole tilt angle on the optimal control inputs and the relative transfer trajectories for formation establishment or reconfiguration is also investigated by comparisons with the results derived from a nontilted dipole model. Furthermore, a closed-loop integral sliding mode controller is designed to guarantee the trajectory tracking in the presence of external disturbances and modeling errors. The stability of the closed-loop system is proved by a Lyapunov-based approach. Numerical simulations are presented to verify the validity of the proposed open-loop control methods and demonstrate the performance of the closed-loop controller. Also, the results indicate the dipole tilt angle should be considered when designing control strategies for Lorentz-propelled spacecraft formation establishment or reconfiguration.

Green Propellant Infusion Mission Program Development and Technology Maturation

NASA Technical Reports Server (NTRS)

McLean, Christopher H.; Deininger, William D.; Joniatis, John; Aggarwal, Pravin K.; Spores, Ronald A.; Deans, Matthew; Yim, John T.; Bury, Kristen; Martinez, Jonathan; Cardiff, Eric H.;

Ignition and combustion characteristics of metallized propellants

NASA Technical Reports Server (NTRS)

Turns, S. R.; Mueller, D. C.; Scott, M. J.

1990-01-01

Research designed to develop detailed knowledge of the secondary atomization and ignition characteristics of aluminum slurry propellants was started. These processes are studied because they are the controlling factors limiting the combustion efficiency of aluminum slurry propellants in rocket applications. A burner and spray rig system allowing the study of individual slurry droplets having diameters from about 10 to 100 microns was designed and fabricated. The burner generates a near uniform high temperature environment from the merging of 72 small laminar diffusion flames above a honeycomb matrix. This design permits essentially adiabatic operation over a wide range of stoichiometries without danger of flashback. A single particle sizing system and velocimeter also were designed and assembled. Light scattered from a focused laser beam is related to the particle (droplet) size, while the particle velocity is determined by its transit time through the focal volume. Light from the combustion of aluminum is also sensed to determine if ignition was achieved. These size and velocity measurements will allow the determination of disruption and ignition times as functions of drop sizes and ambient conditions.

Propulsion

NASA Astrophysics Data System (ADS)

Smith, P. K.

1993-06-01

Current requirements for missile systems increasingly stress the need for stealth capability. For the majority of missile systems and missions, the exhaust plume is likely to be the major contributor to overall missile signature, especially considering the recent developments in low emission and low Radar Cross Section coatings for motor bodies. This implies the need for the lowest possible rocket exhaust signature over a wide range of frequencies from the UV through visible and IR to microwave and radio frequencies. The choice of propellant type, Double Base; Composite etc, plays a significant part in determining the exhaust signature of the rocket motor as does the selection of inert materials for liners, inhibitors, and nozzles. It is also possible with certain propellants to incorporate additives which reduce exhaust signature either by modifying the chemistry or the afterburning plume or more significantly by suppressing secondary combustion and hence dramatically reducing plume temperature. The feasibility of plume signature control on the various missions envisaged by the missile designer is considered. The choice of propellant type and hardware components to give low signature is discussed together with performance implications. Signature reduction results obtained over a wide range of frequencies are also presented.

Self-regulation in self-propelled nematic fluids.

PubMed

Baskaran, A; Marchetti, M C

2012-09-01

We consider the hydrodynamic theory of an active fluid of self-propelled particles with nematic aligning interactions. This class of materials has polar symmetry at the microscopic level, but forms macrostates of nematic symmetry. We highlight three key features of the dynamics. First, as in polar active fluids, the control parameter for the order-disorder transition, namely the density, is dynamically convected by the order parameter via active currents. The resulting dynamical self-regulation of the order parameter is a generic property of active fluids and destabilizes the uniform nematic state near the mean-field transition. Secondly, curvature-driven currents render the system unstable deep in the nematic state, as found previously. Finally, and unique to self-propelled nematics, nematic order induces local polar order that in turn leads to the growth of density fluctuations. We propose this as a possible mechanism for the smectic order of polar clusters seen in numerical simulations.

Cryogenic Fluid Management Technologies for Advanced Green Propulsion Systems

NASA Technical Reports Server (NTRS)

Motil, Susan M.; Meyer, Michael L.; Tucker, Stephen P.

2007-01-01

In support of the Exploration Vision for returning to the Moon and beyond, NASA and its partners are developing and testing critical cryogenic fluid propellant technologies that will meet the need for high performance propellants on long-term missions. Reliable knowledge of low-gravity cryogenic fluid management behavior is lacking and yet is critical in the areas of tank thermal and pressure control, fluid acquisition, mass gauging, and fluid transfer. Such knowledge can significantly reduce or even eliminate tank fluid boil-off losses for long term missions, reduce propellant launch mass and required on-orbit margins, and simplify vehicle operations. The Propulsion and Cryogenic Advanced Development (PCAD) Project is performing experimental and analytical evaluation of several areas within Cryogenic Fluid Management (CFM) to enable NASA's Exploration Vision. This paper discusses the status of the PCAD CFM technology focus areas relative to the anticipated CFM requirements to enable execution of the Vision for Space Exploration.

Stratospheric controlled perturbation experiment (SCoPEx): overview, status, and results from related laboratory experiments

NASA Astrophysics Data System (ADS)

Keith, D.; Dykema, J. A.; Keutsch, F. N.

2017-12-01

Stratospheric Controlled Perturbation Experiment (SCoPEx), is a scientific experiment to advance understanding of stratospheric aerosols. It aims to make quantitative measurements of aerosol microphysics and atmospheric chemistry to improve large-scale models used to assess the risks and benefits of solar geoengineering. A perturbative experiment requires: (a) means to create a well-mixed, small perturbed volume, and (b) observation of time evolution of chemistry and aerosols in the volume. SCoPEx will used a propelled balloon gondola containing all instruments and drive system. The propeller wake forms a well-mixed volume (roughly 1 km long and 100 meters in diameter) that serves as an experimental `beaker' into which aerosols (e.g., < 1 kg of 0.3 µm radius CaCO3 particles) at can be injected; while, the propellers allow the gondola to move at speeds up to 3 m/sec relative to the local air mass driving the gondola back forth through the volume to measure properties of the perturbed air mass. This presentation will provide an overview of the experiment including (a) a systems engineering perspective from high-level scientific questions through instrument selection, mission design, and proposed operations and data analysis; (b) instruments, include current status of integration testing; (c) payload engineering including structure, power and mass budget, etc; (d) results from CFD simulation of propeller wake and simulation of chemistry and aerosol microphysics; and finally (e) proposed concept of operations and schedule. We will also provide an overview of the plans for governance including management of health safety and environmental risks, transparency, public engagement, and larger questions about governance of solar geoengineering experiments. Finally, we will briefly present results of laboratory experiments of the interaction of chemical such as ClONO2 and HCl on particle surfaces relevant for stratospheric solar geoengineering.

Guide to Flow Measurement for Electric Propulsion Systems

NASA Technical Reports Server (NTRS)

Frieman, Jason D.; Walker, Mitchell L. R.; Snyder, Steve

2013-01-01

In electric propulsion (EP) systems, accurate measurement of the propellant mass flow rate of gas or liquid to the thruster and external cathode is a key input in the calculation of thruster efficiency and specific impulse. Although such measurements are often achieved with commercial mass flow controllers and meters integrated into propellant feed systems, the variability in potential propellant options and flow requirements amongst the spectrum of EP power regimes and devices complicates meter selection, integration, and operation. At the direction of the Committee on Standards for Electric Propulsion Testing, a guide was jointly developed by members of the electric propulsion community to establish a unified document that contains the working principles, methods of implementation and analysis, and calibration techniques and recommendations on the use of mass flow meters in laboratory and spacecraft electric propulsion systems. The guide is applicable to EP devices of all types and power levels ranging from microthrusters to high-power ion engines and Hall effect thrusters. The establishment of a community standard on mass flow metering will help ensure the selection of the proper meter for each application. It will also improve the quality of system performance estimates by providing comprehensive information on the physical phenomena and systematic errors that must be accounted for during the analysis of flow measurement data. This paper will outline the standard methods and recommended practices described in the guide titled "Flow Measurement for Electric Propulsion Systems."

Practical Methodology for the Inclusion of Nonlinear Slosh Damping in the Stability Analysis of Liquid-Propelled Space Vehicles

NASA Technical Reports Server (NTRS)

Ottander, John A.; Hall, Robert A.; Powers, J. F.

2018-01-01

A method is presented that allows for the prediction of the magnitude of limit cycles due to adverse control-slosh interaction in liquid propelled space vehicles using non-linear slosh damping. Such a method is an alternative to the industry practice of assuming linear damping and relying on: mechanical slosh baffles to achieve desired stability margins; accepting minimal slosh stability margins; or time domain non-linear analysis to accept time periods of poor stability. Sinusoidal input describing functional analysis is used to develop a relationship between the non-linear slosh damping and an equivalent linear damping at a given slosh amplitude. In addition, a more accurate analytical prediction of the danger zone for slosh mass locations in a vehicle under proportional and derivative attitude control is presented. This method is used in the control-slosh stability analysis of the NASA Space Launch System.

Preliminary design of an auxiliary power unit for the space shuttle: Component and system configuration screening analysis

NASA Technical Reports Server (NTRS)

Binsley, R. L.; Maddox, J. P.; Marcy, R. D.; Siegler, R. S.; Spies, R.

1971-01-01

The auxiliary power unit (APU) for the space shuttle is required to provide hydraulic and electrical power on board the booster and orbiter vehicles. Five systems and their associated components, which utilize hot gas turbines to supply horsepower at gearbox output pads, were studied. Hydrogen-oxygen and storable propellants were considered for the hot gas supply. All APU's were required to be self-contained with respect to dissipating internally generated heat. These five systems were evaluated relative to a consistent criteria. The system supplied with high pressure gaseous hydrogen and oxygen was recommended as the best approach. It included a two-stage pressure-compounded partial-admission turbine, a propellant conditioning system with recuperation, a control system, and a gearbox. The gearbox output used was 240 hp. At the close of the study a 400 hp level was considered more appropriate for meeting the prime shuttle vehicle needs, and an in-depth analysis of the system at the 400 hp output level was recommended.

Forward Air Controllers in the Vietnam War: Exemplars of Audacity, Innovation, and Irreverence

DTIC Science & Technology

2016-06-10

Controllers (FACs). They were bold men who linked air power to ground forces, supersonic jets to propellers, and bombs to targets. They faced their...were bold men who linked air power to ground forces, supersonic jets to propellers, and bombs to targets. They faced their enemy at low altitudes, in...They were bold men who linked air power to ground forces, supersonic jets to propellers, and bombs to targets. FACs were the cutting edge of the