Design of a lunar propellant processing facility. NASA/USRA advanced program

NASA Technical Reports Server (NTRS)

Batra, Rajesh; Bell, Jason; Campbell, J. Matt; Cash, Tom; Collins, John; Dailey, Brian; France, Angelique; Gareau, Will; Gleckler, Mark; Hamilton, Charles

1993-01-01

Mankind's exploration of space will eventually lead to the establishment of a permanent human presence on the Moon. Essential to the economic viability of such an undertaking will be prudent utilization of indigenous lunar resources. The design of a lunar propellant processing system is presented. The system elements include facilities for ore processing, ice transportation, water splitting, propellant storage, personnel and materials transportation, human habitation, power generation, and communications. The design scenario postulates that ice is present in the lunar polar regions, and that an initial lunar outpost was established. Mining, ore processing, and water transportation operations are located in the polar regions. Water processing and propellant storage facilities are positioned near the equator. A general description of design operations is outlined below. Regolith containing the ice is mined from permanently-shaded polar craters. Water is separated from the ore using a microwave processing technique, and refrozen into projectiles for launch to the equatorial site via railgun. A mass-catching device retrieves the ice. This ice is processed using fractional distillation to remove impurities, and the purified liquid water is fed to an electrolytic cell that splits the water into vaporous hydrogen and oxygen. The hydrogen and oxygen are condensed and stored separately in a tank farm. Electric power for all operations is supplied by SP-100 nuclear reactors. Transportation of materials and personnel is accomplished primarily using chemical rockets. Modular living habitats are used which provide flexibility for the placement and number of personnel. A communications system consisting of lunar surface terminals, a lunar relay satellite, and terrestrial surface stations provides capabilities for continuous Moon-Moon and Moon-Earth transmissions of voice, picture, and data.

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.

Measured noise of a scale model high speed propeller at simulated takeoff/approach conditions

NASA Technical Reports Server (NTRS)

Woodward, Richard P.

1987-01-01

A model high-speed advanced propeller, SR-7A, was tested in the NASA Lewis 9x15 foot anechoic wind tunnel at simulated takeoff/approach conditions of 0.2 Mach number. These tests were in support of the full-scale Propfan Text Assessment (PTA) flight program. Acoustic measurements were taken with fixed microphone arrays and with an axially translating microphone probe. Limited aerodynamic measurements were also taken to establish the propeller operating conditions. Tests were conducted with the propeller alone and with three down-stream wing configurations. The propeller was run over a range of blade setting angles from 32.0 deg. to 43.6 deg., tip speeds from 183 to 290 m/sec (600 to 950 ft/sec), and angles of attack from -10 deg. to +15 deg. The propeller alone BPF tone noise was found to increase 10 dB in the flyover plane at 15 deg. propeller axis angle of attack. The installation of the straight wing at minimum spacing of 0.54 wing chord increased the tone noise 5 dB under the wing of 10 deg. propeller axis angle of attack, while a similarly spaced inboard upswept wing only increased the tone noise 2 dB.

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.

Effect of angular inflow on the vibratory response of a counter-rotating propeller

NASA Technical Reports Server (NTRS)

Turnberg, J. E.; Brown, P. C.

1985-01-01

This report presents the results of a propeller vibratory stress survey on the Fairey Gannet aircraft aimed at giving an assessment of the difference in vibratory response between single and counter-rotating propeller operation in angular inflow. The survey showed that counter-rotating operation of the propeller had the effect of increasing the IP response of the rear propeller by approximately 25 percent over comparable single-rotation operation while counter-rotating operation did not significantly influence the IP response of the front propeller.

Developing a safe on-orbit cryogenic depot

NASA Technical Reports Server (NTRS)

Bahr, Nicholas J.

1992-01-01

New U.S. space initiatives will require technology to realize planned programs such as piloted lunar and Mars missions. Key to the optimal execution of such missions are high performance orbit transfer vehicles and propellant storage facilities. Large amounts of liquid hydrogen and oxygen demand a uniquely designed on-orbit cryogenic propellant depot. Because of the inherent dangers in propellant storage and handling, a comprehensive system safety program must be established. This paper shows how the myriad and complex hazards demonstrate the need for an integrated safety effort to be applied from program conception through operational use. Even though the cryogenic depot is still in the conceptual stage, many of the hazards have been identified, including fatigue due to heavy thermal loading from environmental and operating temperature extremes, micrometeoroid and/or depot ancillary equipment impact (this is an important problem due to the large surface area needed to house the large quantities of propellant), docking and maintenance hazards, and hazards associated with extended extravehicular activity. Various safety analysis techniques were presented for each program phase. Specific system safety implementation steps were also listed. Enhanced risk assessment was demonstrated through the incorporation of these methods.

14 CFR 417.417 - Propellants and explosives.

Code of Federal Regulations, 2012 CFR

2012-01-01

... radio frequency radiation sources in a radio frequency radiation exclusion area. A launch operator must determine the vulnerability of its electro-explosive devices and systems to radio frequency radiation and establish radio frequency radiation power limits or radio frequency radiation exclusion areas as required by...

14 CFR 417.417 - Propellants and explosives.

Code of Federal Regulations, 2013 CFR

2013-01-01

... radio frequency radiation sources in a radio frequency radiation exclusion area. A launch operator must determine the vulnerability of its electro-explosive devices and systems to radio frequency radiation and establish radio frequency radiation power limits or radio frequency radiation exclusion areas as required by...

14 CFR 417.417 - Propellants and explosives.

Code of Federal Regulations, 2011 CFR

2011-01-01

... radio frequency radiation sources in a radio frequency radiation exclusion area. A launch operator must determine the vulnerability of its electro-explosive devices and systems to radio frequency radiation and establish radio frequency radiation power limits or radio frequency radiation exclusion areas as required by...

14 CFR 417.417 - Propellants and explosives.

Code of Federal Regulations, 2014 CFR

2014-01-01

... radio frequency radiation sources in a radio frequency radiation exclusion area. A launch operator must determine the vulnerability of its electro-explosive devices and systems to radio frequency radiation and establish radio frequency radiation power limits or radio frequency radiation exclusion areas as required by...

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...

LOX droplet vaporization in a supercritical forced convective environment

NASA Technical Reports Server (NTRS)

Hsiao, Chia-Chun; Yang, Vigor

1993-01-01

Modern liquid rocket engines often use liquid oxygen (LOX) and liquid hydrogen (LH2) as propellants to achieve high performance, with the engine operational conditions in the supercritical regimes of the propellants. Once the propellant exceeds its critical state, it essentially becomes a puff of dense fluid. The entire field becomes a continuous medium, and no distinct interfacial boundary between the liquid and gas exists. Although several studies have been undertaken to investigate the supercritical droplet behavior at quiescent conditions, very little effort has been made to address the fundamental mechanisms associated with LOX droplet vaporization in a supercritical, forced convective environment. The purpose is to establish a theoretical framework within which supercritical droplet dynamics and vaporization can be studied systematically by means of an efficient and robust numerical algorithm.

Green Propellant Landing Demonstration at U.S. Range

NASA Technical Reports Server (NTRS)

Mulkey, Henry W.; Miller, Joseph T.; Bacha, Caitlin E.

2016-01-01

The Green Propellant Loading Demonstration (GPLD) was conducted December 2015 at Wallops Flight Facility (WFF), leveraging work performed over recent years to bring lower toxicity hydrazine replacement green propellants to flight missions. The objective of this collaboration between NASA Goddard Space Flight Center (GSFC), WFF, the Swedish National Space Board (SNSB), and Ecological Advanced Propulsion Systems (ECAPS) was to successfully accept LMP-103S propellant at a U.S. Range, store the propellant, and perform a simulated flight vehicle propellant loading. NASA GSFC Propulsion (Code 597) managed all aspects of the operation, handling logistics, preparing the procedures, and implementing the demonstration. In addition to the partnership described above, Moog Inc. developed an LMP-103S propellant-compatible titanium rolling diaphragm flight development tank and loaned it to GSFC to act as the GPLD flight vessel. The flight development tank offered the GPLD an additional level of flight-like propellant handling process and procedures. Moog Inc. also provided a compatible latching isolation valve for remote propellant expulsion. The GPLD operation, in concert with Moog Inc. executed a flight development tank expulsion efficiency performance test using LMP-103S propellant. As part of the demonstration work, GSFC and WFF documented Range safety analyses and practices including all elements of shipping, storage, handling, operations, decontamination, and disposal. LMP-103S has not been previously handled at a U.S. Launch Range. Requisite for this activity was an LMP-103S Risk Analysis Report and Ground Safety Plan. GSFC and WFF safety offices jointly developed safety documentation for application into the GPLD operation. The GPLD along with the GSFC Propulsion historical hydrazine loading experiences offer direct comparison between handling green propellant versus safety intensive, highly toxic hydrazine propellant. These described motives initiated the GPLD operation in order to investigate the handling and process safety variances in project resources between LMP-103S and typical in-space propellants. The GPLD risk reduction operation proved successful for many reasons including handling the green propellant at a U.S. Range, loading and pressurizing a flight-like tank, expelling the propellant, measuring the tank expulsion efficiency, and most significantly, GSFC propulsion personnel's new insight into the LMP-103S propellant handling details.

Green Propellant Loading Demonstration at U.S. Range

NASA Technical Reports Server (NTRS)

Mulkey, Henry W.; Miller, Joseph T.; Bacha, Caitlin E.

2016-01-01

The Green Propellant Loading Demonstration (GPLD) was conducted December 2015 at Wallops Flight Facility (WFF), leveraging work performed over recent years to bring lower toxicity hydrazine replacement green propellants to flight missions. The objective of this collaboration between NASA Goddard Space Flight Center (GSFC), WFF, the Swedish National Space Board (SNSB), and Ecological Advanced Propulsion Systems (ECAPS) was to successfully accept LMP-103S propellant at a U.S. Range, store the propellant, and perform a simulated flight vehicle propellant loading. NASA GSFC Propulsion (Code 597) managed all aspects of the operation, handling logistics, preparing the procedures, and implementing the demonstration. In addition to the partnership described above, Moog Inc. developed an LMP-103S propellant-compatible titanium rolling diaphragm flight development tank and loaned it to GSFC to act as the GPLD flight vessel. The flight development tank offered the GPLD an additional level of flight-like propellant handling process and procedures. Moog Inc. also provided a compatible latching isolation valve for remote propellant expulsion. The GPLD operation, in concert with Moog Inc. executed a flight development tank expulsion efficiency performance test using LMP-103S propellant. As part of the demonstration work, GSFC and WFF documented Range safety analyses and practices including all elements of shipping, storage, handling, operations, decontamination, and disposal. LMP-103S has not been previously handled at a U.S. Launch Range. Requisite for this activity was an LMP-103S Risk Analysis Report and Ground Safety Plan. GSFC and WFF safety offices jointly developed safety documentation for application into the GPLD operation. The GPLD along with the GSFC Propulsion historical hydrazine loading experiences offer direct comparison between handling green propellant versus safety intensive, highly toxic hydrazine propellant. These described motives initiated the GPLD operation in order to investigate the handling and process safety variances in project resources between LMP-103S and typical in-space propellants. The GPLD risk reduction operation proved successful for many reasons including handling the green propellant at a U.S. Range, loading and pressurizing a flight-like tank, expelling the propellant, measuring the tank expulsion efficiency, and most significantly, GSFC propulsion personnel's new insight into the LMP-103S propellant handling details.

Noise generated by a propeller in a wake

NASA Technical Reports Server (NTRS)

Block, P. J. W.

1984-01-01

Propeller performance and noise were measured on two model scale propellers operating in an anechoic flow environment with and without a wake. Wake thickness of one and three propeller chords were generated by an airfoil which spanned the full diameter of the propeller. Noise measurements were made in the relative near field of the propeller at three streamwise and three azimuthal positions. The data show that as much as 10 dB increase in the OASPL results when a wake is introduced into an operating propeller. Performance data are also presented for completeness.

14 CFR 35.16 - Propeller critical parts.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Propeller critical parts. 35.16 Section 35... AIRWORTHINESS STANDARDS: PROPELLERS Design and Construction § 35.16 Propeller critical parts. The integrity of each propeller critical part identified by the safety analysis required by § 35.15 must be established...

Lifetime Assessment of the NEXT Ion Thruster

NASA Technical Reports Server (NTRS)

VanNoord, Jonathan L.

2010-01-01

Ion thrusters are low thrust, high specific impulse devices with required operational lifetimes on the order of 10,000 to 100,000 hr. The NEXT ion thruster is the latest generation of ion thrusters under development. The NEXT ion thruster currently has a qualification level propellant throughput requirement of 450 kg of xenon, which corresponds to roughly 22,000 hr of operation at the highest throttling point. Currently, a NEXT engineering model ion thruster with prototype model ion optics is undergoing a long duration test to determine wear characteristics and establish propellant throughput capability. The NEXT thruster includes many improvements over previous generations of ion thrusters, but two of its component improvements have a larger effect on thruster lifetime. These include the ion optics with tighter tolerances, a masked region and better gap control, and the discharge cathode keeper material change to graphite. Data from the NEXT 2000 hr wear test, the NEXT long duration test, and further analysis is used to determine the expected lifetime of the NEXT ion thruster. This paper will review the predictions for all of the anticipated failure mechanisms. The mechanisms will include wear of the ion optics and cathode s orifice plate and keeper from the plasma, depletion of low work function material in each cathode s insert, and spalling of material in the discharge chamber leading to arcing. Based on the analysis of the NEXT ion thruster, the first failure mode for operation above a specific impulse of 2000 sec is expected to be the structural failure of the ion optics at 750 kg of propellant throughput, 1.7 times the qualification requirement. An assessment based on mission analyses for operation below a specific impulse of 2000 sec indicates that the NEXT thruster is capable of double the propellant throughput required by these missions.

Experimental study of combustion processes involved in hypergolic propellant coaxial injector operation

NASA Astrophysics Data System (ADS)

Habiballah, M.; Dubois, I.; Gicquel, P.; Foucaud, R.

1992-07-01

The first results are presented of an experimental research program to understand the operation of a coaxial injector using hypergolic propellants. Mechanisms and processes involved in coaxial injector operation are identified for a two-plate injector and a coaxial injector. The usefulness of backlight cinematography and laser sheet visualization in the study of coaxial injector operation is examined. A review of the literature on injector elements using highly reactive hypergolic propellants is presented along with an analysis of fundamental mechanisms involved in these propellants.

Analysis of noise measured from a propeller in a wake

NASA Technical Reports Server (NTRS)

Block, P. J. W.

1984-01-01

In this experimental study, the acoustic characteristics of a propeller operating in a wake were studied. The propeller performance and noise were measured from two 0.25 scale propellers operating in an open jet anechoic flow environment with and without a wake. One propeller had NACA 16 series sections; the other, ARA-D. Wake thicknesses of 1 and 3 propeller chords were generated by an airfoil which spanned the full diameter of the propeller. The airfoil wake profiles were measured. Noise measurements were made in and out of the flow. The propellers were operated at 40, 83, and 100 inf of thrust. The acoustic data are analyzed, and the effects on the overall sound pressure level (OASPL) and scaled A weighted sound level L sub A with propeller thrust, wake thickness, and observer location are presented. The analysis showed that, generally, the wake increased the overall noise (OASPL) produced by the propeller; increased the harmonic content of the noise, thus the scaled L sub a; and produced an azimuthal dependence. With few exceptions, both propellers generally produced the same trends in delta OASPL and delta L sub a with thrust and wake thickness.

Propellant Technologies: A Persuasive Wave of Future Propulsion Benefits

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan; Ianovski, Leonid S.; Carrick, Patrick

1997-01-01

Rocket propellant and propulsion technology improvements can be used to reduce the development time and operational costs of new space vehicle programs. Advanced propellant technologies can make the space vehicles safer, more operable, and higher performing. Five technology areas are described: Monopropellants, Alternative Hydrocarbons, Gelled Hydrogen, Metallized Gelled Propellants, and High Energy Density Materials. These propellants' benefits for future vehicles are outlined using mission study results and the technologies are briefly discussed.

Noise of a model counterrotation propeller with simulated fuselage and support pylon at takeoff/approach conditions

NASA Technical Reports Server (NTRS)

Woodward, Richard P.; Hughes, Christopher E.

1989-01-01

Two modern high-speed advanced counterrotation propellers, F7/A7 and F7/A3 were tested in the NASA Lewis Research Centers's 9- by 15-foot Anechoic Wind Tunnel at simulated takeoff/approach conditions of 0.2 Mach number. Both rotors were of similar diameter on the F7/A7 propeller, while the aft rotor diameter of the F7/A3 propeller was 85 percent of the forward propeller to reduce tip vortex-aft rotor interaction. The two propellers were designed for similar performance. The propellers were tested in both the clean configuration, and installed configuration consisting of a simulated upstream nacelle support pylon and fuselage section. Acoustic measurements were made with an axially translating microphone probe, and with a polar microphone probe which was fixed to the propeller nacelle and could make both sideline and circumferential acoustic surveys. Aerodynamic measurements were also made to establish propeller operating conditions. The propellers were run at blade setting angles (fron angle/rear angle) of 41.1/39.4 deg for the F7/A7 propeller, and 41.1/46.4 deg for the F7/A3 propeller. The forward rotors were tested over a range of tip speeds from 165 to 259 m/sec (540 to 850 ft/sec), and both propellers were tested at the maximum rotor-rotor spacing, based on pitch change axis separation, of 14.99 cm (5.90 in.). The data presented in this paper are for 0 deg propeller axis angle of attack. Results are presented for the baseline, pylon-alone, and strut + fuselage configurations. The presence of the simulated fuselage resulted in higher rotor-alone tone levels in a direction normal to the advancing propeller blade near the fuselage. A corresponding rotor-alone tone reduction was often observed 180 deg circumferentially from this region of increased noise. A significant rotor-alone increase for both rotors was observed diametrically opposite the fuselage. In some cases, interaction tone levels were likewise affected by the simulated installation.

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.

14 CFR 121.641 - Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Flag operations.

Code of Federal Regulations, 2014 CFR

2014-01-01

...-powered airplanes: Flag operations. 121.641 Section 121.641 Aeronautics and Space FEDERAL AVIATION... Flight Release Rules § 121.641 Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Flag operations. (a) No person may dispatch or take off a nonturbine or turbo-propeller-powered airplane unless...

14 CFR 121.641 - Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Flag operations.

Code of Federal Regulations, 2012 CFR

2012-01-01

...-powered airplanes: Flag operations. 121.641 Section 121.641 Aeronautics and Space FEDERAL AVIATION... Flight Release Rules § 121.641 Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Flag operations. (a) No person may dispatch or take off a nonturbine or turbo-propeller-powered airplane unless...

14 CFR 121.641 - Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Flag operations.

Code of Federal Regulations, 2013 CFR

2013-01-01

...-powered airplanes: Flag operations. 121.641 Section 121.641 Aeronautics and Space FEDERAL AVIATION... Flight Release Rules § 121.641 Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Flag operations. (a) No person may dispatch or take off a nonturbine or turbo-propeller-powered airplane unless...

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.

Shuttle Centaur engine cooldown evaluation and effects of expanded inlets on start transient

NASA Technical Reports Server (NTRS)

1987-01-01

As part of the integration of the RL10 engine into the Shuttle Centaur vehicle, a satisfactory method of conditioning the engine to operating temperatures had to be established. This procedure, known as cooldown, is different from the existing Atlas Centaur due to vehicle configuration and mission profile differenced. The program is described, and the results of a Shuttle Centaur cooldown program are reported. Mission peculiarities cause substantial variation in propellant inlet conditions between the substantiated Atlas Centaur and Shuttle Centaur with the Shuttle Centaur having much larger variation in conditions. A test program was conducted to demonstrate operation of the RL10 engine over the expanded inlet conditions. As a result of this program, the Shuttle Centaur requirements were proven satisfactory. Minor configuration changes incorporated as a result of this program provide substantial reduction in cooldown propellant consumption.

Space shuttle aps propellant thermal conditioner study

NASA Technical Reports Server (NTRS)

Fulton, D. L.

1973-01-01

An analytical and experimental effort was completed to evaluate a baffle type thermal conditioner for superheating O2 and H2 at supercritical pressures. The thermal conditioner consisted of a heat exchanger and an integral reactor (gas generator) operating on O2/H2 propellants. Primary emphasis was placed on the hydrogen conditioner with some effort on the oxygen conditioner and a study completed of alternate concepts for use in conditioning oxygen. A hydrogen conditioner was hot fire tested under a range of conditions to establish ignition, heat exchange and response parameters. A parallel technology task was completed to further evaluate the integral reactor and heat exchanger with the side mounted electrical spark igniter.

Design and analysis report for the RL10-2B breadboard low thrust engine

NASA Technical Reports Server (NTRS)

Brown, J. R.; Foust, R. R.; Galler, D. E.; Kanic, P. G.; Kmiec, T. D.; Limerick, C. D.; Peckham, R. J.; Swartwout, T.

1984-01-01

The breadboard low thrust RL10-2B engine is described. A summary of the analysis and design effort to define the multimode thrust concept applicable to the requirements for the upper stage vehicles is provided. Baseline requirements were established for operation of the RL10-2B engine under the following conditions: (1) tank head idle at low propellant tank pressures without vehicle propellant conditioning or settling thrust; (2) pumped idle at a ten percent thrust level for low G deployment and/or vehicle tank pressurization; and (3) full thrust (15,000 lb.). Several variations of the engine configuration were investigated and results of the analyses are included.

A Cis-Lunar Propellant Infrastructure for Flexible Path Exploration and Space Commerce

NASA Technical Reports Server (NTRS)

Oeftering, Richard C.

2012-01-01

This paper describes a space infrastructure concept that exploits lunar water for propellant production and delivers it to users in cis-lunar space. The goal is to provide responsive economical space transportation to destinations beyond low Earth orbit (LEO) and enable in-space commerce. This is a game changing concept that could fundamentally affect future space operations, provide greater access to space beyond LEO, and broaden participation in space exploration. The challenge is to minimize infrastructure development cost while achieving a low operational cost. This study discusses the evolutionary development of the infrastructure from a very modest robotic operation to one that is capable of supporting human operations. The cis-lunar infrastructure involves a mix of technologies including cryogenic propellant production, reusable lunar landers, propellant tankers, orbital transfer vehicles, aerobraking technologies, and electric propulsion. This cislunar propellant infrastructure replaces Earth-launched propellants for missions beyond LEO. It enables users to reach destinations with smaller launchers or effectively multiplies the user s existing payload capacity. Users can exploit the expanded capacity to launch logistics material that can then be traded with the infrastructure for propellants. This mutually beneficial trade between the cis-lunar infrastructure and propellant users forms the basis of in-space commerce.

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.

14 CFR 23.905 - Propellers.

Code of Federal Regulations, 2014 CFR

2014-01-01

... Propellers. (a) Each propeller must have a type certificate. (b) Engine power and propeller shaft rotational... tests, that the propeller is capable of continuous safe operation. (h) All engine cowling, access doors... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Propellers. 23.905 Section 23.905...

14 CFR 23.905 - Propellers.

Code of Federal Regulations, 2013 CFR

2013-01-01

... Propellers. (a) Each propeller must have a type certificate. (b) Engine power and propeller shaft rotational... tests, that the propeller is capable of continuous safe operation. (h) All engine cowling, access doors... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Propellers. 23.905 Section 23.905...

14 CFR 23.905 - Propellers.

Code of Federal Regulations, 2010 CFR

2010-01-01

... Propellers. (a) Each propeller must have a type certificate. (b) Engine power and propeller shaft rotational... tests, that the propeller is capable of continuous safe operation. (h) All engine cowling, access doors... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Propellers. 23.905 Section 23.905...

14 CFR 23.905 - Propellers.

Code of Federal Regulations, 2011 CFR

2011-01-01

... Propellers. (a) Each propeller must have a type certificate. (b) Engine power and propeller shaft rotational... tests, that the propeller is capable of continuous safe operation. (h) All engine cowling, access doors... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Propellers. 23.905 Section 23.905...

14 CFR 23.905 - Propellers.

Code of Federal Regulations, 2012 CFR

2012-01-01

... Propellers. (a) Each propeller must have a type certificate. (b) Engine power and propeller shaft rotational... tests, that the propeller is capable of continuous safe operation. (h) All engine cowling, access doors... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Propellers. 23.905 Section 23.905...

Theory for noise of propellers in angular inflow with parametric studies and experimental verification

NASA Technical Reports Server (NTRS)

Hanson, Donald B.; Parzych, David J.

1993-01-01

This report presents the derivation of a frequency domain theory and working equations for radiation of propeller harmonic noise in the presence of angular inflow. In applying the acoustic analogy, integration over the tangential coordinate of the source region is performed numerically, permitting the equations to be solved without approximation for any degree of angular inflow. Inflow angle is specified in terms of yaw, pitch, and roll angles of the aircraft. Since these can be arbitrarily large, the analysis applies with equal accuracy to propellers and helicopter rotors. For thickness and loading, the derivation is given in complete detail with working equations for near and far field. However, the quadrupole derivation has been carried only far enough to show feasibility of the numerical approach. Explicit formulas are presented for computation of source elements, evaluation of Green's functions, and location of observer points in various visual and retarded coordinate systems. The resulting computer program, called WOBBLE has been written in FORTRAN and follows the notation of this report very closely. The new theory is explored to establish the effects of varying inflow angle on axial and circumferential directivity. Also, parametric studies were performed to evaluate various phenomena outside the capabilities of earlier theories, such as an unsteady thickness effect. Validity of the theory was established by comparison with test data from conventional propellers and Prop Fans in flight and in wind tunnels under a variety of operating conditions and inflow angles.

Rapid Propellant Loading Approach Exploration

DTIC Science & Technology

2010-11-01

the impact upon ground operations of three configuration options. Ground operations management was addressed through a series of studies performed...and operations management system can enable safe rapid propellant loading operations with limited operator knowledge and involvement. A single

Post Accident Procedures for Chemicals and Propellants.

DTIC Science & Technology

1982-09-01

METHODS AND PROCEDURES ............ 4-1 4.1 Overview of Emergency Response Procedures " and Resources Available .......................... 4-1 L1 TABLE...7-1 7.1 Criteria forTwelve Critical Operations ........................ 7-1 7.1.1 On-Scene Methods for Identifying the Ingredients...Establishing A Protocol for Selecting the Hazards Mitigation and Cleanup Methods for Single Material Spills and Multiple Materials Mixing

14 CFR 35.7 - Features and characteristics.

Code of Federal Regulations, 2011 CFR

2011-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS General § 35.7 Features and characteristics. (a) The propeller may not have... propeller. The applicant must make changes to the design and conduct additional tests that the Administrator finds necessary to establish the airworthiness of the propeller. [Amdt. No. 35-8, 73 FR 63346, Oct. 24...

14 CFR 35.7 - Features and characteristics.

Code of Federal Regulations, 2012 CFR

2012-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS General § 35.7 Features and characteristics. (a) The propeller may not have... propeller. The applicant must make changes to the design and conduct additional tests that the Administrator finds necessary to establish the airworthiness of the propeller. [Amdt. 35-8, 73 FR 63346, Oct. 24, 2008] ...

14 CFR 35.7 - Features and characteristics.

Code of Federal Regulations, 2013 CFR

2013-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS General § 35.7 Features and characteristics. (a) The propeller may not have... propeller. The applicant must make changes to the design and conduct additional tests that the Administrator finds necessary to establish the airworthiness of the propeller. [Amdt. 35-8, 73 FR 63346, Oct. 24, 2008] ...

14 CFR 35.7 - Features and characteristics.

Code of Federal Regulations, 2010 CFR

2010-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS General § 35.7 Features and characteristics. (a) The propeller may not have... propeller. The applicant must make changes to the design and conduct additional tests that the Administrator finds necessary to establish the airworthiness of the propeller. [Amdt. No. 35-8, 73 FR 63346, Oct. 24...

14 CFR 35.7 - Features and characteristics.

Code of Federal Regulations, 2014 CFR

2014-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS General § 35.7 Features and characteristics. (a) The propeller may not have... propeller. The applicant must make changes to the design and conduct additional tests that the Administrator finds necessary to establish the airworthiness of the propeller. [Amdt. 35-8, 73 FR 63346, Oct. 24, 2008] ...

Real-Time Inhibitor Recession Measurements in the Space Shuttle Reusable Solid Rocket Motors

NASA Technical Reports Server (NTRS)

McWhorter, Bruce B.; Ewing, Mark E.; McCool, Alex (Technical Monitor)

2001-01-01

Real-time char line recession measurements were made on propellant inhibitors of the Space Shuttle Reusable Solid Rocket Motor (RSRM). The RSRM FSM-8 static test motor propellant inhibitors (composed of a rubber insulation material) were successfully instrumented with eroding potentiometers and thermocouples. The data was used to establish inhibitor recession versus time relationships. Normally, pre-fire and post-fire insulation thickness measurements establish the thermal performance of an ablating insulation material. However, post-fire inhibitor decomposition and recession measurements are complicated by the fact that most of the inhibitor is back during motor operation. It is therefore a difficult task to evaluate the thermal protection offered by the inhibitor material. Real-time measurements would help this task. The instrumentation program for this static test motor marks the first time that real-time inhibitors. This report presents that data for the center and aft field joint forward facing inhibitors. The data was primarily used to measure char line recession of the forward face of the inhibitors which provides inhibitor thickness reduction versus time data. The data was also used to estimate the inhibitor height versus time relationship during motor operation.

The PROPEL Electrodynamic Tether Mission and Connecting to the Ionosphere

NASA Technical Reports Server (NTRS)

Gilchrist, Brian; Bilen, Sven; Hoyt, Rob; Stone,Nobie; Vaughn, Jason; Fuhrhop, Keith; Krause, Linda; Khazanov, George; Johnson, Les

2012-01-01

The exponential increase of launch system size.and cost.with delta-V makes missions that require large total impulse cost prohibitive. Led by NASA's Marshall Space Flight Center, a team from government, industry, and academia has developed a flight demonstration mission concept of an integrated electrodynamic (ED) tethered satellite system called PROPEL: "Propulsion using Electrodynamics". The PROPEL Mission is focused on demonstrating a versatile configuration of an ED tether to overcome the limitations of the rocket equation, enable new classes of missions currently unaffordable or infeasible, and significantly advance the Technology Readiness Level (TRL) to an operational level. We are also focused on establishing a far deeper understanding of critical processes and technologies to be able to scale and improve tether systems in the future. Here, we provide an overview of the proposed PROPEL mission. One of the critical processes for efficient ED tether operation is the ability to inject current to and collect current from the ionosphere. Because the PROPEL mission is planned to have both boost and deboost capability using a single tether, the tether current must be capable of flowing in both directions and at levels well over 1 A. Given the greater mobility of electrons over that of ions, this generally requires that both ends of the ED tether system can both collect and emit electrons. For example, hollow cathode plasma contactors (HCPCs) generally are viewed as state-of-the-art and high TRL devices; however, for ED tether applications important questions remain of how efficiently they can operate as both electron collectors and emitters. Other technologies will be highlighted that are being investigated as possible alternatives to the HCPC such as Solex that generates a plasma cloud from a solid material (Teflon) and electron emission (only) technologies such as cold-cathode electron field emission or photo-electron beam generation (PEBG) techniques.

Operational evaluation of a proppeller test stand in the quiet flow facility at Langley Research Center

NASA Technical Reports Server (NTRS)

Block, P. J. W.

1982-01-01

Operational proof tests of a propeller test stand (PTS) in a quiet flow facility (QFF) are presented. The PTS is an experimental test bed for acoustic propeller research in the quiet flow environment of the QFF. These proof tests validate thrust and torque predictions, examine the repeatability of measurements on the PTS, and determine the effect of applying artificial roughness to the propeller blades. Since a thrusting propeller causes an open jet to contract, the potential flow core was surveyed to examine the magnitude of the contraction. These measurements are compared with predicted values. The predictions are used to determine operational limitations for testing a given propeller design in the QFF.

Noise of the SR-3 propeller model at 2 deg and 4 deg angle of attack

NASA Technical Reports Server (NTRS)

Dittmar, J. H.; Jeracki, R. J.

1981-01-01

The noise effect of operating supersonic tip speed propellers at angle of attack with respect to the incoming flow was determined. Increases in the maximum blade passage noise were observed for the propeller operating at angle of attack. The noise increase was not symmetrical with one wall of the wind tunnel having significantly more noise increase than the other wall. This was apparently the result of the rotational direction of the propeller. The lack of symmetry of the noise at angle of attack to the use of oppositely rotating propellers on opposite sides of an airplane fuselage as a way of minimizing the noise due to operation at angle of attack.

14 CFR 35.23 - Propeller control system.

Code of Federal Regulations, 2011 CFR

2011-01-01

... propeller effect under the intended operating conditions. (4) The failure or corruption of data or signals... corruption of airplane-supplied data does not result in hazardous propeller effects. (e) The propeller...

14 CFR 35.23 - Propeller control system.

Code of Federal Regulations, 2010 CFR

2010-01-01

... propeller effect under the intended operating conditions. (4) The failure or corruption of data or signals... corruption of airplane-supplied data does not result in hazardous propeller effects. (e) The propeller...

Steady and Unsteady Loadings and Hydrodynamic Forces on Counterrotating Propellers.

DTIC Science & Technology

1976-07-01

forces and bending moments) of counterrotating propeller systems with equal and unequal number of blades operating in uniform and nonuniform inflow...1899 July 1976 STEADY AND UNSTEADY LOADIN GS AND HYDRODYNAM IC FORCES ON COUNTERROTATING PROPELLERS by S. Tsakonas, W. Jacobs and M. Afl This study...operator II , LINEARIZED UNSTEADY LIFTING SURFACE THEORY index of sunviiation Two counterrotating propellers are operatin g i n the flow of an ideal

14 CFR 21.129 - Tests: propellers.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Tests: propellers. 21.129 Section 21.129... PROCEDURES FOR PRODUCTS AND PARTS Production Under Type Certificate § 21.129 Tests: propellers. Each person... functional test to determine if it operates properly throughout the normal range of operation. ...

14 CFR 21.129 - Tests: propellers.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Tests: propellers. 21.129 Section 21.129... PROCEDURES FOR PRODUCTS AND PARTS Production Under Type Certificate § 21.129 Tests: propellers. Each person... functional test to determine if it operates properly throughout the normal range of operation. ...

14 CFR 35.3 - Instructions for propeller installation and operation.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Instructions for propeller installation and... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: PROPELLERS General § 35.3 Instructions for propeller.... Those approved instructions must contain: (a) Instructions for installing the propeller, which: (1...

14 CFR 35.3 - Instructions for propeller installation and operation.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Instructions for propeller installation and... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: PROPELLERS General § 35.3 Instructions for propeller.... Those approved instructions must contain: (a) Instructions for installing the propeller, which: (1...

14 CFR 35.3 - Instructions for propeller installation and operation.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Instructions for propeller installation and... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: PROPELLERS General § 35.3 Instructions for propeller.... Those approved instructions must contain: (a) Instructions for installing the propeller, which: (1...

14 CFR 35.3 - Instructions for propeller installation and operation.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Instructions for propeller installation and... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: PROPELLERS General § 35.3 Instructions for propeller.... Those approved instructions must contain: (a) Instructions for installing the propeller, which: (1...

14 CFR 35.3 - Instructions for propeller installation and operation.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Instructions for propeller installation and... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: PROPELLERS General § 35.3 Instructions for propeller.... Those approved instructions must contain: (a) Instructions for installing the propeller, which: (1...

14 CFR 121.643 - Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Supplemental operations.

Code of Federal Regulations, 2014 CFR

2014-01-01

...-powered airplanes: Supplemental operations. 121.643 Section 121.643 Aeronautics and Space FEDERAL AVIATION... Flight Release Rules § 121.643 Fuel supply: Nonturbine and turbo-propeller-powered airplanes... flight or takeoff a nonturbine or turbo-propeller-powered airplane unless, considering the wind and other...

14 CFR 121.643 - Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Supplemental operations.

Code of Federal Regulations, 2013 CFR

2013-01-01

...-powered airplanes: Supplemental operations. 121.643 Section 121.643 Aeronautics and Space FEDERAL AVIATION... Flight Release Rules § 121.643 Fuel supply: Nonturbine and turbo-propeller-powered airplanes... flight or takeoff a nonturbine or turbo-propeller-powered airplane unless, considering the wind and other...

14 CFR 121.643 - Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Supplemental operations.

Code of Federal Regulations, 2012 CFR

2012-01-01

...-powered airplanes: Supplemental operations. 121.643 Section 121.643 Aeronautics and Space FEDERAL AVIATION... Flight Release Rules § 121.643 Fuel supply: Nonturbine and turbo-propeller-powered airplanes... flight or takeoff a nonturbine or turbo-propeller-powered airplane unless, considering the wind and other...

14 CFR 21.129 - Tests: propellers.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Tests: propellers. 21.129 Section 21.129... PROCEDURES FOR PRODUCTS AND PARTS Production Under Type Certificate Only § 21.129 Tests: propellers. Each... acceptable functional test to determine if it operates properly throughout the normal range of operation. ...

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...

Investigation of lightweight designs and materials for LO2 and LH2 propellant tanks for space vehicles, phase 1

NASA Technical Reports Server (NTRS)

1975-01-01

Design, analysis, and fabrication studies were performed on nonintegral (suspended) tanks using a representative space tug design. The LH2 and LO2 tank concept selection was developed. Tank geometries and support relationships were investigated using tug design propellant inertias and ullage pressures, then compared based on total tug systems effects. The tank combinations which resulted in the maximum payload were selected. Tests were conducted on samples of membrane material which was processed in a manner simulating production tank fabrication operations to determine fabrication effects on the fracture toughness of the tank material. Fracture mechanics analyses were also performed to establish a preliminary set of allowables for initial defects.

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...

Evaluation of aero commander propeller acoustic data: Taxi operations

NASA Technical Reports Server (NTRS)

Piersol, A. G.; Wilby, E. G.; Wilby, J. F.

1979-01-01

The acoustic data from ground tests performed on an Aero Commander propeller driven aircraft are analyzed. An array of microphones flush mounted on the side of the fuselage were used to record data. The propeller blade passage noise during operations at several different taxi speeds is considered and calculations of the magnitude and phase of the blade passage tones, the amplitude stability of the tones, and the spatial phase and coherence of the tones are included. The measured results are compared to theoretical predictions for propeller noise and various evaluations which reveal important details of propeller noise characteristics are presented.

The acoustic experimental investigation of counterrotating propeller configurations

NASA Technical Reports Server (NTRS)

Gazzaniga, John A.

1987-01-01

An experimental study of scale counterrotating propellers operating in an anechoic facility has been conducted. Various configurations of counterrotation for equal numbers of blades per disk have been tested along with single-rotation propellers, underscoring the fundamental acoustic differences between single and counterrotation propeller operation. In addition it is shown that, as the loading on the counterrotating system is increased, the overall sound-pressure level is also increased in both the disk plane and axial direction.

14 CFR 121.641 - Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Flag operations.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Flag operations. 121.641 Section 121.641 Aeronautics and Space FEDERAL AVIATION... Flight Release Rules § 121.641 Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Flag...

14 CFR 121.641 - Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Flag operations.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 3 2011-01-01 2011-01-01 false Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Flag operations. 121.641 Section 121.641 Aeronautics and Space FEDERAL AVIATION... Flight Release Rules § 121.641 Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Flag...

14 CFR 121.643 - Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Supplemental operations.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 3 2011-01-01 2011-01-01 false Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Supplemental operations. 121.643 Section 121.643 Aeronautics and Space FEDERAL AVIATION... Flight Release Rules § 121.643 Fuel supply: Nonturbine and turbo-propeller-powered airplanes...

14 CFR 121.643 - Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Supplemental operations.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Fuel supply: Nonturbine and turbo-propeller-powered airplanes: Supplemental operations. 121.643 Section 121.643 Aeronautics and Space FEDERAL AVIATION... Flight Release Rules § 121.643 Fuel supply: Nonturbine and turbo-propeller-powered airplanes...

Tripropellant Engine Study

NASA Technical Reports Server (NTRS)

Wheeler, D. B.

1977-01-01

The feasibility of modifying the space shuttle main engine (SSME) for dual mode operation was investigated. Various high power cycle engine configurations derived from the SSME were configurations that will allow sequential burning of LOX/hydrocarbon and LOX/hydrogen were studied in order to identify concepts that make maximum use of SSME hardware and best satisfy the dual mode booster engine system application. Engine cycles were formulated for LOX/RP-1, LOX/CH4, and LOX/C3H8 propellants. Flow rates and operating cycles were established and the adaptability of the major components of the SSME was evaluated.

76 FR 74749 - Critical Parts for Airplane Propellers

Federal Register 2010, 2011, 2012, 2013, 2014

2011-12-01

... manufacturer, and establish engineering, manufacture, and maintenance processes for those parts. The intended... a specific definition for a propeller critical part, or-- [rtarr9] Require type certificate holders..., however, has regulations that identify a specific definition for propeller critical part, and regulations...

High-Lift Propeller Noise Prediction for a Distributed Electric Propulsion Flight Demonstrator

NASA Technical Reports Server (NTRS)

Nark, Douglas M.; Buning, Pieter G.; Jones, William T.; Derlaga, Joseph M.

2017-01-01

Over the past several years, the use of electric propulsion technologies within aircraft design has received increased attention. The characteristics of electric propulsion systems open up new areas of the aircraft design space, such as the use of distributed electric propulsion (DEP). In this approach, electric motors are placed in many different locations to achieve increased efficiency through integration of the propulsion system with the airframe. Under a project called Scalable Convergent Electric Propulsion Technology Operations Research (SCEPTOR), NASA is designing a flight demonstrator aircraft that employs many "high-lift propellers" distributed upstream of the wing leading edge and two cruise propellers (one at each wingtip). As the high-lift propellers are operational at low flight speeds (take-off/approach flight conditions), the impact of the DEP configuration on the aircraft noise signature is also an important design consideration. This paper describes efforts toward the development of a mulit-fidelity aerodynamic and acoustic methodology for DEP high-lift propeller aeroacoustic modeling. Specifically, the PAS, OVERFLOW 2, and FUN3D codes are used to predict the aerodynamic performance of a baseline high-lift propeller blade set. Blade surface pressure results from the aerodynamic predictions are then used with PSU-WOPWOP and the F1A module of the NASA second generation Aircraft NOise Prediction Program to predict the isolated high-lift propeller noise source. Comparisons of predictions indicate that general trends related to angle of attack effects at the blade passage frequency are captured well with the various codes. Results for higher harmonics of the blade passage frequency appear consistent for the CFD based methods. Conversely, evidence of the need for a study of the effects of increased azimuthal grid resolution on the PAS based results is indicated and will be pursued in future work. Overall, the results indicate that the computational approach is acceptable for fundamental assessment of low-noise high-lift propeller designs. The extent to which the various approaches may be used in a complementary manner will be further established as measured data becomes available for validation. Ultimately, it is anticipated that this combined approach may be used to provide realistic incident source fields for acoustic shielding/scattering studies on various aircraft configurations.

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.

Robust Exploration and Commercial Missions to the Moon Using NTR LANTR Propulsion and Lunar-Derived Propellants

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Ryan, Stephen W.; Burke, Laura M.; McCurdy, David R.; Fittje, James E.; Joyner, Claude R.

2017-01-01

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable access through cislunar space necessary for commercial development and sustained human presence on the Moon. In his post-Apollo Integrated Space Program Plan (1970-1990), Wernher von Braun, proposed a reusable nuclear thermal propulsion stage (NTPS) to deliver cargo and crew to the Moon to establish a lunar base before undertaking human missions to Mars. The NTR option was selected by von Braun because it was a demonstrated technology capable of generating both high thrust and high specific impulse (Isp 900 s) twice that of todays best chemical rockets. In NASAs Mars Design Reference Architecture (DRA) 5.0 study, the crewed Mars transfer vehicle used three 25 klbf Pewee engines the smallest and highest performing engine tested in the Rover program along with graphite composite fuel. Smaller, lunar transfer vehicles consisting of a NTPS using three approximately 16.5 klbf Small Nuclear Rocket Engines (SNREs), an in-line propellant tank, plus the payload can enable a variety of reusable lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong tourism missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing an affordable in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in-situ propellant production used to supply strategically located propellant depots and transportation nodes. The utilization of iron-rich volcanic glass or lunar polar ice (LPI) deposits (each estimated at billions of metric tons) for propellant production can significantly reduce the launch mass requirements from Earth and can enable reusable, surface-based lunar landing vehicles (LLVs) using liquid oxygen/hydrogen (LOX/LH2) chemical rocket engines. Afterwards, LOX/LH2 propellant depots can be established in lunar equatorial and polar orbits to supply the LTS. At this point a modified version of the conventional NTR called the LOX-augmented NTR, or LANTR would be introduced into the LTS allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an afterburner into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engines choked sonic throat essentially scramjet propulsion in reverse. By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and Isp values while the reactor core power level remains relatively constant. Eventually, a LANTR-based LTS can enable a rapid commuter shuttle with one-way trip times to and from the Moon ranging from 36 to 24 hours. Even if only 1 of the extracted propellant from identified volcanic glass and polar ice deposits were available for use in lunar orbit, such a supply could support daily commuter flights to the Moon for many thousands of years! An evolutionary mission architecture is outlined and a variety of lunar missions and transfer vehicle designs are examined, along with the increasing demands on propellant production as mission complexity increases. A comparison of vehicle features and engine operating characteristics, for both NTR and LANTR engines, is also provided along with a brief discussion on the propellant production issues associated with using volcanic glass and LPI as source material.

Experimental study of the effects of installation on singleand counter-rotation propeller noise

NASA Technical Reports Server (NTRS)

Block, P. J. W.

1986-01-01

Measurements which are required to define the directivity and the level of propeller noise were studied. The noise radiation pattern for various single-rotation (SR) propeller and counter-rotation (CR) propeller installations were mapped. The measurements covered + or - 60 deg from the propeller disk plane and + or - 60 deg in the cross-stream direction. Configurations examined included SR and CR propellers at angle of attack and an SR pusher installation. The increases in noise that arise from an unsteady loading operation such as an SR pusher or a CR exceeded 15 dB in the forward axial direction. Most of the additional noise radiates in the axial directions for unsteady loading operations of both the SR pusher and the CR tractor.

Direct electrical arc ignition of hybrid rocket motors

NASA Astrophysics Data System (ADS)

Judson, Michael I., Jr.

Hybrid rockets motors provide distinct safety advantages when compared to traditional liquid or solid propellant systems, due to the inherent stability and relative inertness of the propellants prior to established combustion. As a result of this inherent propellant stability, hybrid motors have historically proven difficult to ignite. State of the art hybrid igniter designs continue to require solid or liquid reactants distinct from the main propellants. These ignition methods however, reintroduce to the hybrid propulsion system the safety and complexity disadvantages associated with traditional liquid or solid propellants. The results of this study demonstrate the feasibility of a novel direct electrostatic arc ignition method for hybrid motors. A series of small prototype stand-alone thrusters demonstrating this technology were successfully designed and tested using Acrylonitrile Butadiene Styrene (ABS) plastic and Gaseous Oxygen (GOX) as propellants. Measurements of input voltage and current demonstrated that arc-ignition will occur using as little as 10 watts peak power and less than 5 joules total energy. The motor developed for the stand-alone small thruster was adapted as a gas generator to ignite a medium-scale hybrid rocket motor using nitrous oxide /and HTPB as propellants. Multiple consecutive ignitions were performed. A large data set as well as a collection of development `lessons learned' were compiled to guide future development and research. Since the completion of this original groundwork research, the concept has been developed into a reliable, operational igniter system for a 75mm hybrid motor using both gaseous oxygen and liquid nitrous oxide as oxidizers. A development map of the direct spark ignition concept is presented showing the flow of key lessons learned between this original work and later follow on development.

Space storable propellant performance program coaxial injector characterization

NASA Technical Reports Server (NTRS)

Burick, R. J.

1972-01-01

An experimental program was conducted to characterize the circular coaxial injector concept for application with the space-storable gas/liquid propellant combination FLOX(82.6% F2)/CH4(g) at high pressure. The primary goal of the program was to obtain high characteristic velocity efficiency in conjunction with acceptable injector/chamber compatibility. A series of subscale (single element) cold flow and hot fire experiments was employed to establish design criteria for a 3000-lbf (sea level) engine operating at 500 psia. The subscale experiments characterized both high performance core elements and peripheral elements with enhanced injector/chamber compatibility. The full-scale injector which evolved from the study demonstrated a performance level of 99 percent of the theoretical shifting characteristic exhaust velocity with low chamber heat flux levels. A 44-second-duration firing demonstrated the durability of the injector. Parametric data are presented that are applicable for the design of circular, coaxial injectors that operate with injection dynamics (fuel and oxidizer velocity, etc.) similar to those employed in the work reported.

Safety Standard for Hydrogen and Hydrogen Systems: Guidelines for Hydrogen System Design, Materials Selection, Operations, Storage and Transportation. Revision

NASA Technical Reports Server (NTRS)

1997-01-01

The NASA Safety Standard, which establishes a uniform process for hydrogen system design, materials selection, operation, storage, and transportation, is presented. The guidelines include suggestions for safely storing, handling, and using hydrogen in gaseous (GH2), liquid (LH2), or slush (SLH2) form whether used as a propellant or non-propellant. The handbook contains 9 chapters detailing properties and hazards, facility design, design of components, materials compatibility, detection, and transportation. Chapter 10 serves as a reference and the appendices contained therein include: assessment examples; scaling laws, explosions, blast effects, and fragmentation; codes, standards, and NASA directives; and relief devices along with a list of tables and figures, abbreviations, a glossary and an index for ease of use. The intent of the handbook is to provide enough information that it can be used alone, but at the same time, reference data sources that can provide much more detail if required.

Investigation of cryogenic rupture disc design

NASA Technical Reports Server (NTRS)

Keough, J. B.; Oldland, A. H.

1973-01-01

Rupture disc designs of both the active (command actuated) and passive (pressure ruptured) types were evaluated for performance characteristics at cryogenic temperatures and for capability to operate in a variety of cryogens, including gaseous and liquid fluorine. The test results, coupled with information from literature and industry searches, were used to establish a statement of design criteria and recommended practices for application of rupture discs to cryogenic rocket propellant feed and vent systems.

14 CFR 33.89 - Operation test.

Code of Federal Regulations, 2011 CFR

2011-01-01

..., acceleration, overspeeding, ignition, functioning of the propeller (if the engine is designated to operate with a propeller); (2) Compliance with the engine response requirements of § 33.73; and (3) The minimum... STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.89 Operation test. (a) The operation...

14 CFR 33.89 - Operation test.

Code of Federal Regulations, 2012 CFR

2012-01-01

..., acceleration, overspeeding, ignition, functioning of the propeller (if the engine is designated to operate with a propeller); (2) Compliance with the engine response requirements of § 33.73; and (3) The minimum... STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.89 Operation test. (a) The operation...

14 CFR 35.41 - Overspeed and overtorque.

Code of Federal Regulations, 2014 CFR

2014-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS Tests and Inspections § 35.41 Overspeed and overtorque. (a) When the applicant seeks approval of a transient maximum propeller overspeed, the applicant must demonstrate that the propeller is capable of further operation without maintenance action at the maximum propeller overspeed...

14 CFR 35.41 - Overspeed and overtorque.

Code of Federal Regulations, 2012 CFR

2012-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS Tests and Inspections § 35.41 Overspeed and overtorque. (a) When the applicant seeks approval of a transient maximum propeller overspeed, the applicant must demonstrate that the propeller is capable of further operation without maintenance action at the maximum propeller overspeed...

14 CFR 35.41 - Overspeed and overtorque.

Code of Federal Regulations, 2010 CFR

2010-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS Tests and Inspections § 35.41 Overspeed and overtorque. (a) When the applicant seeks approval of a transient maximum propeller overspeed, the applicant must demonstrate that the propeller is capable of further operation without maintenance action at the maximum propeller overspeed...

14 CFR 35.41 - Overspeed and overtorque.

Code of Federal Regulations, 2011 CFR

2011-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS Tests and Inspections § 35.41 Overspeed and overtorque. (a) When the applicant seeks approval of a transient maximum propeller overspeed, the applicant must demonstrate that the propeller is capable of further operation without maintenance action at the maximum propeller overspeed...

14 CFR 35.41 - Overspeed and overtorque.

Code of Federal Regulations, 2013 CFR

2013-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS Tests and Inspections § 35.41 Overspeed and overtorque. (a) When the applicant seeks approval of a transient maximum propeller overspeed, the applicant must demonstrate that the propeller is capable of further operation without maintenance action at the maximum propeller overspeed...

49 CFR 390.21 - Marking of self-propelled CMVs and intermodal equipment.

Code of Federal Regulations, 2011 CFR

2011-10-01

... 49 Transportation 5 2011-10-01 2011-10-01 false Marking of self-propelled CMVs and intermodal... Marking of self-propelled CMVs and intermodal equipment. (a) General. Every self-propelled CMV subject to...) The legal name or a single trade name of the motor carrier operating the self-propelled CMV, as listed...

49 CFR 390.21 - Marking of self-propelled CMVs and intermodal equipment.

Code of Federal Regulations, 2010 CFR

2010-10-01

... 49 Transportation 5 2010-10-01 2010-10-01 false Marking of self-propelled CMVs and intermodal... Marking of self-propelled CMVs and intermodal equipment. (a) General. Every self-propelled CMV subject to...) The legal name or a single trade name of the motor carrier operating the self-propelled CMV, as listed...

49 CFR 390.21 - Marking of self-propelled CMVs and intermodal equipment.

Code of Federal Regulations, 2012 CFR

2012-10-01

... 49 Transportation 5 2012-10-01 2012-10-01 false Marking of self-propelled CMVs and intermodal... Marking of self-propelled CMVs and intermodal equipment. (a) General. Every self-propelled CMV subject to...) The legal name or a single trade name of the motor carrier operating the self-propelled CMV, as listed...

14 CFR 23.33 - Propeller speed and pitch limits.

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 limits. 23.33... Propeller speed and pitch limits. (a) General. The propeller speed and pitch must be limited to values that... the all engine(s) operating climb speed specified in § 23.65, the propeller must limit the engine r.p...

14 CFR 23.33 - Propeller speed and pitch limits.

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 limits. 23.33... Propeller speed and pitch limits. (a) General. The propeller speed and pitch must be limited to values that... the all engine(s) operating climb speed specified in § 23.65, the propeller must limit the engine r.p...

In-Space Cryogenic Propellant Depot Stepping Stone

NASA Technical Reports Server (NTRS)

Howell, Joe T.; Mankins, John C.; Fikes, John C.

2005-01-01

An In-Space Cryogenic Propellant Depot (ISCPD) is an important stepping stone to provide the capability to preposition, store, manufacture, and later use the propellants for Earth-Neighborhood campaigns and beyond. An in-space propellant depot will provide affordable propellants and other similar consumables to support the development of sustainable and affordable exploration strategies as well as commercial space activities. An in-space propellant depot not only requires technology development in key areas such as zero boil-off storage and fluid transfer, but in other areas such as lightweight structures, highly reliable connectors, and autonomous operations. These technologies can be applicable to a broad range of propellant depot concepts or specific to a certain design. In addition, these technologies are required for spacecraft and orbit transfer vehicle propulsion and power systems, and space life support. Generally, applications of this technology require long-term storage, on-orbit fluid transfer and supply, cryogenic propellant production from water, unique instrumentation and autonomous operations. This paper discusses the reasons why such advances are important to future affordable and sustainable operations in space. This paper also discusses briefly R&D objectives comprising a promising approach to the systems planning and evolution into a meaningful stepping stone design, development, and implementation of an In-Space Cryogenic Propellant Depot. The success of a well-planned and orchestrated approach holds great promise for achieving innovation and revolutionary technology development for supporting future exploration and development of space.

14 CFR 21.16 - Special conditions.

Code of Federal Regulations, 2013 CFR

2013-01-01

... aircraft, aircraft engine, or propeller because of a novel or unusual design feature of the aircraft, aircraft engine or propeller, he prescribes special conditions and amendments thereto for the product. The... for the aircraft, aircraft engine or propeller as the FAA finds necessary to establish a level of...

14 CFR 21.16 - Special conditions.

Code of Federal Regulations, 2012 CFR

2012-01-01

... aircraft, aircraft engine, or propeller because of a novel or unusual design feature of the aircraft, aircraft engine or propeller, he prescribes special conditions and amendments thereto for the product. The... for the aircraft, aircraft engine or propeller as the FAA finds necessary to establish a level of...

14 CFR 21.16 - Special conditions.

Code of Federal Regulations, 2014 CFR

2014-01-01

... aircraft, aircraft engine, or propeller because of a novel or unusual design feature of the aircraft, aircraft engine or propeller, he prescribes special conditions and amendments thereto for the product. The... for the aircraft, aircraft engine or propeller as the FAA finds necessary to establish a level of...

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.

Sources and characteristics of interior noise in general aviation aircraft

NASA Technical Reports Server (NTRS)

Catherines, J. J.; Jha, S. K.

1976-01-01

A field study was conducted to examine the interior noise characteristics of a general aviation aircraft. The goals were to identify the major noise sources and their relative contribution and to establish the noise transmission paths and their relative importance. Tests were performed on an aircraft operating under stationary conditions on the ground. Results show that the interior noise level of light aircraft is dominated by broadband, low frequencies (below 1,000 Hz). Both the propeller and the engine are dominant sources, however, the contribution from the propeller is significantly more than the engine at its fundamental blade passage frequency. The data suggest that the airborne path is more dominant than the structure-borne path in the transmission of broadband, low frequency noise which apparently results from the exhaust.

Noise of the SR-6 propeller model at 2 deg and 4 deg angles of attack

NASA Technical Reports Server (NTRS)

Dittmar, J. H.; Stefko, G. L.

1983-01-01

The noise generated by supersonic-tip-speed propellers creates a cabin noise problem for future airplanes powered by these propellers. Noise of a number of propeller models were measured in the NASA Lewis 8- by 6-Foot Wind Tunnel with flow parallel to the propeller axis. In flight, as a result of the induced upwash from the airplane wing, the propeller is at an angle of attack with respect to the incoming flow. Therefore, the 10-blade SR-6 propeller was operated at angle of attack to determine its noise behavior. Higher blade passage tones were observed for the propeller operating at angle of attack in a 0.6 axial Mach number flow. The noise increase was not symmetrical, with one wall of the wind tunnel showing a larger noise increase than the other wall. No noise increase was observed at angle of attack in a 0.8 axial Mach number flow. For this propeller the dominance of thickness noise, which does not increase with angle of attack, explains the lack of noise increase at the higher 0.8 Mach number.

Study of alternate methods of disposal of propellants and gases at KSC

NASA Technical Reports Server (NTRS)

Moore, W. I.

1970-01-01

A comprehensive study was conducted at KSC launch support facilities to determine the nature and extent of potential hazards from propellant and gas releases to the environment. The results of the study, alternate methods for reducing or eliminating the hazards, and recommendations pertaining to these alternatives are presented. The operational modes of the propellant or hazardous gas systems considered include: system charging, system standby, system operation, and post-test operations. The results are outlined on an area-by-area basis.

14 CFR 35.37 - Fatigue limits and evaluation.

Code of Federal Regulations, 2014 CFR

2014-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS Tests and Inspections § 35.37 Fatigue limits and evaluation. This section does not apply to fixed-pitch wood propellers of conventional design. (a) Fatigue limits must be established by tests, or analysis based on tests, for propeller: (1) Hubs. (2) Blades. (3) Blade retention...

14 CFR 35.37 - Fatigue limits and evaluation.

Code of Federal Regulations, 2012 CFR

2012-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS Tests and Inspections § 35.37 Fatigue limits and evaluation. This section does not apply to fixed-pitch wood propellers of conventional design. (a) Fatigue limits must be established by tests, or analysis based on tests, for propeller: (1) Hubs. (2) Blades. (3) Blade retention...

14 CFR 35.37 - Fatigue limits and evaluation.

Code of Federal Regulations, 2013 CFR

2013-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS Tests and Inspections § 35.37 Fatigue limits and evaluation. This section does not apply to fixed-pitch wood propellers of conventional design. (a) Fatigue limits must be established by tests, or analysis based on tests, for propeller: (1) Hubs. (2) Blades. (3) Blade retention...

14 CFR 35.37 - Fatigue limits and evaluation.

Code of Federal Regulations, 2011 CFR

2011-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS Tests and Inspections § 35.37 Fatigue limits and evaluation. This section does not apply to fixed-pitch wood propellers of conventional design. (a) Fatigue limits must be established by tests, or analysis based on tests, for propeller: (1) Hubs. (2) Blades. (3) Blade retention...

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."

An analysis for high speed propeller-nacelle aerodynamic performance prediction. Volume 1: Theory and application

NASA Technical Reports Server (NTRS)

Egolf, T. Alan; Anderson, Olof L.; Edwards, David E.; Landgrebe, Anton J.

1988-01-01

A computer program, the Propeller Nacelle Aerodynamic Performance Prediction Analysis (PANPER), was developed for the prediction and analysis of the performance and airflow of propeller-nacelle configurations operating over a forward speed range inclusive of high speed flight typical of recent propfan designs. A propeller lifting line, wake program was combined with a compressible, viscous center body interaction program, originally developed for diffusers, to compute the propeller-nacelle flow field, blade loading distribution, propeller performance, and the nacelle forebody pressure and viscous drag distributions. The computer analysis is applicable to single and coaxial counterrotating propellers. The blade geometries can include spanwise variations in sweep, droop, taper, thickness, and airfoil section type. In the coaxial mode of operation the analysis can treat both equal and unequal blade number and rotational speeds on the propeller disks. The nacelle portion of the analysis can treat both free air and tunnel wall configurations including wall bleed. The analysis was applied to many different sets of flight conditions using selected aerodynamic modeling options. The influence of different propeller nacelle-tunnel wall configurations was studied. Comparisons with available test data for both single and coaxial propeller configurations are presented along with a discussion of the results.

The effects of installation on single- and counter-rotation propeller noise

NASA Technical Reports Server (NTRS)

Block, P. J. W.

1984-01-01

In order to understand the effects of installation on propeller noise, numerous measurements are required to define the directivity of the noise as well as the level. An experimental study was designed to map the noise radiation pattern for various single-rotation propeller (SRP) and counter-rotation propeller (CRP) installations covering + or 60 deg from the propeller disk plane and + or - 60 deg laterally. Configurations that were considered included an SRP at angle of attack and in tractor and pusher operations and a CRP. A first principles linear theory was validated for the SRP tractor operation over the angle range mentioned above. The increases in noise that arise from an unsteady loading operation such as an SRP pusher or CRP exceed 15 dB and depend on the observer location. In particular, the majority of the additional noise appears to radiate in the axial directions.

Fuels and Space Propellants for Reusable Launch Vehicles: A Small Business Innovation Research Topic and Its Commercial Vision

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan A.

1997-01-01

Under its Small Business Innovation Research (SBIR) program (and with NASA Headquarters support), the NASA Lewis Research Center has initiated a topic entitled "Fuels and Space Propellants for Reusable Launch Vehicles." The aim of this project would be to assist in demonstrating and then commercializing new rocket propellants that are safer and more environmentally sound and that make space operations easier. Soon it will be possible to commercialize many new propellants and their related component technologies because of the large investments being made throughout the Government in rocket propellants and the technologies for using them. This article discusses the commercial vision for these fuels and propellants, the potential for these propellants to reduce space access costs, the options for commercial development, and the benefits to nonaerospace industries. This SBIR topic is designed to foster the development of propellants that provide improved safety, less environmental impact, higher density, higher I(sub sp), and simpler vehicle operations. In the development of aeronautics and space technology, there have been limits to vehicle performance imposed by traditionally used propellants and fuels. Increases in performance are possible with either increased propellant specific impulse, increased density, or both. Flight system safety will also be increased by the use of denser, more viscous propellants and fuels.

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.

Establishment of design criteria for acceptable failure modes and fail safe considerations for the space shuttle structural system

NASA Technical Reports Server (NTRS)

Westrup, R. W.

1972-01-01

Investigations of fatigue life, and safe-life and fail-safe design concepts as applied to space shuttle structure are summarized. The results are evaluated to select recommended structural design criteria to provide assurance that premature failure due to propagation of undetected crack-like defects will not occur during shuttle operational service. The space shuttle booster, GDC configuration B-9U, is selected as the reference vehicle. Structural elements used as basis of detail analyses include wing spar caps, vertical stabilizer skins, crew compartment skin, orbiter support frame, and propellant tank shell structure. Fatigue life analyses of structural elements are performed to define potential problem areas and establish upper limits of operating stresses. Flaw growth analyses are summarized in parametric form over a range of initial flaw types and sizes, operating stresses and service life requirements. Service life of 100 to 500 missions is considered.

14 CFR 25.771 - Pilot compartment.

Code of Federal Regulations, 2014 CFR

2014-01-01

... the propellers so that no member of the minimum flight crew (established under § 25.1523), or part of 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 five...

14 CFR 25.771 - Pilot compartment.

Code of Federal Regulations, 2013 CFR

2013-01-01

... the propellers so that no member of the minimum flight crew (established under § 25.1523), or part of 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 five...

14 CFR 25.771 - Pilot compartment.

Code of Federal Regulations, 2012 CFR

2012-01-01

... the propellers so that no member of the minimum flight crew (established under § 25.1523), or part of 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 five...

14 CFR 25.771 - Pilot compartment.

Code of Federal Regulations, 2011 CFR

2011-01-01

... the propellers so that no member of the minimum flight crew (established under § 25.1523), or part of 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 five...

Refueling with In-Situ Produced Propellants

NASA Technical Reports Server (NTRS)

Chato, David J.

2014-01-01

In-situ produced propellants have been identified in many architecture studies as key to implementing feasible chemical propulsion missions to destinations beyond lunar orbit. Some of the more noteworthy ones include: launching from Mars to return to Earth (either direct from the surface, or via an orbital rendezvous); using the Earth-Moon Lagrange point as a place to refuel Mars transfer stages with Lunar surface produced propellants; and using Mars Moon Phobos as a place to produce propellants for descent and ascent stages bound for the Mars surface. However successful implementation of these strategies require an ability to successfully transfer propellants from the in-situ production equipment into the propellant tankage of the rocket stage used to move to the desired location. In many circumstances the most desirable location for this transfer to occur is in the low-gravity environment of space. In support of low earth orbit propellant depot concepts, extensive studies have been conducted on transferring propellants in-space. Most of these propellant transfer techniques will be applicable to low gravity operations in other locations. Even ground-based transfer operations on the Moon, Mars, and especially Phobos could benefit from the propellant conserving techniques used for depot refueling. This paper will review the literature of in-situ propellants and refueling to: assess the performance benefits of the use in-situ propellants for mission concepts; review the parallels with propellant depot efforts; assess the progress of the techniques required; and provide recommendations for future research.

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...

14 CFR 25.33 - Propeller speed and pitch limits.

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 limits. 25.33... AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Flight General § 25.33 Propeller speed and pitch limits. (a) The propeller speed and pitch must be limited to values that will ensure— (1) Safe operation...

14 CFR 25.33 - Propeller speed and pitch limits.

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 limits. 25.33... AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Flight General § 25.33 Propeller speed and pitch limits. (a) The propeller speed and pitch must be limited to values that will ensure- (1) Safe operation...

A Review of Solar-Powered Aircraft Flight Activity at the Pacific Missile Range Test Facility, Kauai, Hawaii

NASA Technical Reports Server (NTRS)

Ehernberger, L. J.; Donohue, Casey; Teets, Edward H., Jr.

2004-01-01

A series of solar-powered aircraft have been designed and operated by AeroVironment, Inc. (Monrovia, CA) as a part of National Aeronautics and Space Administration (NASA) objectives to develop energy-efficient high-altitude long-endurance platforms for earth observations and communications applications. Flight operations have been conducted at NASA's Dryden Flight Research Center, Edwards CA and at the U.S. Navy Pacific Missile Range Facility (PMRF) at Barking Sands, Kauai, HI. These aircraft flown at PMRF are named Pathfinder , Pathfinder Plus and Helios . Sizes of these three aircraft range from 560 lb with a 99-ft wingspan to 2300 lb with a 247-ft wingspan. Available payload capacity reaches approximately 200 lb. Pathfinder uses six engines and propellers: Pathfinder Plus 8; and Helios 14. The 2003 Helios fuel cell configurations used 10 engines and propellers. The PMRF was selected as a base of operations because if offers optimal summertime solar exposure, low prevailing wind-speeds on the runway, modest upper-air wind-speeds and the availability of suitable airspace. Between 1997 and 2001, successive altitude records of 71,530 ft, 80,200 ft, and 96,863 ft were established. Flight durations extended to 18 hours.

Technology Area Roadmap for In-Space Propulsion Technologies

NASA Technical Reports Server (NTRS)

Johnson, Les; Meyer, Michael; Palaszewski, Bryan; Coote, David; Goebel, Dan; White, Harold

2012-01-01

The exponential increase of launch system size.and cost.with delta-V makes missions that require large total impulse cost prohibitive. Led by NASA fs Marshall Space Flight Center, a team from government, industry, and academia has developed a flight demonstration mission concept of an integrated electrodynamic (ED) tethered satellite system called PROPEL: \\Propulsion using Electrodynamics.. The PROPEL Mission is focused on demonstrating a versatile configuration of an ED tether to overcome the limitations of the rocket equation, enable new classes of missions currently unaffordable or infeasible, and significantly advance the Technology Readiness Level (TRL) to an operational level. We are also focused on establishing a far deeper understanding of critical processes and technologies to be able to scale and improve tether systems in the future. Here, we provide an overview of the proposed PROPEL mission. One of the critical processes for efficient ED tether operation is the ability to inject current to and collect current from the ionosphere. Because the PROPEL mission is planned to have both boost and deboost capability using a single tether, the tether current must be capable of flowing in both directions and at levels well over 1 A. Given the greater mobility of electrons over that of ions, this generally requires that both ends of the ED tether system can both collect and emit electrons. For example, hollow cathode plasma contactors (HCPCs) generally are viewed as state-of-the-art and high TRL devices; however, for ED tether applications important questions remain of how efficiently they can operate as both electron collectors and emitters. Other technologies will be highlighted that are being investigated as possible alternatives to the HCPC such as Solex that generates a plasma cloud from a solid material (Teflon) and electron emission (only) technologies such as cold-cathode electron field emission or photo-electron beam generation (PEBG) techniques

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.

Acoustic and aerodynamic study of a pusher-propeller aircraft model

NASA Astrophysics Data System (ADS)

Soderman, Paul T.; Horne, W. Clifton

1990-09-01

An aerodynamic and acoustic study was made of a pusher-propeller aircraft model in the NASA-Ames 7 x 10 ft Wind Tunnel. The test section was changed to operate as an open jet. The 591 mm diameter unswept propeller was operated alone and in the wake of three empennages: an I tail, Y tail, and a V tail. The radiated noise and detailed wake properties were measured. Results indicate that the unsteady blade loading caused by the blade interactions with the wake mean velocity distribution had a strong effect on the harmonics of blade passage noise. The blade passage harmonics above the first were substantially increased in all horizontal directions by the empennage/propeller interaction. Directivity in the plane of the propeller was maximum perpendicular to the blade surface. Increasing the tail loading caused the propeller harmonics to increase 3 to 5 dB for an empennage/propeller spacing of 0.38 mean empennage chords. The interaction noise became weak as empennage propeller spacing was increased beyond 1.0 mean empennage chord lengths. Unlike the mean wake deficit, the wake turbulence had only a small effect on the propeller noise, that effect being a small increase in the broadband noise.

Acoustic and aerodynamic study of a pusher-propeller aircraft model

NASA Technical Reports Server (NTRS)

Soderman, Paul T.; Horne, W. Clifton

1990-01-01

An aerodynamic and acoustic study was made of a pusher-propeller aircraft model in the NASA-Ames 7 x 10 ft Wind Tunnel. The test section was changed to operate as an open jet. The 591 mm diameter unswept propeller was operated alone and in the wake of three empennages: an I tail, Y tail, and a V tail. The radiated noise and detailed wake properties were measured. Results indicate that the unsteady blade loading caused by the blade interactions with the wake mean velocity distribution had a strong effect on the harmonics of blade passage noise. The blade passage harmonics above the first were substantially increased in all horizontal directions by the empennage/propeller interaction. Directivity in the plane of the propeller was maximum perpendicular to the blade surface. Increasing the tail loading caused the propeller harmonics to increase 3 to 5 dB for an empennage/propeller spacing of 0.38 mean empennage chords. The interaction noise became weak as empennage propeller spacing was increased beyond 1.0 mean empennage chord lengths. Unlike the mean wake deficit, the wake turbulence had only a small effect on the propeller noise, that effect being a small increase in the broadband noise.

Oxygen-Methane Thruster

NASA Technical Reports Server (NTRS)

Pickens, Tim

2012-01-01

An oxygen-methane thruster was conceived with integrated igniter/injector capable of nominal operation on either gaseous or liquid propellants. The thruster was designed to develop 100 lbf (approximately 445 N) thrust at vacuum conditions and use oxygen and methane as propellants. This continued development included refining the design of the thruster to minimize part count and manufacturing difficulties/cost, refining the modeling tools and capabilities that support system design and analysis, demonstrating the performance of the igniter and full thruster assembly with both gaseous and liquid propellants, and acquiring data from this testing in order to verify the design and operational parameters of the thruster. Thruster testing was conducted with gaseous propellants used for the igniter and thruster. The thruster was demonstrated to work with all types of propellant conditions, and provided the desired performance. Both the thruster and igniter were tested, as well as gaseous propellants, and found to provide the desired performance using the various propellant conditions. The engine also served as an injector testbed for MSFC-designed refractory combustion chambers made of rhenium.

Annoyance caused by propeller airplane flyover noise

NASA Technical Reports Server (NTRS)

Mccurdy, D. A.; Powell, C. A.

1984-01-01

Laboratory experiments were conducted to provide information on quantifying the annoyance response of people to propeller airplane noise. The items of interest were current noise metrics, tone corrections, duration corrections, critical band corrections, and the effects of engine type, operation type, maximum takeoff weight, blade passage frequency, and blade tip speed. In each experiment, 64 subjects judged the annoyance of recordings of propeller and jet airplane operations presented at d-weighted sound pressure levels of 70, 80, and 90 dB in a testing room which simulates the outdoor acoustic environment. The first experiment examined 11 propeller airplanes with maximum takeoff weights greater than or equal to 5700 kg. The second experiment examined 14 propeller airplanes weighting 5700 kg or less. Five jet airplanes were included in each experiment. For both the heavy and light propeller airplanes, perceived noise level and perceived level (Stevens Mark VII procedure) predicted annoyance better than other current noise metrics.

Evaluation of aero Commander propeller acoustic data: Static operations

NASA Technical Reports Server (NTRS)

Piersol, A. G.; Wilby, E. G.; Wilby, J. F.

1978-01-01

Acoustic data are analyzed from a series of ground tests performed on an Aero Commander propeller-driven aircraft with an array of microphones flush-mounted on one side of the fuselage. The analyses were concerned with the propeller blade passage noise during static operation at several different engine speeds and included calculations of the magnitude and phase of the blade passage tones, the amplitude stability of the tones, and the spatial phase and coherence of the tones. The results indicate that the pressure field impinging on the fuselage represents primarily aerodynamic (near field) effects in the plane of the propeller at all frequencies. Forward and aft of the propeller plane aerodynamic effects still dominate the pressure field at frequencies below 200 Hz; but at higher frequencies, the pressure field is due to acoustic propagation from an equivalent center located about 0.15 to 0.30 blade diameters inboard from the propeller hub.

Effective Utilization of Resources and Infrastructure for a Spaceport Network Architecture

NASA Technical Reports Server (NTRS)

Gill, Tracy; Larson, Wiley; Mueller, Robert; Roberson, Luke

2012-01-01

Providing routine, affordable access to a variety of orbital and deep space destinations requires an intricate network of ground, planetary surface, and space-based spaceports like those on Earth (land and sea), in various Earth orbits, and on other extraterrestrial surfaces. Advancements in technology and international collaboration are critical to establish a spaceport network that satisfies the requirements for private and government research, exploration, and commercial objectives. Technologies, interfaces, assembly techniques, and protocols must be adapted to enable mission critical capabilities and interoperability throughout the spaceport network. The conceptual space mission architecture must address the full range of required spaceport services, from managing propellants for a variety of spacecraft to governance structure. In order to accomplish affordability and sustainability goals, the network architecture must consider deriving propellants from in situ planetary resources to the maximum extent possible. Water on the Moon and Mars, Mars' atmospheric CO2, and O2 extracted from lunar regolith are examples of in situ resources that could be used to generate propellants for various spacecraft, orbital stages and trajectories, and the commodities to support habitation and human operations at these destinations. The ability to use in-space fuel depots containing in situ derived propellants would drastically reduce the mass required to launch long-duration or deep space missions from Earth's gravity well. Advances in transformative technologies and common capabilities, interfaces, umbilicals, commodities, protocols, and agreements will facilitate a cost-effective, safe, reliable infrastructure for a versatile network of Earth- and extraterrestrial spaceports. Defining a common infrastructure on Earth, planetary surfaces, and in space, as well as deriving propellants from in situ planetary resources to construct in-space propellant depots to serve the spaceport network, will reduce exploration costs due to standardization of infrastructure commonality and reduction in number and types of interfaces and commodities.

Sources and characteristics of interior noise in general aviation aircraft

NASA Technical Reports Server (NTRS)

Catherines, J. J.; Jha, S. K.

1976-01-01

A field study has been conducted to examine the interior noise characteristics of a general aviation aircraft. The purposes of the study were to identify the major noise sources and their relative contribution and to establish the noise transmission paths and their relative importance. Tests were performed on an aircraft operating under stationary conditions on the ground. The results show that the interior noise level of light aircraft is dominated by broadband, low frequencies (below 1,000 Hz). Both the propeller and the engine are dominant sources; however, the contribution from the propeller is significantly more than the engine at its fundamental blade passage frequency. The data suggests that the airborne path is more dominant than the structure-borne path in the transmission of broadband, low-frequency noise which apparently results from the exhaust.

Investigation of the aerothermodynamics of hypervelocity reacting flows in the ram accelerator

NASA Technical Reports Server (NTRS)

Hertzberg, A.; Bruckner, A. P.; Mattick, A. T.; Knowlen, C.

1992-01-01

New diagnostic techniques for measuring the high pressure flow fields associated with high velocity ram accelerator propulsive modes was experimentally investigated. Individual propulsive modes are distinguished by their operating Mach number range and the manner in which the combustion process is initiated and stabilized. Operation of the thermally choked ram accelerator mode begins by injecting the projectile into the accelerator tube at a prescribed entrance velocity by means of a conventional light gas gun. A specially designed obturator, which is used to seal the bore of the gun, plays a key role in the ignition of the propellant gases in the subsonic combustion mode of the ram accelerator. Once ignited, the combustion process travels with the projectile and releases enough heat to thermally choke the flow within several tube diameters behind it, thereby stabilizing a high pressure zone on the rear of the projectile. When the accelerating projectile approaches the Chapman-Jouguet detonation speed of the propellant mixture, the combustion region is observed to move up onto the afterbody of the projectile as the pressure field evolves to a distinctively different form that implies the presence of supersonic combustion processes. Eventually, a high enough Mach number is reached that the ram effect is sufficient to cause the combustion process to occur entirely on the body. Propulsive cycles utilizing on-body heat release can be established either by continuously accelerating the projectile in a single propellant mixture from low initial in-tube Mach numbers (M less than 4) or by injecting the projectile at a speed above the propellant's Chapman-Jouguet detonation speed. The results of experimental and theoretical explorations of ram accelerator gas dynamic phenomena and the effectiveness of the new diagnostic techniques are presented in this report.

Tradespace Exploration of Distributed Propulsors for Advanced On-Demand Mobility Concepts

NASA Technical Reports Server (NTRS)

Borer, Nicholas K.; Moore, Mark D.; Turnbull, Andrew R.

2014-01-01

Combustion-based sources of shaft power tend to significantly penalize distributed propulsion concepts, but electric motors represent an opportunity to advance the use of integrated distributed propulsion on an aircraft. This enables use of propellers in nontraditional, non-thrust-centric applications, including wing lift augmentation, through propeller slipstream acceleration from distributed leading edge propellers, as well as wingtip cruise propulsors. Developing propellers for these applications challenges long-held constraints within propeller design, such as the notion of optimizing for maximum propulsive efficiency, or the use of constant-speed propellers for high-performance aircraft. This paper explores the design space of fixed-pitch propellers for use as (1) lift augmentation when distributed about a wing's leading edge, and (2) as fixed-pitch cruise propellers with significant thrust at reduced tip speeds for takeoff. A methodology is developed for evaluating the high-level trades for these types of propellers and is applied to the exploration of a NASA Distributed Electric Propulsion concept. The results show that the leading edge propellers have very high solidity and pitch well outside of the empirical database, and that the cruise propellers can be operated over a wide RPM range to ensure that thrust can still be produced at takeoff without the need for a pitch change mechanism. To minimize noise exposure to observers on the ground, both the leading edge and cruise propellers are designed for low tip-speed operation during takeoff, climb, and approach.

Low-speed wind tunnel performance of high-speed counterrotation propellers at angle-of-attack

NASA Technical Reports Server (NTRS)

Hughes, Christopher E.; Gazzaniga, John A.

1989-01-01

The low-speed aerodynamic performance characteristics of two advanced counterrotation pusher-propeller configurations with cruise design Mach numbers of 0.72 were investigated in the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel. The tests were conducted at Mach number 0.20, which is representative of the aircraft take-off/landing flight regime. The investigation determined the effect of nonuniform inflow on the propeller performance characteristics for several blade angle settings and a range of rotational speeds. The inflow was varied by yawing the propeller model to angle-of-attack by as much as plus or minus 16 degrees and by installing on the counterrotation propeller test rig near the propeller rotors a model simulator of an aircraft engine support pylon and fuselage. The results of the investigation indicated that the low-speed performance of the counterrotation propeller configurations near the take-off target operating points were reasonable and were fairly insensitive to changes in model angle-of-attack without the aircraft pylon/fuselage simulators installed on the propeller test rig. When the aircraft pylon/fuselage simulators were installed, small changes in propeller performance were seen at zero angle-of-attack, but fairly large changes in total power coefficient and very large changes of aft-to-forward-rotor torque ratio were produced when the propeller model was taken to angle-of-attack. The propeller net efficiency, though, was fairly insensitive to any changes in the propeller flowfield conditions near the take-off target operating points.

A new method in accelerating PROPELLER MRI.

PubMed

Li, Bing Keong; D'Arcy, Michael; Weber, Ewald; Crozier, Stuart

2008-01-01

In this work, a new method has been proposed to accelerate the PROPELLER MRI operation. The proposed method uses a rotary phased array coil and a new method in acquiring the k-space strips and preparing the complete k-space trajectories data set. It is numerically shown that for a 12 strips PROPELLER MR brain imaging sequence, the operation time can be reduced by four folds, with no apparent loss in the image quality.

Working Charts for the Selection of Aluminum Alloy Propellers of a Standard Form to Operate with Various Aircraft Engines and Bodies

NASA Technical Reports Server (NTRS)

Weick, Fred E

1931-01-01

Working charts are given for the convenient selection of aluminum alloy propellers of a standard form, to operate in connection with six different engine-fuselage combinations. The charts have been prepared from full-scale test data obtained in the 20-foot propeller research tunnel of the National Advisory Committee for Aeronautics. An example is also given showing the use of the charts.

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.

14 CFR 23.33 - Propeller speed and pitch limits.

Code of Federal Regulations, 2014 CFR

2014-01-01

... the all engine(s) operating climb speed specified in § 23.65, the propeller must limit the engine r.p... approved overspeed, a means to limit the maximum engine and propeller speed to not more than the maximum... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Propeller speed and pitch limits. 23.33...

14 CFR 23.33 - Propeller speed and pitch limits.

Code of Federal Regulations, 2013 CFR

2013-01-01

... the all engine(s) operating climb speed specified in § 23.65, the propeller must limit the engine r.p... approved overspeed, a means to limit the maximum engine and propeller speed to not more than the maximum... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Propeller speed and pitch limits. 23.33...

14 CFR 23.33 - Propeller speed and pitch limits.

Code of Federal Regulations, 2012 CFR

2012-01-01

... the all engine(s) operating climb speed specified in § 23.65, the propeller must limit the engine r.p... approved overspeed, a means to limit the maximum engine and propeller speed to not more than the maximum... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Propeller speed and pitch limits. 23.33...

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.

Navigation/Prop Software Suite

NASA Technical Reports Server (NTRS)

Bruchmiller, Tomas; Tran, Sanh; Lee, Mathew; Bucker, Scott; Bupane, Catherine; Bennett, Charles; Cantu, Sergio; Kwong, Ping; Propst, Carolyn

2012-01-01

Navigation (Nav)/Prop software is used to support shuttle mission analysis, production, and some operations tasks. The Nav/Prop suite containing configuration items (CIs) resides on IPS/Linux workstations. It features lifecycle documents, and data files used for shuttle navigation and propellant analysis for all flight segments. This suite also includes trajectory server, archive server, and RAT software residing on MCC/Linux workstations. Navigation/Prop represents tool versions established during or after IPS Equipment Rehost-3 or after the MCC Rehost.

A Design Method and an Application for Contrarotating Propellers

DTIC Science & Technology

1990-01-01

force gen- stricted to uniform flow , it fhowed that the analysis of CR pro- erated by the contrarotating propeller to be balanced by the drag... uniform flow at where the operating point of the propeller for a typical high-speed sur- ,/2 face ship. Force measurements for the CR propelier in... experimental thrust coefficient, torque Bronze. Since this propeller set is designed for uniform flow , coefficient, and efficiency for the CR propellers

Impact of Advanced Propeller Technology on Aircraft/Mission Characteristics of Several General Aviation Aircraft

NASA Technical Reports Server (NTRS)

Keiter, I. D.

1982-01-01

Studies of several General Aviation aircraft indicated that the application of advanced technologies to General Aviation propellers can reduce fuel consumption in future aircraft by a significant amount. Propeller blade weight reductions achieved through the use of composites, propeller efficiency and noise improvements achieved through the use of advanced concepts and improved propeller analytical design methods result in aircraft with lower operating cost, acquisition cost and gross weight.

14 CFR 35.17 - Materials and manufacturing methods.

Code of Federal Regulations, 2011 CFR

2011-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS Design and Construction § 35.17 Materials and manufacturing methods. (a) The suitability and durability of materials used in the propeller must: (1) Be established on the basis of...

14 CFR 35.17 - Materials and manufacturing methods.

Code of Federal Regulations, 2013 CFR

2013-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS Design and Construction § 35.17 Materials and manufacturing methods. (a) The suitability and durability of materials used in the propeller must: (1) Be established on the basis of...

14 CFR 35.17 - Materials and manufacturing methods.

Code of Federal Regulations, 2012 CFR

2012-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS Design and Construction § 35.17 Materials and manufacturing methods. (a) The suitability and durability of materials used in the propeller must: (1) Be established on the basis of...

14 CFR 35.17 - Materials and manufacturing methods.

Code of Federal Regulations, 2014 CFR

2014-01-01

... AIRWORTHINESS STANDARDS: PROPELLERS Design and Construction § 35.17 Materials and manufacturing methods. (a) The suitability and durability of materials used in the propeller must: (1) Be established on the basis of...

Conceptual design of an orbital propellant transfer experiment. Volume 2: Study results

NASA Technical Reports Server (NTRS)

Drake, G. L.; Bassett, C. E.; Merino, F.; Siden, L. E.; Bradley, R. E.; Carr, E. J.; Parker, R. E.

1980-01-01

The OTV configurations, operations and requirements planned for the period from the 1980's to the 1990's were reviewed and a propellant transfer experiment was designed that would support the needs of these advanced OTV operational concepts. An overall integrated propellant management technology plan for all NASA centers was developed. The preliminary cost estimate (for planning purposes only) is $56.7 M, of which approximately $31.8 M is for shuttle user costs.

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.

Wind tunnel performance results of swirl recovery vanes as tested with an advanced high speed propeller

NASA Technical Reports Server (NTRS)

Gazzaniga, John A.; Rose, Gayle E.

1992-01-01

Tests of swirl recovery vanes designed for use in conjunction with advanced high speed propellers were carried out at the NASA Lewis Research Center. The eight bladed 62.23 cm vanes were tested with a 62.23 cm SR = 7A high speed propeller in the NASA Lewis 2.44 x 1.83 m Supersonic Wind Tunnel for a Mach number range of 0.60 to 0.80. At the design operating condition for cruise of Mach 0.80 at an advance ratio of 3.26, the vane contribution to the total efficiency approached 2 percent. At lower off-design Mach numbers, the vane efficiency is even higher, approaching 4.5 percent for the Mach 0.60 condition. Use of the swirl recovery vanes essentially shifts the peak of the high speed propeller efficiency to a higher operating speed. This allows a greater degree of freedom in the selection of rpm over a wider operating range. Another unique result of the swirl recovery vane configuration is their essentially constant torque split between the propeller and the swirl vanes over a wide range of operating conditions for the design vane angle.

Study and design of cryogenic propellant acquisition systems. Volume 2: Supporting experimental program

NASA Technical Reports Server (NTRS)

Burge, G. W.; Blackmon, J. B.

1973-01-01

Areas of cryogenic fuel systems were identified where critical experimental information was needed either to define a design criteria or to establish the feasibility of a design concept or a critical aspect of a particular design. Such data requirements fell into three broad categories: (1) basic surface tension screen characteristics; (2) screen acquisition device fabrication problems; and (3) screen surface tension device operational failure modes. To explore these problems and to establish design criteria where possible, extensive laboratory or bench test scale experiments were conducted. In general, these proved to be quite successful and, in many instances, the test results were directly used in the system design analyses and development. In some cases, particularly those relating to operational-type problems, areas requiring future research were identified, especially screen heat transfer and vibrational effects.

Large-Eddy Simulation of Propeller Crashback

NASA Astrophysics Data System (ADS)

Kumar, Praveen; Mahesh, Krishnan

2013-11-01

Crashback is an operating condition to quickly stop a propelled vehicle, where the propeller is rotated in the reverse direction to yield negative thrust. The crashback condition is dominated by the interaction of free stream flow with strong reverse flow. Crashback causes highly unsteady loads and flow separation on blade surface. This study uses Large-Eddy Simulation to predict the highly unsteady flow field in propeller crashback. Results are shown for a stand-alone open propeller, hull-attached open propeller and a ducted propeller. The simulations are compared to experiment, and used to discuss the essential physics behind the unsteady loads. This work is supported by the Office of Naval Research.

De-aliasing for signal restoration in Propeller MR imaging.

PubMed

Chiu, Su-Chin; Chang, Hing-Chiu; Chu, Mei-Lan; Wu, Ming-Long; Chung, Hsiao-Wen; Lin, Yi-Ru

2017-02-01

Objects falling outside of the true elliptical field-of-view (FOV) in Propeller imaging show unique aliasing artifacts. This study proposes a de-aliasing approach to restore the signal intensities in Propeller images without extra data acquisition. Computer simulation was performed on the Shepp-Logan head phantom deliberately placed obliquely to examine the signal aliasing. In addition, phantom and human imaging experiments were performed using Propeller imaging with various readouts on a 3.0 Tesla MR scanner. De-aliasing using the proposed method was then performed, with the first low-resolution single-blade image used to find out the aliasing patterns in all the single-blade images, followed by standard Propeller reconstruction. The Propeller images without and with de-aliasing were compared. Computer simulations showed signal loss at the image corners along with aliasing artifacts distributed along directions corresponding to the rotational blades, consistent with clinical observations. The proposed de-aliasing operation successfully restored the correct images in both phantom and human experiments. The de-aliasing operation is an effective adjunct to Propeller MR image reconstruction for retrospective restoration of aliased signals. Copyright © 2016 Elsevier Inc. All rights reserved.

Liquefaction and Storage of In-Situ Oxygen on the Surface of Mars

NASA Technical Reports Server (NTRS)

Hauser, Daniel M.; Johnson, Wesley L.; Sutherlin, Steven G.

2016-01-01

The In-Situ production of propellants for Martian and Lunar missions has been heavily discussed since the mid 1990's. One portion of the production of the propellants is the liquefaction, storage, and delivery of the propellants to the stage tanks. Two key technology development efforts are required: large refrigeration systems (cryocoolers) to perform the liquefaction and high performance insulation within a soft vacuum environment. Several different concepts of operation may be employed to liquefy the propellants based on how and where these two technologies are implemented. The concepts that were investigated include: using an accumulator tank to store the propellant until it is needed, liquefying in the flow stream going into the tank, and liquefying in the flight propellant tank itself. The different concept of operations were studied to assess the mass and power impacts of each concept. Additionally, the trade between insulation performance and cryocooler mass was performed to give performance targets for soft vacuum insulation development. It was found that liquefying within the flight propellant tank itself adds the least mass and power requirements to the mission.

Injector design guidelines for gas/liquid propellant systems

NASA Technical Reports Server (NTRS)

Falk, A. Y.; Burick, R. J.

1973-01-01

Injector design guidelines are provided for gas/liquid propellant systems. Information was obtained from a 30-month applied research program encompassing an analytical, design, and experimental effort to relate injector design parameters to simultaneous attainment of high performance and component (injector/thrust chamber) compatibility for gas/liquid space storable propellants. The gas/liquid propellant combination studied was FLOX (82.6% F2)/ ambient temperature gaseous methane. Design criteria that provide for simultaneous attainment of high performance and chamber compatibility are presented for both injector types. Parametric data are presented that are applicable for the design of circular coaxial and like-doublet injectors that operate with design parameters similar to those employed. However, caution should be exercised when applying these data to propellant combinations whose elements operate in ranges considerably different from those employed in this study.

Selected Foreign Counterparts of U.S. Army Ground Combat Systems and Implications for Combat Operations and Modernization

DTIC Science & Technology

2017-01-18

System program. 7 While it may not be realistic to have definitive design criteria for a vehicle to be fielded in 2035 established now, a clearer...is focused on fielding an existing “hard kill” APS capability in the near term for the Army’s M-1 Abrams tanks, M-2/3 Bradley fighting vehicles , and...infantry fighting vehicles (IFVs), tracked self-propelled (SP) artillery, and multiple launch rocket systems (MLRS), which constitutes the nucleus

RL10 ignition limits test for Shuttle Centaur

NASA Technical Reports Server (NTRS)

1987-01-01

During routine development testing of the RL10A-3-3B engine a potential no-ignition condition was encountered when operating at certain propellant inlet conditions within the Shuttle Centaur G operating region. The conditions, the resulting investigative program, and methods to correct the potential problem are discussed. The Shuttle Centaur program was cancelled prior to completion of this effort. Although the RL10 engine in the Atlas Centaur vehicle is required by specification to operate over a wide range of propellant inlet conditions. The vehicle actually operates over a narrow range of conditions. This factor, combined with configuration differences between Atlas Centaur (or Titan Centaur) and the Shuttle Centaur RL10 engines, indicates the ignition problem does not exist for these vehicles. As a precautionary measure the vehicle manufacturer was requested to coordinate with Pratt and Whitney any anticipated changes in propellant inlet conditions from the current narrow range. An engineering change will be proposed for future RL10 deliveries to provide more consistent propellant flow to the igniter. This will permit operation of the engine throughout the wide range specification inlet conditions if desired.

Cooperative Threat Reduction: Cooperation Threat Reduction Program Liquid Propellant Disposition Project

NASA Astrophysics Data System (ADS)

2002-09-01

This audit is one in a series of audits the Deputy Secretary of Defense requested. As part of the Cooperative Threat Reduction (CTR) Program, DoD agreed to assist the Russian Federation in disposing of its liquid rocket propellant. Public Law 102-228 (section 2551 NOTE, title 22, United States Code), the Soviet Nuclear Threat Reduction Act of 1991 designates DoD as the executive agent for the CTR Program. Specific objectives of the act are to destroy chemical, nuclear, and other weapons; transport, store, disable, and safeguard weapons in connection with their destruction; and establish verifiable safeguards against proliferation of weapons of mass destruction. The Office of the Assistant Secretary of Defense (International Security Policy), under the Office of the Under Secretary of Defense for Policy, develops, coordinates, and oversees implementation of policy for the CTR Program. The CTR Directorate, Defense Threat Reduction Agency operates the program.

Advanced high pressure engine study for mixed-mode vehicle applications

NASA Technical Reports Server (NTRS)

Luscher, W. P.; Mellish, J. A.

1977-01-01

High pressure liquid rocket engine design, performance, weight, envelope, and operational characteristics were evaluated for a variety of candidate engines for use in mixed-mode, single-stage-to-orbit applications. Propellant property and performance data were obtained for candidate Mode 1 fuels which included: RP-1, RJ-5, hydrazine, monomethyl-hydrazine, and methane. The common oxidizer was liquid oxygen. Oxygen, the candidate Mode 1 fuels, and hydrogen were evaluated as thrust chamber coolants. Oxygen, methane, and hydrogen were found to be the most viable cooling candidates. Water, lithium, and sodium-potassium were also evaluated as auxiliary coolant systems. Water proved to be the best of these, but the system was heavier than those systems which cooled with the engine propellants. Engine weight and envelope parametric data were established for candidate Mode 1, Mode 2, and dual-fuel engines. Delivered engine performance data were also calculated for all candidate Mode 1 and dual-fuel engines.

Tripropellant engine study

NASA Technical Reports Server (NTRS)

Wheeler, D. B.

1977-01-01

Work conducted was devoted to three main tasks. Thermochemical equilibrium performance data were assembled to establish the expected performance calculations of the mode 1 engine propellant combinations and thermodynamic and transport data for the products of combustion. Turbine drive gas characteristics were also established. Thrust chamber and nozzle cooling studies were devoted to the evaluation of H2, C3H8, CH4, and RP-1 as coolants in the existing SSME cooling circuit geometry. It was found that all these candidate coolants are feasible without limiting the desired operating conditions with the exception of RP-1, which would limit the maximum P(c) to 2000 psia. RP-1 could be used, however, to cool the nozzle only without imposing the chamber pressure limit. A total of 15 candidate engine system cycles were selected and a preliminary engine system balance was conducted for 12 of these systems to establish component operating flowrates, pressures and temperatures. It was found that the staged combustion cycles employing fuel rich LOX/hydrocarbon turbine drive gases are power limited.

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

Code of Federal Regulations, 2010 CFR

2010-10-01

... 46 Shipping 5 2010-10-01 2010-10-01 false Operation of oceangoing non-self-propelled ships Carrying Category D NLS. 151.12-10 Section 151.12-10 Shipping COAST GUARD, DEPARTMENT OF HOMELAND SECURITY (CONTINUED) CERTAIN BULK DANGEROUS CARGOES BARGES CARRYING BULK LIQUID HAZARDOUS MATERIAL CARGOES Equipment and Operating Requirements for Control of...

Transient Analysis of Pressurization and Pneumatic Subsystems of the X-34 Main Propulsion System

NASA Technical Reports Server (NTRS)

Hedayat, A.; Knight, K. C.; Chamption, R. H., Jr.; Kennedy, Jim W. (Technical Monitor)

2000-01-01

Transient models for the pressurization, vent/relief, and pneumatic subsystems of the X-34 Main Propulsion System are presented and simulation of their operation within prescribed requirements are provided. First, using ROCket Engine Transient Simulation (ROCETS) program, pressurization subsystem operation was simulated and helium requirements and the ullage thermodynamic condition within each propellant tank were calculated. Then, Overpressurization scenarios of propellant tanks and the response of vent/relief valves were evaluated using ROCETS simulation of simultaneous operation of the pressurization and vent/relief subsystems by incorporating the valves data into the model. Finally, the ROCETS simulation of in-flight operation of pneumatic subsystem predicted the overall helium consumption, Inter-Propellant Seal (IPS) purge flowrate and thermodynamic conditions, and Spin Start power.

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...

14 CFR 25.907 - Propeller vibration and fatigue.

Code of Federal Regulations, 2010 CFR

2010-01-01

...) The applicant must determine the magnitude of the propeller vibration stresses or loads, including any stress peaks and resonant conditions, throughout the operational envelope of the airplane by either: (1) Measurement of stresses or loads through direct testing or analysis based on direct testing of the propeller...

14 CFR 25.907 - Propeller vibration and fatigue.

Code of Federal Regulations, 2012 CFR

2012-01-01

...) The applicant must determine the magnitude of the propeller vibration stresses or loads, including any stress peaks and resonant conditions, throughout the operational envelope of the airplane by either: (1) Measurement of stresses or loads through direct testing or analysis based on direct testing of the propeller...

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...

14 CFR 25.907 - Propeller vibration and fatigue.

Code of Federal Regulations, 2011 CFR

2011-01-01

...) The applicant must determine the magnitude of the propeller vibration stresses or loads, including any stress peaks and resonant conditions, throughout the operational envelope of the airplane by either: (1) Measurement of stresses or loads through direct testing or analysis based on direct testing of the propeller...

NASA's Evolutionary Xenon Thruster (NEXT) Prototype Model 1R (PM1R) Ion Thruster and Propellant Management System Wear Test Results

NASA Technical Reports Server (NTRS)

VanNoord, Jonathan L.; Soulas, George C.; Sovey, James S.

2010-01-01

The results of the NEXT wear test are presented. This test was conducted with a 36-cm ion engine (designated PM1R) and an engineering model propellant management system. The thruster operated with beam extraction for a total of 1680 hr and processed 30.5 kg of xenon during the wear test, which included performance testing and some operation with an engineering model power processing unit. A total of 1312 hr was accumulated at full power, 277 hr at low power, and the remainder was at intermediate throttle levels. Overall ion engine performance, which includes thrust, thruster input power, specific impulse, and thrust efficiency, was steady with no indications of performance degradation. The propellant management system performed without incident during the wear test. The ion engine and propellant management system were also inspected following the test with no indication of anomalous hardware degradation from operation.

Testing of a Liquid Oxygen/Liquid Methane Reaction Control Thruster in a New Altitude Rocket Engine Test Facility

NASA Technical Reports Server (NTRS)

Meyer, Michael L.; Arrington, Lynn A.; Kleinhenz, Julie E.; Marshall, William M.

2012-01-01

A relocated rocket engine test facility, the Altitude Combustion Stand (ACS), was activated in 2009 at the NASA Glenn Research Center. This facility has the capability to test with a variety of propellants and up to a thrust level of 2000 lbf (8.9 kN) with precise measurement of propellant conditions, propellant flow rates, thrust and altitude conditions. These measurements enable accurate determination of a thruster and/or nozzle s altitude performance for both technology development and flight qualification purposes. In addition the facility was designed to enable efficient test operations to control costs for technology and advanced development projects. A liquid oxygen-liquid methane technology development test program was conducted in the ACS from the fall of 2009 to the fall of 2010. Three test phases were conducted investigating different operational modes and in addition, the project required the complexity of controlling propellant inlet temperatures over an extremely wide range. Despite the challenges of a unique propellant (liquid methane) and wide operating conditions, the facility performed well and delivered up to 24 hot fire tests in a single test day. The resulting data validated the feasibility of utilizing this propellant combination for future deep space applications.

Thrust Deduction in Contrarotating Propellers

DTIC Science & Technology

1974-11-01

nuder gavc At = 0.056 Design CR propellers (Table 2) At = 0.029 Single Screw. Stromn-Tejsen 14 Very good agreement between the experimental and... design experimental points do not lie on the theoretical curve. This is believed to be due to either experimental test accuracy. or tile rudder effect, or...propellers. Con trarotating propellers operating at off- design loading and spacing as well as the contribution of a rudder were investigated. Theli

Evolutionary space station fluids management strategies

NASA Technical Reports Server (NTRS)

1989-01-01

Results are summarized for an 11-month study to define fluid storage and handling strategies and requirements for various specific mission case studies and their associated design impacts on the Space Station. There are a variety of fluid users which require a variety of fluids and use rates. Also, the cryogenic propellants required for NASA's STV, Planetary, and Code Z missions are enormous. The storage methods must accommodate fluids ranging from a high pressure gas or supercritical state fluid to a sub-cooled liquid (and superfluid helium). These requirements begin in the year 1994, reach a maximum of nearly 1800 metric tons in the year 2004, and trail off to the year 2018, as currently planned. It is conceivable that the cryogenic propellant needs for the STV and/or Lunar mission models will be met by LTCSF LH2/LO2 tanksets attached to the SS truss structure. Concepts and corresponding transfer and delivery operations have been presented for STV propellant provisioning from the SS. A growth orbit maneuvering vehicle (OMV) and associated servicing capability will be required to move tanksets from delivery launch vehicles to the SS or co-orbiting platforms. Also, appropriate changes to the software used for OMV operation are necessary to allow for the combined operation of the growth OMV. To support fluid management activities at the Space Station for the experimental payloads and propellant provisioning, there must be truss structure space allocated for fluid carriers and propellant tanksets, and substantial beam strengthening may be required. The Station must have two Mobile Remote Manipulator Systems (MRMS) and the growth OMV propellant handling operations for the STV at the SS. Propellant needs for the Planetary Initiatives and Code Z mission models will most likely be provided by co-orbiting propellant platform(s). Space Station impacts for Code Z mission fluid management activities will be minimal.

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.

78 FR 45471 - Airworthiness Directives; Schempp-Hirth Flugzeugbau GmbH Gliders

Federal Register 2010, 2011, 2012, 2013, 2014

2013-07-29

.... The MCAI describes the unsafe condition as the instructions provided to inspect the propeller hub and...: It was found that the accomplishment instructions provided to check the powered sailplane's propeller... propeller hub and blades are insufficient for detecting cracks and/or other damage, and other operating...

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...

14 CFR 420.69 - Solid and liquid propellants located together.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 4 2011-01-01 2011-01-01 false Solid and liquid propellants located together. 420.69 Section 420.69 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION... Licensee § 420.69 Solid and liquid propellants located together. (a) A launch site operator proposing an...

14 CFR 420.69 - Solid and liquid propellants located together.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 4 2012-01-01 2012-01-01 false Solid and liquid propellants located together. 420.69 Section 420.69 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION... Licensee § 420.69 Solid and liquid propellants located together. (a) A launch site operator proposing an...

14 CFR 420.69 - Solid and liquid propellants located together.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 4 2010-01-01 2010-01-01 false Solid and liquid propellants located together. 420.69 Section 420.69 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION... Licensee § 420.69 Solid and liquid propellants located together. (a) A launch site operator proposing an...

46 CFR 69.121 - Engine room deduction.

Code of Federal Regulations, 2011 CFR

2011-10-01

... necessary for the safe operation and maintenance of the propelling machinery, the entire space, or, if... machinery space is not bulkheaded off or is larger than necessary for the safe operation and maintenance of... room deduction is either a percentage of the vessel's total propelling machinery spaces or a percentage...

Lessons Learned with Metallized Gelled Propellants

NASA Technical Reports Server (NTRS)

1996-01-01

During testing of metallized gelled propellants in a rocket engine, many changes had to be made to the normal test program for traditional liquid propellants. The lessons learned during the testing and the solutions for many of the new operational conditions posed with gelled fuels will help future programs run more smoothly. The major factors that influenced the success of the testing were propellant settling, piston-cylinder tank operation, control of self pressurization, capture of metal oxide particles, and a gelled-fuel protective layer. In these ongoing rocket combustion experiments at the NASA Lewis Research Center, metallized, gelled liquid propellants are used in a small modular engine that produces 30 to 40 lb of thrust. Traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt% loadings of aluminum are used with gaseous oxygen as the oxidizer. The figure compares the thrust chamber efficiencies of different engines.

RANS simulation of cavitation and hull pressure fluctuation for marine propeller operating behind-hull condition

NASA Astrophysics Data System (ADS)

Paik, Kwang-Jun; Park, Hyung-Gil; Seo, Jongsoo

2013-12-01

Simulations of cavitation flow and hull pressure fluctuation for a marine propeller operating behind a hull using the unsteady Reynolds-Averaged Navier-Stokes equations (RANS) are presented. A full hull body submerged under the free surface is modeled in the computational domain to simulate directly the wake field of the ship at the propeller plane. Simulations are performed in design and ballast draught conditions to study the effect of cavitation number. And two propellers with slightly different geometry are simulated to validate the detectability of the numerical simulation. All simulations are performed using a commercial CFD software FLUENT. Cavitation patterns of the simulations show good agreement with the experimental results carried out in Samsung CAvitation Tunnel (SCAT). The simulation results for the hull pressure fluctuation induced by a propeller are also compared with the experimental results showing good agreement in the tendency and amplitude, especially, for the first blade frequency.

Numerical study of fairing installed between brackets based on CFD

NASA Astrophysics Data System (ADS)

Xi, Peng; Xiong, Ying; Tang, Xin

2017-10-01

In view of the low speed and instability of the flow between the two arms of the bracket in front of the propeller, the fairing is installed between the arms of the bracket taking example of compensating duct, in order to speed up the flow between the bracket arms and improve the flow quality. A four-propeller surface ship was studied and an integral mathematic model including hull, appendage and propellers was established. Using a RANS solver, its installation height, angle and airfoil is optimized. Then ship models with fairing and without fairing are calculated. The result shows that fairing improves propeller efficiency behind ship with 1.1% of the outer propeller and 1.6% of the inner propeller, which indicates that fairing helps improve the flow quality

Method for providing real-time control of a gaseous propellant rocket propulsion system

NASA Technical Reports Server (NTRS)

Morris, Brian G. (Inventor)

1991-01-01

The new and improved methods and apparatus disclosed provide effective real-time management of a spacecraft rocket engine powered by gaseous propellants. Real-time measurements representative of the engine performance are compared with predetermined standards to selectively control the supply of propellants to the engine for optimizing its performance as well as efficiently managing the consumption of propellants. A priority system is provided for achieving effective real-time management of the propulsion system by first regulating the propellants to keep the engine operating at an efficient level and thereafter regulating the consumption ratio of the propellants. A lower priority level is provided to balance the consumption of the propellants so significant quantities of unexpended propellants will not be left over at the end of the scheduled mission of the engine.

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.

Water Contaminant Mitigation in Ionic Liquid Propellant

NASA Technical Reports Server (NTRS)

Conroy, David; Ziemer, John

2009-01-01

Appropriate system and operational requirements are needed in order to ensure mission success without unnecessary cost. Purity requirements applied to thruster propellants may flow down to materials and operations as well as the propellant preparation itself. Colloid electrospray thrusters function by applying a large potential to a room temperature liquid propellant (such as an ionic liquid), inducing formation of a Taylor cone. Ions and droplets are ejected from the Taylor cone and accelerated through a strong electric field. Electrospray thrusters are highly efficient, precise, scaleable, and demonstrate low thrust noise. Ionic liquid propellants have excellent properties for use as electrospray propellants, but can be hampered by impurities, owing to their solvent capabilities. Of foremost concern is the water content, which can result from exposure to atmosphere. Even hydrophobic ionic liquids have been shown to absorb water from the air. In order to mitigate the risks of bubble formation in feed systems caused by water content of the ionic liquid propellant, physical properties of the ionic liquid EMI-Im are analyzed. The effects of surface tension, material wetting, physisorption, and geometric details of the flow manifold and electrospray emitters are explored. Results are compared to laboratory test data.

Solid propellant processing factor in rocket motor design

NASA Technical Reports Server (NTRS)

1971-01-01

The ways are described by which propellant processing is affected by choices made in designing rocket engines. Tradeoff studies, design proof or scaleup studies, and special design features are presented that are required to obtain high product quality, and optimum processing costs. Processing is considered to include the operational steps involved with the lining and preparation of the motor case for the grain; the procurement of propellant raw materials; and propellant mixing, casting or extrusion, curing, machining, and finishing. The design criteria, recommended practices, and propellant formulations are included.

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.

The design of propeller blade roots

NASA Technical Reports Server (NTRS)

Cordes, G

1942-01-01

Predicated on the assumption of certain normal conditions for engine and propeller, simple expressions for the static and dynamic stresses of propeller blade roots are evolved. They, in combination with the fatigue strength diagram of the employed material, afford for each engine power one certain operating point by which the state of stress serving as a basis for the design of the root is defined. Different stress cases must be analyzed, depending on the vibration tendency of engine and use of propeller. The solution affords an insight into the possible introduction of different size classes of propeller.

Aeroelastic analysis for propellers - mathematical formulations and program user's manual

NASA Technical Reports Server (NTRS)

Bielawa, R. L.; Johnson, S. A.; Chi, R. M.; Gangwani, S. T.

1983-01-01

Mathematical development is presented for a specialized propeller dedicated version of the G400 rotor aeroelastic analysis. The G400PROP analysis simulates aeroelastic characteristics particular to propellers such as structural sweep, aerodynamic sweep and high subsonic unsteady airloads (both stalled and unstalled). Formulations are presented for these expanded propeller related methodologies. Results of limited application of the analysis to realistic blade configurations and operating conditions which include stable and unstable stall flutter test conditions are given. Sections included for enhanced program user efficiency and expanded utilization include descriptions of: (1) the structuring of the G400PROP FORTRAN coding; (2) the required input data; and (3) the output results. General information to facilitate operation and improve efficiency is also provided.

Bleed cycle propellant pumping in a gas-core nuclear rocket engine system

NASA Technical Reports Server (NTRS)

Kascak, A. F.; Easley, A. J.

1972-01-01

The performance of ideal and real staged primary propellant pumps and bleed-powered turbines was calculated for gas-core nuclear rocket engines over a range of operating pressures from 500 to 5000 atm. This study showed that for a required engine operating pressure of 1000 atm the pump work was about 0.8 hp/(lb/sec), the specific impulse penalty resulting from the turbine propellant bleed flow as about 10 percent; and the heat required to preheat the propellant was about 7.8 MN/(lb/sec). For a specific impulse above 2400 sec, there is an excess of energy available in the moderator due to the gamma and neutron heating that occurs there. Possible alternative pumping cycles are the Rankine or Brayton cycles.

Satellite Power Systems (SPS) concept definition study. Volume 5: Special emphasis studies. [rectenna and solar power satellite design studies

NASA Technical Reports Server (NTRS)

Hanley, G. M.

1980-01-01

Satellite configurations based on the Satellite Power System baseline requirements were analyzed and a preferred concept selected. A satellite construction base was defined, precursor operations incident to establishment of orbital support facilities identified, and the satellite construction sequence and procedures developed. Rectenna construction requirement were also addressed. Mass flow to orbit requirements were revised and traffic models established based on construction of 60 instead of 120 satellites. Analyses were conducted to determine satellite control, resources, manufacturing, and propellant requirements. The impact of the laser beam used for space-to-Earth power transmission upon the intervening atmosphere was examined as well as the inverse effect. The significant space environments and their effects on spacecraft components were investigated to define the design and operational limits imposed by the environments on an orbit transfer vehicle. The results show that LEO altitude 300 nmi and transfer orbit duration 6 months are preferrable.

Induction simulation of gas core nuclear engine

NASA Technical Reports Server (NTRS)

Poole, J. W.; Vogel, C. E.

1973-01-01

The design, construction and operation of an induction heated plasma device known as a combined principles simulator is discussed. This device incorporates the major design features of the gas core nuclear rocket engine such as solid feed, propellant seeding, propellant injection through the walls, and a transpiration cooled, choked flow nozzle. Both argon and nitrogen were used as propellant simulating material, and sodium was used for fuel simulating material. In addition, a number of experiments were conducted utilizing depleted uranium as the fuel. The test program revealed that satisfactory operation of this device can be accomplished over a range of operating conditions and provided additional data to confirm the validity of the gas core concept.

Orbiting propellent depot safety. Volume 2: Technical discussion

NASA Technical Reports Server (NTRS)

1971-01-01

The results of a study assessing the gross requirements and concomitant safety hazards associated with the operation of several configurations of orbiting propellant depots (OPD) are presented. The major structural and operational aspects of the integral, semimodular, and modular concepts are described. The concepts are evaluated to determine their safety hazards, and emphasis is placed on propellant transfer to and from the OPD. The study indicates that a modular mode of depot resupply is desirable regardless of the OPD configuration. This resupply mode requires no flow transfer to resupply either the semimodular or modular OPD. Of the three concepts, the semimodular appears to provide the best operational advantage and lowest safety risk.

LES of propeller crashback

NASA Astrophysics Data System (ADS)

Kumar, Praveen; Mahesh, Krishnan

2014-11-01

Crashback is an operating condition to quickly stop a propelled vehicle, where the propeller is rotated in the reverse direction to yield a negative thrust. In crashback, the freestream interacts with the strong reverse flow from the propeller leading to massive flow separation and highly unsteady loads. We have used Large-Eddy Simulation (LES) in recent years to accurately simulate the flowfield in crashback around a stand-alone open propeller, hull-attached (posterior alone) open propeller and a ducted propeller with stator blades. This talk will discuss our work towards LES of crashback inclusive of the entire hull. The results will be compared to available experimental data, and the flow physics will be discussed. This work is supported by the Office of Naval Research.

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.

Characteristics of Five Propellers in Flight

NASA Technical Reports Server (NTRS)

Crowley, J W , Jr; Mixson, R E

1928-01-01

This investigation was made for the purpose of determining the characteristics of five full-scale propellers in flight. The equipment consisted of five propellers in conjunction with a VE-7 airplane and a Wright E-2 engine. The propellers were of the same diameter and aspect ratio. Four of them differed uniformly in thickness and pitch and the fifth propeller was identical with one of the other four with exception of a change of the airfoil section. The propeller efficiencies measured in flight are found to be consistently lower than those obtained in model tests. It is probable that this is mainly a result of the higher tip speeds used in the full-scale tests. The results show also that because of differences in propeller deflections it is difficult to obtain accurate comparisons of propeller characteristics. From this it is concluded that for accurate comparisons it is necessary to know the propeller pitch angles under actual operating conditions. (author)

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.

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.

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.

Holographic investigation of solid propellant particulates

NASA Astrophysics Data System (ADS)

Gillespie, T. R.

1981-12-01

The investigation completed the development process to establish a technique to obtain holographic recordings of particulate behavior during the combustion process of solid propellants in a two-dimensional rocket motor. Holographic and photographic recordings were taken in a crossflow environment using various compositions of metallized propellants. The reconstructed holograms are used to provide data on the behavior of aluminum/aluminum oxide particulates in a steady state combustion environment as a function of the initial aluminum size cast into the propellant. High speed, high resolution motion pictures were taken to compare the cinematic data with that available from the holograms.

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.

Recommendations for Safe Separation Distances from the Kennedy Space Center (KSC) Vehicle Assembly Building (VAB) Using a Heat-Flux-Based Analytical Approach (Abridged)

NASA Technical Reports Server (NTRS)

Cragg, Clinton H.; Bowman, Howard; Wilson, John E.

2011-01-01

The NASA Engineering and Safety Center (NESC) was requested to provide computational modeling to support the establishment of a safe separation distance surrounding the Kennedy Space Center (KSC) Vehicle Assembly Building (VAB). The two major objectives of the study were 1) establish a methodology based on thermal flux to determine safe separation distances from the Kennedy Space Center's (KSC's) Vehicle Assembly Building (VAB) with large numbers of solid propellant boosters containing hazard division 1.3 classification propellants, in case of inadvertent ignition; and 2) apply this methodology to the consideration of housing eight 5-segment solid propellant boosters in the VAB. The results of the study are contained in this report.

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.

Tests of Five Full-Scale Propellers in the Presence of a Radial and a Liquid-Cooled Engine Nacelle, Including Tests of Two Spinners

NASA Technical Reports Server (NTRS)

Biermann, David; Hartman, Edwin P

1938-01-01

Wind-tunnel tests are reported of five 3-blade 10-foot propellers operating in front of a radial and a liquid-cooled engine nacelle. The range of blade angles investigated extended from 15 degrees to 45 degrees. Two spinners were tested in conjunction with the liquid-cooled engine nacelle. Comparisons are made between propellers having different blade-shank shapes, blades of different thickness, and different airfoil sections. The results show that propellers operating in front of the liquid-cooled engine nacelle had higher take-off efficiencies than when operating in front of the radial engine nacelle; the peak efficiency was higher only when spinners were employed. One spinner increased the propulsive efficiency of the liquid-cooled unit 6 percent for the highest blade-angle setting investigated and less for lower blade angles. The propeller having airfoil sections extending into the hub was superior to one having round blade shanks. The thick propeller having a Clark y section had a higher take-off efficiency than the thinner one, but its maximum efficiency was possibly lower. Of the three blade sections tested, Clark y, R.A.F. 6, and NACA 2400-34, the Clark y was superior for the high-speed condition, but the R.A.F. 6 excelled for the take-off condition.

Runtime and Pressurization Analyses of Propellant Tanks

NASA Technical Reports Server (NTRS)

Field, Robert E.; Ryan, Harry M.; Ahuja, Vineet; Hosangadi, Ashvin; Lee, Chung P.

2007-01-01

Multi-element unstructured CFD has been utilized at NASA SSC to carry out analyses of propellant tank systems in different modes of operation. The three regimes of interest at SSC include (a) tank chill down (b) tank pressurization and (c) runtime propellant draw-down and purge. While tank chill down is an important event that is best addressed with long time-scale heat transfer calculations, CFD can play a critical role in the tank pressurization and runtime modes of operation. In these situations, problems with contamination of the propellant by inclusion of the pressurant gas from the ullage causes a deterioration of the quality of the propellant delivered to the test article. CFD can be used to help quantify the mixing and propellant degradation. During tank pressurization under some circumstances, rapid mixing of relatively warm pressurant gas with cryogenic propellant can lead to rapid densification of the gas and loss of pressure in the tank. This phenomenon can cause serious problems during testing because of the resulting decrease in propellant flow rate. With proper physical models implemented, CFD can model the coupling between the propellant and pressurant including heat transfer and phase change effects and accurately capture the complex physics in the evolving flowfields. This holds the promise of allowing the specification of operational conditions and procedures that could minimize the undesirable mixing and heat transfer inherent in propellant tank operation. It should be noted that traditional CFD modeling is inadequate for such simulations because the fluids in the tank are in a range of different sub-critical and supercritical states and elaborate phase change and mixing rules have to be developed to accurately model the interaction between the ullage gas and the propellant. We show a typical run-time simulation of a spherical propellant tank, containing RP-1 in this case, being pressurized with room-temperature nitrogen at 540 R. Nitrogen, shown in blue on the right-hand side of the figures, enters the tank from the diffuser at the top of the figures and impinges on the RP-1, shown in red, while the propellant is being continuously drained at the rate of 1050 lbs/sec through a pipe at the bottom of the tank. The sequence of frames in Figure 1 shows the resultant velocity fields and mixing between nitrogen and RP-1 in a cross-section of the tank at different times. A vortex is seen to form in the incoming nitrogen stream that tends to entrain propellant, mixing it with the pressurant gas. The RP-1 mass fraction contours in Figure 1 are also indicative of the level of mixing and contamination of the propellant. The simulation is used to track the propagation of the pure propellant front as it is drawn toward the exit with the evolution of the mixing processes in the tank. The CFD simulation modeled a total of 10 seconds of run time. As is seen from Figure 1d, after 5.65 seconds the propellant front is nearing the drain pipe, especially near the center of the tank. Behind this pure propellant front is a mixed fluid of compromised quality that would require the test to end when it reaches the exit pipe. Such unsteady simulations provide an estimate of the time that a high-quality propellant supply to the test article can be guaranteed at the modeled mass flow rate. In the final paper, we will discuss simulations of the LOX and propellant tanks at NASA SSC being pressurized by an inert ullage. Detailed comparisons will be made between the CFD simulations and lower order models as well as with test data. Conditions leading to cryo collapse in the tank will also be identified.

Water-propellant resistojets for man-tended platforms

NASA Technical Reports Server (NTRS)

Louviere, Allen J.; Jones, Robert E.; Morren, W. Earl; Sovey, James S.

1987-01-01

The selection of a propulsion system for a man-tended platform has been influenced by the planned use of resistojets for drag make-up on the manned space station. For that application a resistojet has been designed that is capable of operation with a wide variety of propellants, including water. The reasons for the selection of water as the propellant and the performance of water as a propellant are discussed. The man-tended platform and its mission requirements are described.

ISRU Propellant Selection for Space Exploration Vehicles

NASA Technical Reports Server (NTRS)

Chen, Timothy T.

2013-01-01

Chemical propulsion remains the only viable solution as technically matured technology for the near term human space transportation to Lunar and Mars. Current mode of space travel requires us to "take everything we will need", including propellant for the return trip. Forcing the mission designers to carry propellant for the return trip limits payload mass available for mission operations and results in a large and costly (and often unaffordable) design. Producing propellant via In-Situ Resource Utilization (ISRU) will enable missions with chemical propulsion by the "refueling" of return-trip propellant. It will reduce vehicle propellant mass carrying requirement by over 50%. This mass reduction can translates into increased payload to enhance greater mission capability, reduces vehicle size, weight and cost. It will also reduce size of launch vehicle fairing size as well as number of launches for a given space mission and enables exploration missions with existing chemical propulsion. Mars remains the ultimate destination for Human Space Exploration within the Solar System. The Mars atmospheric consist of 95% carbon dioxide (CO2) and the presence of Ice (water) was detected on Mars surfaces. This presents a basic chemical building block for the ISRU propellant manufacturing. However, the rationale for the right propellant to produce via ISRU appears to be limited to the perception of "what we can produce" as oppose to "what is the right propellant". Methane (CH4) is often quoted as a logical choice for Mars ISRU propellant, however; it is believed that there are better alternatives available that can result in a better space transportation architecture. A system analysis is needed to determine on what is the right propellant choice for the exploration vehicle. This paper examines the propellant selection for production via ISRU method on Mars surfaces. It will examine propellant trades for the exploration vehicle with resulting impact on vehicle performance, size, and on launch vehicles. It will investigate propellant manufacturing techniques that will be applicable on Mars surfaces and address related issues on storage, transfer, and safety. Finally, it will also address the operability issues associated with the impact of propellant selection on ground processing and launch vehicle integration.

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.

Numerical prediction of marine propeller noise in non-uniform inflow

NASA Astrophysics Data System (ADS)

Pan, Yu-cun; Zhang, Huai-xin

2013-03-01

A numerical study on the acoustic radiation of a propeller interacting with non-uniform inflow has been conducted. Real geometry of a marine propeller DTMB 4118 is used in the calculation, and sliding mesh technique is adopted to deal with the rotational motion of the propeller. The performance of the DES (Detached Eddy Simulation) approach at capturing the unsteady forces and moments on the propeller is compared with experiment. Far-field sound radiation is predicted by the formation 1A developed by Farassat, an integral solution of FW-H (Ffowcs Williams-Hawkings) equation in time domain. The sound pressure and directivity patterns of the propeller operating in two specific velocity distributions are discussed.

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.

AP reclamation and reuse in RSRM propellant

NASA Technical Reports Server (NTRS)

Miks, Kathryn F.; Harris, Stacey A.

1995-01-01

A solid propellant ingredient reclamation pilot plant has been evaluated at the Strategic Operations of Thiokol Corporation, located in Brigham City, Utah. The plant produces AP wet cake (95 percent AP, 5 percent water) for recycling at AP vendors. AP has been obtained from two standard propellant binder systems (PBAN and HTPB). Analytical work conducted at Thiokol indicates that the vendor-recrystallized AP meets Space Shuttle propellant specification requirements. Thiokol has processed 1-, 5-, and 600-gallon propellant mixes with the recrystallized AP. Processing, cast, cure, ballistic, mechanical, and safety properties have been evaluated. Phillips Laboratory static-test-fired 70-pound and 800-pound BATES motors. The data indicate that propellant processed with reclaimed AP has nominal properties.

75 FR 922 - Notification and Reporting of Aircraft Accidents or Incidents and Overdue Aircraft, and...

Federal Register 2010, 2011, 2012, 2013, 2014

2010-01-07

... pose to crews, passengers, and bystanders. However, the NTSB notes that propeller blades are designed... intact and in place during normal operation. Propeller blades are not designed or expected to continue to... release of all or a portion of a propeller blade from an aircraft, inconsistent with its design parameters...

Flow-field Survey of an Empennage Wake Interacting with a Pusher Propeller

NASA Technical Reports Server (NTRS)

Horne, W. Clifton; Soderman, Paul T.

1988-01-01

The flow field between a model empennage and a 591-mm-diameter pusher propeller was studied in the Ames 7- by 10-Foot Wind Tunnel with directional pressure probes and hot-wire anemometers. The region probed was bounded by the empennage trailing edge and downstream propeller. The wake properties, including effects of propeller operation on the empennage wake, were investigated for two empennage geometries: one, a vertical tail fin, the other, a Y-tail with a 34 deg dihedral. Results showed that the effect of the propeller on the empennage wake upstream of the propeller was not strong. The flow upstream of the propeller was accelerated in the streamwise direction by the propeller, but the empennage wake width and velocity defect were relatively unaffected by the presence of the propeller. The peak turbulence in the wake near the propeller tip station, 0.66 diameter behind the vertical tail fin, was approximately 3 percent of the free-stream velocity. The velocity field data can be used in predictions of the acoustic field due to propeller-wake interaction.

Liquid-hydrogen rocket engine development at Aerojet, 1944 - 1950

NASA Technical Reports Server (NTRS)

Osborn, G. H.; Gordon, R.; Coplen, H. L.; James, G. S.

1977-01-01

This program demonstrated the feasibility of virtually all the components in present-day, high-energy, liquid-rocket engines. Transpiration and film-cooled thrust chambers were successfully operated. The first liquid-hydrogen tests of the coaxial injector was conducted and the first pump to successfully produce high pressures in pumping liquid hydrogen was tested. A 1,000-lb-thrust gaseous propellant and a 3,000-lb-thrust liquid-propellant thrust chamber were operated satisfactorily. Also, the first tests were conducted to evaluate the effects of jet overexpansion and separation on performance of rocket thrust chambers with hydrogen-oxygen propellants.

The Advancing State of AF-M315E Technology

NASA Technical Reports Server (NTRS)

Masse, Robert; Spores, Ronald A.; McLean, Chris

2014-01-01

The culmination of twenty years of applied research in hydroxyl ammonium nitrate (HAN)-based monopropellants, the NASA Space Technology mission Directorate's (STMD) Green Propellant Infusion Mission (GPIM) will achieve the first on-orbit demonstration of an operational AF-M315E green propellant propulsion system by the end of 2015. Following an contextual overview of the completed flight design of the GPIM propellant storage and feed system, results of first operation of a flight-representative heavyweight 20-N engineering model thruster (to be conducted in mid-2014) are presented with performance comparisons to prior lab model (heavyweight) test articles.

Cryogenic propellant management: Integration of design, performance and operational requirements

NASA Technical Reports Server (NTRS)

Worlund, A. L.; Jamieson, J. R., Jr.; Cole, T. W.; Lak, T. I.

1985-01-01

The integration of the design features of the Shuttle elements into a cryogenic propellant management system is described. The implementation and verification of the design/operational changes resulting from design deficiencies and/or element incompatibilities encountered subsequent to the critical design reviews are emphasized. Major topics include: subsystem designs to provide liquid oxygen (LO2) tank pressure stabilization, LO2 facility vent for ice prevention, liquid hydrogen (LH2) feedline high point bleed, pogo suppression on the Space Shuttle Main Engine (SSME), LO2 low level cutoff, Orbiter/engine propellant dump, and LO2 main feedline helium injection for geyser prevention.

Solar-Thermal Engine Testing

NASA Technical Reports Server (NTRS)

Tucker, Stephen; Salvail, Pat; Haynes, Davy (Technical Monitor)

2001-01-01

A solar-thermal engine serves as a high-temperature solar-radiation absorber, heat exchanger, and rocket nozzle. collecting concentrated solar radiation into an absorber cavity and transferring this energy to a propellant as heat. Propellant gas can be heated to temperatures approaching 4,500 F and expanded in a rocket nozzle, creating low thrust with a high specific impulse (I(sub sp)). The Shooting Star Experiment (SSE) solar-thermal engine is made of 100 percent chemical vapor deposited (CVD) rhenium. The engine 'module' consists of an engine assembly, propellant feedline, engine support structure, thermal insulation, and instrumentation. Engine thermal performance tests consist of a series of high-temperature thermal cycles intended to characterize the propulsive performance of the engines and the thermal effectiveness of the engine support structure and insulation system. A silicone-carbide electrical resistance heater, placed inside the inner shell, substitutes for solar radiation and heats the engine. Although the preferred propellant is hydrogen, the propellant used in these tests is gaseous nitrogen. Because rhenium oxidizes at elevated temperatures, the tests are performed in a vacuum chamber. Test data will include transient and steady state temperatures on selected engine surfaces, propellant pressures and flow rates, and engine thrust levels. The engine propellant-feed system is designed to Supply GN2 to the engine at a constant inlet pressure of 60 psia, producing a near-constant thrust of 1.0 lb. Gaseous hydrogen will be used in subsequent tests. The propellant flow rate decreases with increasing propellant temperature, while maintaining constant thrust, increasing engine I(sub sp). In conjunction with analytical models of the heat exchanger, the temperature data will provide insight into the effectiveness of the insulation system, the structural support system, and the overall engine performance. These tests also provide experience on operational aspects of the engine and associated subsystems, and will include independent variation of both steady slate heat-exchanger temperature prior to thrust operation and nitrogen inlet pressure (flow rate) during thrust operation. Although the Shooting Star engines were designed as thermal-storage engines to accommodate mission parameters, they are fully capable of operating as scalable, direct-gain engines. Tests are conducted in both operational modes. Engine thrust and propellant flow rate will be measured and thereby I(sub sp). The objective of these tests is to investigate the effectiveness of the solar engine as a heat exchanger and a rocket. Of particular interest is the effectiveness of the support structure as a thermal insulator, the integrity of both the insulation system and the insulation containment system, the overall temperature distribution throughout the engine module, and the thermal power required to sustain steady state fluid temperatures at various flow rates.

14 CFR 121.645 - Fuel supply: Turbine-engine powered airplanes, other than turbo propeller: Flag and supplemental...

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 3 2013-01-01 2013-01-01 false Fuel supply: Turbine-engine powered airplanes, other than turbo propeller: Flag and supplemental operations. 121.645 Section 121.645 Aeronautics... SUPPLEMENTAL OPERATIONS Dispatching and Flight Release Rules § 121.645 Fuel supply: Turbine-engine powered...

14 CFR 121.645 - Fuel supply: Turbine-engine powered airplanes, other than turbo propeller: Flag and supplemental...

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 3 2014-01-01 2014-01-01 false Fuel supply: Turbine-engine powered airplanes, other than turbo propeller: Flag and supplemental operations. 121.645 Section 121.645 Aeronautics... SUPPLEMENTAL OPERATIONS Dispatching and Flight Release Rules § 121.645 Fuel supply: Turbine-engine powered...

14 CFR 121.645 - Fuel supply: Turbine-engine powered airplanes, other than turbo propeller: Flag and supplemental...

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 3 2012-01-01 2012-01-01 false Fuel supply: Turbine-engine powered airplanes, other than turbo propeller: Flag and supplemental operations. 121.645 Section 121.645 Aeronautics... SUPPLEMENTAL OPERATIONS Dispatching and Flight Release Rules § 121.645 Fuel supply: Turbine-engine powered...

Tripropellant engine study

NASA Technical Reports Server (NTRS)

Wheeler, D. B.

1978-01-01

Engine performance data, combustion gas thermodynamic properties, and turbine gas parameters were determined for various high power cycle engine configurations derived from the space shuttle main engine that will allow sequential burning of LOX/hydrocarbon and LOX/hydrogen fuels. Both stage combustion and gas generator pump power cycles were considered. Engine concepts were formulated for LOX/RP-1, LOX/CH4, and LOX/C3H8 propellants. Flowrates and operating conditions were established for this initial set of engine systems, and the adaptability of the major components of shuttle main engine was investigated.

Large-eddy simulation of propeller wake at design operating conditions

NASA Astrophysics Data System (ADS)

Kumar, Praveen; Mahesh, Krishnan

2016-11-01

Understanding the propeller wake is crucial for efficient design and optimized performance. The dynamics of the propeller wake are also central to physical phenomena such as cavitation and acoustics. Large-eddy simulation is used to study the evolution of the wake of a five-bladed marine propeller from near to far field at design operating condition. The computed mean loads and phase-averaged flow field show good agreement with experiments. The propeller wake consisting of tip and hub vortices undergoes streamtube contraction, which is followed by the onset of instabilities as evident from the oscillations of the tip vortices. Simulation results reveal a mutual induction mechanism of instability where instead of the tip vortices interacting among themselves, they interact with the smaller vortices generated by the roll-up of the blade trailing edge wake in the near wake. Phase-averaged and ensemble-averaged flow fields are analyzed to explain the flow physics. This work is supported by ONR.

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.

An unsteady Euler scheme for the analysis of ducted propellers

NASA Technical Reports Server (NTRS)

Srivastava, R.

1992-01-01

An efficient unsteady solution procedure has been developed for analyzing inviscid unsteady flow past ducted propeller configurations. This scheme is first order accurate in time and second order accurate in space. The solution procedure has been applied to a ducted propeller consisting of an 8-bladed SR7 propeller with a duct of NACA 0003 airfoil cross section around it, operating in a steady axisymmetric flowfield. The variation of elemental blade loading with radius, compares well with other published numerical results.

Comparative evaluation of gas-turbine engine combustion chamber starting and stalling characteristics for mechanical and air-injection

NASA Technical Reports Server (NTRS)

Dyatlov, I. N.

1983-01-01

The effectiveness of propellant atomization with and without air injection in the combustion chamber nozzle of a gas turbine engine is studied. Test show that the startup and burning performance of these combustion chambers can be improved by using an injection during the mechanical propellant atomization process. It is shown that the operational range of combustion chambers can be extended to poorer propellant mixtures by combined air injection mechanical atomization of the propellant.

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.

Evaluation of the Langley 4- by 7-meter tunnel for propeller noise measurements

NASA Technical Reports Server (NTRS)

Block, P. J. W.; Gentry, G. L., Jr.

1984-01-01

An experimental and theoretical evaluation of the Langley 4- by 7- Meter Tunnel was conducted to determine its suitability for obtaining propeller noise data. The tunnel circuit and open test section are described. An experimental evaluation is performed using microphones placed in and on the tunnel floor. The reflection characteristics and background noise are determined. The predicted source (propeller) near-field/far-field boundary is given using a first-principles method. The effect of the tunnel-floor boundry layer on the noise from the propeller is also predicted. A propeller test stand used for part of his evaluation is also described. The measured propeller performance characteristics are compared with those obtained at a larger scale, and the effect of the test-section configuration on the propeller performance is examined. Finally, propeller noise measurements were obtained on an eight-bladed SR-2 propeller operating at angles of attack -8 deg, 0 deg, and 4.6 deg to give an indication of attainable signal-to-noise ratios.

Cavitation erosion in blocked flow with a ducted ice-class propeller

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

Doucet, J.M.; Bose, N.; Walker, D.

1996-12-31

Ships that operate in ice often encounter momentary increased propeller cavitation because ice pieces block the flow into the propeller. For ducted propellers, this additional cavitation is more significant than it is for open propellers; ice pieces may become lodged against and within the duct and subject the propeller to longer periods of increased cavitation due to the blocked flow. Associated with this blocked flow is the possibility of cavitation erosion on the propeller. An erosion study, using paint films, was conducted in a cavitation tunnel with a model propeller of the type fitted to the Canadian Marine Drilling Ltd.more » vessel MV Robert LeMeur. A simulated ice blockage was installed ahead of the propeller model and within the duct. Tests were carried out over a range of advance coefficients for various test conditions. The resulting types of cavitation were documented, the erosion patterns were photographed and comparisons between each test were made.« less

Green Applications for Space Power Project

NASA Technical Reports Server (NTRS)

Robinson, Joel (Principal Investigator)

2014-01-01

Spacecraft propulsion and power for many decades has relied on Hydrazine monopropellant technology for auxiliary power units (APU), orbital circularization, orbit raising/lowering and attitude control. However, Hydrazine is toxic and therefore requires special ground handling procedures to ensure launch crew safety. The Swedish Company ECAPS has developed a technology based upon the propellant Ammonium Dinitramide (ADN) that offers higher performance, higher density and reduced ground handling support than Hydrazine. This blended propellant is called LMP-103S. Currently, the United States Air Force (USAF) is pursuing a technology based on Hydroxyl Ammonium Nitrate (HAN, otherwise known as AF-M315E) with industry partners Aerojet and Moog. Based on the advantages offered by these propellants, MSFC should explore powering APU's with these propellants. Due to the availability of space hardware, the principal investigator has found a collection of USAF hardware, that will act as a surrogate, which operates on a Hydrazine derivative. The F-16 fighter jet uses H-70 or 30% diluted Hydrazine for an Emergency Power Unit (EPU) which supplies power to the plane. The PI has acquired two EPU's from planes slated for destruction at the Davis Monthan AFB. This CIF will include a partnership with 2 other NASA Centers who are individually seeking seed funds from their respective organizations: Kennedy Space Center (KSC) and Dryden Flight Research Center (DFRC). KSC is preparing for future flights from their launch pads that will utilize green propellants and desire a low-cost testbed in which to test and calibrate new leak detection sensors. DFRC has access to F-16's which can be used by MSFC & KSC to perform a ground test that demonstrates emergency power supplied to the jet. Neither of the green propellant alternatives have been considered nor evaluated for an APU application. Work has already been accomplished to characterize and obtain the properties of these 2 propellants. However, the spacecraft are using existing leak detection sensors that are typically used for Hydrazine. Using these green propellants for the APU application requires decrementing their TRL down to 3. This task would aim to establish a TRL of 4 at conclusion by showing a proof of concept with a KSC-instrumented EPU asset at the MSFC Component Development Area (CDA). The task to accomplish this is called Green Application for Space Power or GRASP.

LADEE Propulsion System Cold Flow Test

NASA Technical Reports Server (NTRS)

Williams, Jonathan Hunter; Chapman, Jack M.; Trinh, Hau, P.; Bell, James H.

2013-01-01

Lunar Atmosphere and Dust Environment Explorer (LADEE) is a NASA mission that will orbit the Moon. Its main objective is to characterize the atmosphere and lunar dust environment. The spacecraft development is being led by NASA Ames Research Center and scheduled for launch in 2013. The LADEE spacecraft will be operated with a bi-propellant hypergolic propulsion system using MMH and NTO as the fuel and oxidizer, respectively. The propulsion system utilizes flight-proven hardware on major components. The propulsion layout is composed of one 100-lbf main thruster and four 5-lbf RCS thrusters. The propellants are stored in four tanks (two parallel-connected tanks per propellant component). The propellants will be pressurized by regulated helium. A simulated propulsion system has been built for conducting cold flow test series to characterize the transient fluid flow of the propulsion system feed lines and to verify the critical operation modes, such as system priming, waterhammer, and crucial mission duty cycles. Propellant drainage differential between propellant tanks will also be assessed. Since the oxidizer feed line system has a higher flow demand than the fuel system does, the cold flow test focuses on the oxidizer system. The objective of the cold flow test is to simulate the LADEE propulsion fluid flow operation through water cold flow test and to obtain data for anchoring analytical models. The models will be used to predict the transient and steady state flow behaviors in the actual flight operations. The test activities, including the simulated propulsion test article, cold flow test, and analytical modeling, are being performed at NASA Marshall Space Flight Center. At the time of the abstract submission, the test article checkout is being performed. The test series will be completed by November, 2012

Model-Based Diagnostics for Propellant Loading Systems

NASA Technical Reports Server (NTRS)

Daigle, Matthew John; Foygel, Michael; Smelyanskiy, Vadim N.

2011-01-01

The loading of spacecraft propellants is a complex, risky operation. Therefore, diagnostic solutions are necessary to quickly identify when a fault occurs, so that recovery actions can be taken or an abort procedure can be initiated. Model-based diagnosis solutions, established using an in-depth analysis and understanding of the underlying physical processes, offer the advanced capability to quickly detect and isolate faults, identify their severity, and predict their effects on system performance. We develop a physics-based model of a cryogenic propellant loading system, which describes the complex dynamics of liquid hydrogen filling from a storage tank to an external vehicle tank, as well as the influence of different faults on this process. The model takes into account the main physical processes such as highly nonequilibrium condensation and evaporation of the hydrogen vapor, pressurization, and also the dynamics of liquid hydrogen and vapor flows inside the system in the presence of helium gas. Since the model incorporates multiple faults in the system, it provides a suitable framework for model-based diagnostics and prognostics algorithms. Using this model, we analyze the effects of faults on the system, derive symbolic fault signatures for the purposes of fault isolation, and perform fault identification using a particle filter approach. We demonstrate the detection, isolation, and identification of a number of faults using simulation-based experiments.

A Detailed Historical Review of Propellant Management Devices for Low Gravity Propellant Acquisition

NASA Technical Reports Server (NTRS)

Hartwig, Jason W.

2016-01-01

This paper presents a comprehensive background and historical review of Propellant Management Devices (PMDs) used throughout spaceflight history. The purpose of a PMD is to separate liquid and gas phases within a propellant tank and to transfer vapor-free propellant from a storage tank to a transfer line en route to either an engine or receiver depot tank, in any gravitational or thermal environment. The design concept, basic flow physics, and principle of operation are presented for each type of PMD. The three primary capillary driven PMD types of vanes, sponges, and screen channel liquid acquisition devices are compared and contrasted. For each PMD type, a detailed review of previous applications using storable propellants is given, which include space experiments as well as space missions and vehicles. Examples of previous cryogenic propellant management are also presented.

Recent Developments in Chemically Reactive Sensors for Propellants

NASA Technical Reports Server (NTRS)

Davis, Dennis D.; Mast, Dion J.; Baker, David L.; Fries, Joseph (Technical Monitor)

1999-01-01

Propellant system leaks can pose a significant hazard in aerospace operations. For example, a leak in the hydrazine supply system of the shuttle auxiliary power unit (APU) has resulted in hydrazine ignition and fire in the aft compartment of the shuttle. Sensors indicating the location of a leak could provide valuable information required for operational decisions. WSTF has developed a small, single-use sensor for detection of propellant leaks. The sensor is composed of a thermistor bead coated with a substance which is chemically reactive with the propellant. The reactive thermistor is one of a pair of closely located thermistors, the other being a reference. On exposure to the propellant, the reactive coating responds exothermically to it and increases the temperature of the coated-thermistor by several degrees. The temperature rise is sensed by a resistive bridge circuit, and an alarm is registered by data acquisition software. The concept is general and has been applied to sensors for hydrazine, monomethylhydrazine, unsym-dimethylhydrazine, ammonia, hydrogen peroxide, ethanol, and dinitrogen tetroxide. Responses of these sensors to humidity, propellant concentration, distance from the liquid leak, and ambient pressure levels arc presented. A multi-use sensor has also been developed for hydrazine based on its catalytic reactivity with noble metals.

78 FR 5859 - Agency Information Collection Activities: Requests for Comments; Clearance of New Approval of...

Federal Register 2010, 2011, 2012, 2013, 2014

2013-01-28

... about our intention to request the Office of Management and Budget (OMB) approval for a new information... propeller critical parts by the manufacturer, and establish engineering, manufacture, and maintenance... their manuals to record engineering, manufacture, and maintenance processes for propeller critical parts...

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.

Particle swarm optimization: an alternative in marine propeller optimization?

NASA Astrophysics Data System (ADS)

Vesting, F.; Bensow, R. E.

2018-01-01

This article deals with improving and evaluating the performance of two evolutionary algorithm approaches for automated engineering design optimization. Here a marine propeller design with constraints on cavitation nuisance is the intended application. For this purpose, the particle swarm optimization (PSO) algorithm is adapted for multi-objective optimization and constraint handling for use in propeller design. Three PSO algorithms are developed and tested for the optimization of four commercial propeller designs for different ship types. The results are evaluated by interrogating the generation medians and the Pareto front development. The same propellers are also optimized utilizing the well established NSGA-II genetic algorithm to provide benchmark results. The authors' PSO algorithms deliver comparable results to NSGA-II, but converge earlier and enhance the solution in terms of constraints violation.

The cohesive law of particle/binder interfaces in solid propellants

NASA Astrophysics Data System (ADS)

Tan, H.

2011-10-01

Solid propellants are treated as composites with high volume fraction of particles embedded in the polymeric binder. A micromechanics model is developed to establish the link between the microscopic behavior of particle/binder interfaces and the macroscopic constitutive information. This model is then used to determine the tension/shearing coupled interface cohesive law of a redesigned solid rocket motor propellant, based on the experimental data of the stress-strain and dilatation-strain curves for the material under slow rate uniaxial tension.

Motorboat propeller injuries.

PubMed

Di Nunno, N; Di Nunno, C

2000-07-01

The authors analyze the case of an Albanian refugee who was killed by the propellers of the outboard engine of a rubber dinghy while illegally attempting to reach Italy. The finding of multiple parallel, deep clear-cut injuries is uncommon, but highly characteristic of the object producing the lesions. These are typical and cannot be mistaken with those produced by sharp objects or shark bites. The description of the injuries is vital for establishing the position of the victim with regard to the propeller that struck him.

The alleged contributions of Pedro E. Paulet to liquid-propellant rocketry

NASA Technical Reports Server (NTRS)

Ordway, F. I., III

1977-01-01

The first practical working liquid propellant rocket motor was claimed by Pedro E. Paulet, a South American engineer from Peru (1895). He operated a conical motor, 10 centimeters in diameter, using nitrogen peroxide and gasoline as propellants and measuring thrust up to 90 kilograms, and apparently used spark ignition and intermittent propellant injection. The test device which he used contained elements of later test stands, such as a spring thrust-measuring device. However, he did not publish his work until twenty-five years later. Evidence is examined concerning this only known claim to liquid propellant rocket engine experiments in the nineteenth century.

Evaluation of cholesteatoma: our experience with DW Propeller imaging.

PubMed

Karandikar, Amit; Loke, Siu Cheng; Goh, Julian; Yeo, Seng Beng; Tan, Tiong Yong

2015-09-01

Cholesteatoma management includes early detection and surgical exploration. Due to its tendency to recur, it can be potentially locally aggressive. Magnetic resonance imaging (MRI), and in particular diffusion weighted imaging (DWI), plays an important role in management of these lesions. To assess the accuracy of Propeller (Periodically Rotated Overlapping ParallEL Lines with Enhanced Reconstruction) DW sequence in detecting middle ear and mastoid cholesteatomas in non-operated ears by surgical correlation. A retrospective review of 15 patients was done who underwent Propeller DWI with either clinically confirmed or suspected cholesteatomas. Surgical correlation was done in all cases. All patients had hyperintense foci on Propeller DWI. Surgical correlation performed revealed that 13 patients had cholesteatomas while two patients had mastoid abscesses. The location, extent, and size of cholesteatomas on Propeller DWI matched with the operative findings. Of the 13 patients with cholesteatomas, three patients had multiple foci of hyperintensity on Propeller DWI, which corroborated with the surgical finding of multiple cholesteatomas. The average apparent diffusion coefficient value of cholesteatoma was 0.868 × 10(-3) mm(2)/s, found to be higher than that of abscess, which was 0.425 × 10(-3) mm(2)/s. Propeller DWI was accurate in assessing the location, extent, and size of cholesteatomas as corroborated with surgical findings. Propeller DWI is useful in detecting number of cholesteatoma foci, a vital finding as it may impact the choice of surgery. © The Foundation Acta Radiologica 2014.

Liquid Oxygen Propellant Densification Production and Performance Test Results With a Large-Scale Flight-Weight Propellant Tank for the X33 RLV

NASA Technical Reports Server (NTRS)

Tomsik, Thomas M.; Meyer, Michael L.

2010-01-01

This paper describes in-detail a test program that was initiated at the Glenn Research Center (GRC) involving the cryogenic densification of liquid oxygen (LO2). A large scale LO2 propellant densification system rated for 200 gpm and sized for the X-33 LO2 propellant tank, was designed, fabricated and tested at the GRC. Multiple objectives of the test program included validation of LO2 production unit hardware and characterization of densifier performance at design and transient conditions. First, performance data is presented for an initial series of LO2 densifier screening and check-out tests using densified liquid nitrogen. The second series of tests show performance data collected during LO2 densifier test operations with liquid oxygen as the densified product fluid. An overview of LO2 X-33 tanking operations and load tests with the 20,000 gallon Structural Test Article (STA) are described. Tank loading testing and the thermal stratification that occurs inside of a flight-weight launch vehicle propellant tank were investigated. These operations involved a closed-loop recirculation process of LO2 flow through the densifier and then back into the STA. Finally, in excess of 200,000 gallons of densified LO2 at 120 oR was produced with the propellant densification unit during the demonstration program, an achievement that s never been done before in the realm of large-scale cryogenic tests.

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.

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.

Breadboard RL10-2B low-thrust operating mode (second iteration) test report

NASA Technical Reports Server (NTRS)

Kanic, Paul G.; Kaldor, Raymond B.; Watkins, Pia M.

1988-01-01

Cryogenic rocket engines requiring a cooling process to thermally condition the engine to operating temperature can be made more efficient if cooling propellants can be burned. Tank head idle and pumped idle modes can be used to burn propellants employed for cooling, thereby providing useful thrust. Such idle modes required the use of a heat exchanger to vaporize oxygen prior to injection into the combustion chamber. During December 1988, Pratt and Whitney conducted a series of engine hot firing demonstrating the operation of two new, previously untested oxidizer heat exchanger designs. The program was a second iteration of previous low thrust testing conducted in 1984, during which a first-generation heat exchanger design was used. Although operation was demonstrated at tank head idle and pumped idle, the engine experienced instability when propellants could not be supplied to the heat exchanger at design conditions.

Robust Exploration and Commercial Missions to the Moon Using LANTR Propulsion and In-Situ Propellants Derived from Lunar Polar Ice (LPI) Deposits

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Ryan, Stephen W.; Burke, Laura M.; McCurdy, David R.; Fittje, James E.; Joyner, Claude R.

2017-01-01

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable access through cislunar space necessary for commercial development and sustained human presence on the Moon. It is a demonstrated technology capable of generating both high thrust and high specific impulse (Isp 900 s) twice that of todays best chemical rockets. Nuclear lunar transfer vehicles consisting of a propulsion stage using three approx.16.5 klbf "Small Nuclear Rocket Engines (SNREs)", an in-line propellant tank, plus the payload can enable a variety of reusable lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong "tourism" missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing a robust in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in-situ propellant production used to supply strategically located propellant depots and transportation nodes. The processing of LPI deposits (estimated to be approx. 2 billion metric tons) for propellant production - specifically liquid oxygen (LO2) and hydrogen (LH2) can significantly reduce the launch mass requirements from Earth and can enable reusable, surface-based lunar landing vehicles (LLVs) using LO2/LH2 chemical rocket engines. Afterwards, LO2/LH2 propellant depots can be established in lunar polar and equatorial orbits to supply the LTS. At this point a modified version of the conventional NTR called the LO2-augmented NTR, or LANTR would be introduced into the LTS allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants (LDPs) for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an afterburner into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engines choked sonic throat essentially scramjet propulsion in reverse. By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and Isp values while the reactor core power level remains relatively constant. A LANTR-based LTS offers unique mission capabilities including short transit time crewed cargo transports. Even a commuter shuttle service may be possible allowing one-way trip times to and from the Moon on the order of 36 hours or less. If only 1 of the postulated water ice trapped in deep shadowed craters at the lunar poles were available for use in lunar orbit, such a supply could support daily commuter flights to the Moon for many thousands of years! The proposed paper outlines an evolutionary mission architecture and examines a variety of mission types and transfer vehicle designs, along with the increasing demands on LDP production as mission complexity and delta V requirements increase. A comparison of vehicle features and engine operating characteristics are also provided together with a discussion of the propellant production and mining requirements, and issues, associated with using LPI as the source material.

A Design Tool for Matching UAV Propeller and Power Plant Performance

NASA Astrophysics Data System (ADS)

Mangio, Arion L.

A large body of knowledge is available for matching propellers to engines for large propeller driven aircraft. Small UAV's and model airplanes operate at much lower Reynolds numbers and use fixed pitch propellers so the information for large aircraft is not directly applicable. A design tool is needed that takes into account Reynolds number effects, allows for gear reduction, and the selection of a propeller optimized for the airframe. The tool developed in this thesis does this using propeller performance data generated from vortex theory or wind tunnel experiments and combines that data with an engine power curve. The thrust, steady state power, RPM, and tip Mach number vs. velocity curves are generated. The Reynolds number vs. non dimensional radial station at an operating point is also found. The tool is then used to design a geared power plant for the SAE Aero Design competition. To measure the power plant performance, a purpose built engine test stand was built. The characteristics of the engine test stand are also presented. The engine test stand was then used to characterize the geared power plant. The power plant uses a 26x16 propeller, 100/13 gear ratio, and an LRP 0.30 cubic inch engine turning at 28,000 RPM and producing 2.2 HP. Lastly, the measured power plant performance is presented. An important result is that 17 lbf of static thrust is produced.

A fractional calculus perspective of distributed propeller design

NASA Astrophysics Data System (ADS)

Tenreiro Machado, J.; Galhano, Alexandra M.

2018-02-01

A new generation of aircraft with distributed propellers leads to operational performances superior to those exhibited by standard designs. Computational simulations and experimental tests show a reduction of fuel consumption and noise. This paper proposes an analogy between aerodynamics and electrical circuits. The model reveals properties similar to those of fractional-order systems and gives a deeper insight into the dynamics of multi-propeller coupling.

Preliminary Results of an Altitude-Wind-Tunnel Investigation of a TG-100A Gas Turbine-Propeller Engine II - Windmilling Characteristics

NASA Technical Reports Server (NTRS)

Conrad, E. W.; Durham, J. D.

1947-01-01

An investigation was conducted to determine the operational and performance characteristics of the TG-100A gas turbine-propeller engine II. Windmilling characteristics were deterined for a range of altitudes from 5000 to 35,000 feet, true airspeeds from 100 to 273 miles per hour, and propeller blade angles from 4 degrees to 46 degrees.

Mars ISPP Precursor (MIP): The First Flight Demonstration of In-Situ Propellant Production

NASA Technical Reports Server (NTRS)

Kaplan, David

1997-01-01

Strategic planning for human missions of exploration to Mars has conclusively identified in-situ propellant production (ISPP) as an enabling technology. The Mars reference mission concept predeploys a robotic propellant production plant to the planet two years before the planned departure of the crew from Earth. The successful operation of this plant is necessary for the human journey to begin.

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.

Computer prediction of three-dimensional potential flow fields in which aircraft propellers operate: Computer program description and users manual

NASA Technical Reports Server (NTRS)

Jumper, S. J.

1979-01-01

A method was developed for predicting the potential flow velocity field at the plane of a propeller operating under the influence of a wing-fuselage-cowl or nacelle combination. A computer program was written which predicts the three dimensional potential flow field. The contents of the program, its input data, and its output results are described.

In-Situ Propellant Supplied Lunar Lander Concept

NASA Astrophysics Data System (ADS)

Donahue, Benjamin; Maulsby, Curtis

2008-01-01

Future NASA and commercial Lunar missions will require innovative spacecraft configurations incorporating reliable, sustainable propulsion, propellant storage, power and crew life support technologies that can evolve into long duration, partially autonomous systems that can be used to emplace and sustain the massive supplies required for a permanently occupied lunar base. Ambitious surface science missions will require efficient Lunar transfer systems to provide the consumables, science equipment, energy generation systems, habitation systems and crew provisions necessary for lengthy tours on the surface. Lunar lander descent and ascent stages become significantly more efficient when they can be refueled on the Lunar surface and operated numerous times. Landers enabled by Lunar In-Situ Propellant Production (ISPP) facilities will greatly ease constraints on spacecraft mass and payload delivery capability, and may operate much more affordably (in the long term) then landers that are dependant on Earth supplied propellants. In this paper, a Lander concept that leverages ISPP is described and its performance is quantified. Landers, operating as sortie vehicles from Low Lunar Orbit, with efficiencies facilitated by ISPP will enable economical utilization and enhancements that will provide increasingly valuable science yields from Lunar Bases.

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.

Development and demonstration of flueric sounding rocket motor ignition

NASA Technical Reports Server (NTRS)

Marchese, V. P.

1974-01-01

An analytical and experimental program is described which established a flueric rocket motor ignition system concept incorporating a pneumatic match with a simple hand pump as the only energy source. An evaluation was made of this concept to determine the margins of the operating range and capabilities of every component of the system. This evaluation included a determination of power supply requirements, ignitor geometry and alinement, ignitor/propellant interfacing and materials and the effects of ambient temperatures and pressure. It was demonstrated that an operator using a simple hand pump for 30 seconds could ignite BKNO3 at a standoff distance of 100 m (330 ft) with the only connection to the ignitor being a piece of plastic pneumatic tubing.

MPD thruster technology

NASA Technical Reports Server (NTRS)

Myers, Roger M.; Mantenieks, Maris A.; Lapointe, Michael R.

1991-01-01

MPD (MagnetoPlasmaDynamic) thrusters demonstrated between 2000 and 7000 seconds specific impulse at efficiencies approaching 40 percent, and were operated continuously at power levels over 500 kW. These demonstrated capabilities, combined with the simplicity and robustness of the thruster, make them attractive candidates for application to both unmanned and manned orbit raising, lunar, and planetary missions. To date, however, only a limited number of thruster configurations, propellants, and operating conditions were studied. The present status of MPD research is reviewed, including developments in the measured performance levels and electrode erosion rates. Theoretical studies of the thruster dynamics are also described. Significant progress was made in establishing empirical scaling laws, performance and lifetime limitations and in the development of numerical codes to simulate the flow field and electrode processes.

Oxygen carrier for gas chromatographic analysis of inert gases in propellants

NASA Technical Reports Server (NTRS)

Cannon, W. A.

1972-01-01

Gas chromatographic determination of small quantities of inert gases in reactive propellants is discussed. Operating conditions used for specific analyses of helium in diborane and nitrogen in oxygen difluoride are presented in tabular form.

APEX 3D Propeller Test Preliminary Design

NASA Technical Reports Server (NTRS)

Colozza, Anthony J.

2002-01-01

A low Reynolds number, high subsonic mach number flight regime is fairly uncommon in aeronautics. Most flight vehicles do not fly under these aerodynamic conditions. However, recently there have been a number of proposed aircraft applications (such as high altitude observation platforms and Mars aircraft) that require flight within this regime. One of the main obstacles to flight under these conditions is the ability to reliably generate sufficient thrust for the aircraft. For a conventional propulsion system, the operation and design of the propeller is the key aspect to its operation. Due to the difficulty in experimentally modeling the flight conditions in ground-based facilities, it has been proposed to conduct propeller experiments from a high altitude gliding platform (APEX). A preliminary design of a propeller experiment under the low Reynolds number, high mach number flight conditions has been devised. The details of the design are described as well as the potential data that will be collected.

Occupational hazards of missile operations with special regard to the hydrazine propellants.

PubMed

Back, K C; Carter, V L; Thomas, A A

1978-04-01

The second generation of ballistic missiles and boosters, characterized by increased range and quick reaction capability, required the development of new high-energy storage propellants. This exploration led to the introduction of hydrazine (Hz), monomethylhydrazine (MMH), and 1,1-dimethylhydrazine (UDMH) into the USAF inventory. These compounds are all storable, noncryogenic, high-energy fuels which may be used alone or in combination as mixed amine fuels. Early toxicology experiments were to produce data on acute and subacute effects of the propellants in order to set standards for test and operational procedures to protect propellant handlers. The early work indicated that, despite similar chemical characteristics, there were marked differences between the compounds in terms of toxicological mechanisms. Since the propellant systems have been used for some 15 years, recent emphasis on toxicology has been centered on the more chronic effects and on an increasing body of evidence from animal experiments that the compounds may possess oncogenic potential as well as chronic systemic effects. This paper addresses itself to data leading up to current occupational standards.

STV fueling options

NASA Technical Reports Server (NTRS)

Flemming, Ken

1991-01-01

Lunar vehicles that will be space based and reusable will require resupply of propellants in orbit. Approximately 75 pct. of the total mass delivered to low earth orbit will be propellants. Consequently, the propellant management techniques selected for Space Exploration Initiative (SEI) orbital operations will have a major influence on the overall SEI architecture. Five proposed propellant management facility (PMF) concepts were analyzed and compared in order to determine the best method of resupplying reusable, space based Lunar Transfer Vehicles (LTVs). The processing time needed at the Space Station to prepare LTV for its next lunar mission was estimated for each of the PMF concepts. The estimated times required to assemble and maintain the different PMF concepts were also compared. The results of the maintenance analysis were similar, with co-orbiting depots needing 100 to 350 pct. more annual maintenance. The first few external tanks mating operations at KSC encountered many problems that could cause serious lunar mission schedule delays. The use of drop tanks on lunar vehicles increases by a factor of four the number of critical propellant interface disturbances.

Computer prediction of three-dimensional potential flow fields in which aircraft propellers operate. M.S. Thesis

NASA Technical Reports Server (NTRS)

Jumper, S. J.

1982-01-01

A computer program was developed to calculate the three dimensional, steady, incompressible, inviscid, irrotational flow field at the propeller plane (propeller removed) located upstream of an arbitrary airframe geometry. The program uses a horseshoe vortex of known strength to model the wing. All other airframe surfaces are modeled by a network source panels of unknown strength which is exposed to a uniform free stream and the wing-induced velocity field. By satisfying boundary conditions on each panel (the Neumann problem), relaxed boundary conditions being used on certain panels to simulate inlet inflow, the source strengths are determined. From the known source and wing vortex strengths, the resulting velocity fields on the airframe surface and at the propeller plane are obtained. All program equations are derived in detail, and a brief description of the program structure is presented. A user's manual which fully documents the program is cited. Computer predictions of the flow on the surface of a sphere and at a propeller plane upstream of the sphere are compared with the exact mathematical solutions. Agreement is good, and correct program operation is verified.

Development of a miniature solid propellant rocket motor for use in plume simulation studies

NASA Technical Reports Server (NTRS)

Baran, W. J.

1974-01-01

A miniature solid propellant rocket motor has been developed to be used in a program to determine those parameters which must be duplicated in a cold gas flow to produce aerodynamic effects on an experimental model similar to those produced by hot, particle-laden exhaust plumes. Phenomena encountered during the testing of the miniature solid propellant motors included erosive propellant burning caused by high flow velocities parallel to the propellant surface, regressive propellant burning as a result of exposed propellant edges, the deposition of aluminum oxide on the nozzle surfaces sufficient to cause aerodynamic nozzle throat geometry changes, and thermal erosion of the nozzle throat at high chamber pressures. A series of tests was conducted to establish the stability of the rocket chamber pressure and the repeatibility of test conditions. Data are presented which define the tests selected to represent the final test matrix. Qualitative observations are also presented concerning the phenomena experienced based on the results of a large number or rocket tests not directly applicable to the final test matrix.

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.

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.

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.

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).

Primary Electric Propulsion Technology Study. [for thruster wear-out mechanisms

NASA Technical Reports Server (NTRS)

Poeschel, R. L.; Beattie, J. R.

1979-01-01

An investigation of the 30-cm engineering-model-thruster technology with emphasis placed on the development of models for understanding and predicting the operational characteristics and wear-out mechanisms of the thruster as a function of operating or design parameters is presented. The task studies include: (1) the wear mechanisms and wear rates that determine the useful lifetime of the thruster discharge chamber; (2) cathode lifetime as determined by the depletion of barium from the barium-aluminate-impregnated-porous-tungsten insert that serves as a barium reservoir; (3) accelerator-grid-system technology; (4) a verification of the high-voltage propellant-flow-electrical-isolator design developed under NASA contract NAS3-20395 for operation at 10-kV applied voltage and 10-A equivalent propellant flow with mercury and argon propellants. A model was formulated for predicting performance.

Adaptive control of a jet turboshaft engine driving a variable pitch propeller using multiple models

NASA Astrophysics Data System (ADS)

Ahmadian, Narjes; Khosravi, Alireza; Sarhadi, Pouria

2017-08-01

In this paper, a multiple model adaptive control (MMAC) method is proposed for a gas turbine engine. The model of a twin spool turbo-shaft engine driving a variable pitch propeller includes various operating points. Variations in fuel flow and propeller pitch inputs produce different operating conditions which force the controller to be adopted rapidly. Important operating points are three idle, cruise and full thrust cases for the entire flight envelope. A multi-input multi-output (MIMO) version of second level adaptation using multiple models is developed. Also, stability analysis using Lyapunov method is presented. The proposed method is compared with two conventional first level adaptation and model reference adaptive control techniques. Simulation results for JetCat SPT5 turbo-shaft engine demonstrate the performance and fidelity of the proposed method.

Experimental research and design planning in the field of liquid-propellant rocket engines conducted between 1934 - 1944 by the followers of F. A. Tsander

NASA Technical Reports Server (NTRS)

Dushkin, L. S.

1977-01-01

The development of the following Liquid-Propellant Rocket Engines (LPRE) is reviewed: (1) an alcohol-oxygen single-firing LPRE for use in wingless and winged rockets, (2) a similar multifiring LPRE for use in rocket gliders, (3) a combined solid-liquid propellant rocket engine, and (4) an aircraft LPRE operating on nitric acid and kerosene.

Vibrations of a Marine Propeller Operating in a Nonuniform Inflow.

DTIC Science & Technology

1980-04-01

Expanded Blade Midsurface ......... ........................ ... 73 16 - Calculated Normalized Propeller RMS Vibration Velocity as a Function of...averaged over the blade midsurface ), rather thaft the maximum velocities near the blade tip. Then, for the two test propellers, the rms nonuniform inflow...time- averaged midsurface of the blade, then the instantaneous position S of the vibrating midsurface is _S (ric)+ qct S(r,c,t) = (rc) + q(t) i(rc

Evaluation of supercritical cryogen storage and transfer systems for future NASA missions

NASA Technical Reports Server (NTRS)

Arif, Hugh; Aydelott, John C.; Chato, David J.

1990-01-01

Conceptual designs of Space Transportation Vehicles (STV), and their orbital servicing facilities, that utilize supercritical, single phase, cryogenic propellant were established and compared with conventional subcritical, two phases, STV concepts. The analytical study was motivated by the desire to avoid fluid management problems associated with the storage, acquisition and transfer of subcritical liquid oxygen and hydrogen propellants in the low gravity environment of space. Although feasible, the supercritical concepts suffer from STV weight penalties and propellant resupply system power requirements which make the concepts impractical.

Evaluation of supercritical cryogen storage and transfer systems for future NASA missions

NASA Technical Reports Server (NTRS)

Arif, Hugh; Aydelott, John C.; Chato, David J.

1989-01-01

Conceptual designs of Space Transportation Vehicles (STV), and their orbital servicing facilities, that utilize supercritical, single phase, cryogenic propellants were established and compared with conventional subcritical, two phase, STV concepts. The analytical study was motivated by the desire to avoid fluid management problems associated with the storage, acquisition and transfer of subcritical liquid oxygen and hydrogen propellants in the low gravity environment of space. Although feasible, the supercritical concepts suffer from STV weight penalties and propellant resupply system power requirements which make the concepts impractical.

External And Internal Work Of A T-6 Paraplegic Propelling A Wheelchair And Arm Cranking A Cycle Ergometer: Case Study

NASA Astrophysics Data System (ADS)

Novak, Charles W.

1982-02-01

In this, the International Year of the Disabled, attention is directed among other areas toward rehabilitation and sports participation of wheelchair users. As an application of movement analysis in medicine and rehabilitation and as an application of sports research using biomechanics, this investigation was performed to compare the results of two methods of gathering data on the stress of wheelchair propelling at equivalent work loads and to account for differences in physiological responses with a mechanical analysis of wheelchair propelling. Physiological data collected were heart rate, systolic blood pressure, and rate-pressure product. A biomechanical cinematography analysis was used to determine external work in wheelchair propelling and to determine the extent to which modifications in segment actionsoccurred during increasing magnitude of work. A cycle ergometer was adjusted to replicate external work loads performed during wheelchair propelling. A t-test of equivalent external work loads indicated that heart rate was not different between the two exercise modes at the .05 level of significance. The t-test did indicate a significant difference in systolic blood pressure and rate-pressure product at the .05 level of significance. The biomechanical analysis of wheelchair propelling established that an increase in external work was accomplished by a decrease in the range of motion and an increase in the speed of movement. During cycle ergometry the range and speed of movement remained the same while resistance was increased. Results of the study established that while heart rate for equivalent external work loads was the same for wheelchair propelling and arm cranking cycle ergometry, systolic blood pressure and rate-pressure product were not the same. The suggestion was that some means of propelling a wheelchair other than that which is con-sidered "standard" might be considered which produces less stressful responses in wheelchair users.

Recovery of Waste Heat from Propellant Forced-Air Dry House

DTIC Science & Technology

1978-12-01

function of bulk air side film heat transfer coefficient and diffusivity 66 15. Dry house waste heat recovery system instrumentation 67 16. Sample data...inlet condition by, maintaining the exhaust temperature above the NG dew point. The set point is adjustable to accommodate various propel- lant and...system. In dry cycle operation, an overall energy recovery effectiveness of about 40% was measured for winter operation when the exhaust temperature

Interior noise levels of two propeller-driven light aircraft

NASA Technical Reports Server (NTRS)

Catherines, J. J.; Mayes, W. H.

1975-01-01

The relationships between aircraft operating conditions and interior noise and the degree to which ground testing can be used in lieu of flight testing for performing interior noise research were studied. The results show that the noise inside light aircraft is strongly influenced by the rotational speed of the engine and propeller. Both the overall noise and low frequency spectra levels were observed to decrease with increasing high speed rpm operations during flight. This phenomenon and its significance is not presently understood. Comparison of spectra obtained in flight with spectra obtained on the ground suggests that identification of frequency components and relative amplitude of propeller and engine noise sources may be evaluated on stationary aircraft.

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.

Ultra High Voltage Propellant Isolators and Insulators for JIMO Ion Thrusters

NASA Technical Reports Server (NTRS)

Banks, Bruce A.; Gaier, James R.; Hung, Ching-Cheh; Walters, Patty A.; Sechkar, Ed; Panko, Scott; Kamiotis, Christina A.

2004-01-01

Within NASA's Project Prometheus, high specific impulse ion thrusters for electric propulsion of spacecraft for the proposed Jupiter Icy Moon Orbiter (JIMO) mission to three of Jupiter's moons: Callisto, Ganymede and Europa will require high voltage operation to meet mission propulsion. The anticipated approx.6,500 volt net ion energy will require electrical insulation and propellant isolation which must exceed that used successfully by the NASA Solar Electric Propulsion Technology Readiness (NSTAR) Deep Space 1 mission thruster by a factor of approx.6. Xenon propellant isolator prototypes that operate at near one atmosphere and prototypes that operate at low pressures (<100 Torr) have been designed and are being tested for suitability to the JIMO mission requirements. Propellant isolators must be durable to Paschen breakdown, sputter contamination, high temperature, and high voltage while operating for factors longer duration than for the Deep Space 1 Mission. Insulators used to mount the thrusters as well as those needed to support the ion optics have also been designed and are under evaluation. Isolator and insulator concepts, design issues, design guidelines, fabrication considerations and performance issues are presented. The objective of the investigation was to identify candidate isolators and insulators that are sufficiently robust to perform durably and reliably during the proposed JIMO mission.

Investigation of the Discharge Characteristics of the T6 Hollow Cathode Operating on Several Inert Gases and a Kr/Xe Mixture

NASA Astrophysics Data System (ADS)

Ahmed Rudwan, M.; Gabriel, S. B.

2002-01-01

Investigation of the discharge characteristics of the T6 hollow cathode operating on several inert Xenon is currently the propellant of choice for gridded ion thrusters. But in order to make deep space missions feasible, an increase in the Specific Impulse (SI) that these thrusters can achieve is necessary. One method of achieving this is to use a propellant with a lower atomic mass (e.g. argon), as the propellant exhaust velocity is inversely proportional to the square root of the propellant mass. However, the feasibility of operating the hollow cathode using these alternative propellants has to be demonstrated. Moreover, interest in decreasing the propellant cost in missions and ground testing (especially life tests) have led to the comprehensive discharge characterisation of several gases that will be presented in this paper. A Kr/Xe mixture in the naturally occurring ratio, for example, could offer a 15 times cost saving when compared to pure xenon and 2-3 times cost saving when compared to pure krypton. The T6 hollow cathode discharge behaviour as well as its initiation characteristics have been studied. The tests were carried out in diode configuration using a T6 hollow cathode with an enclosed keeper design employing xenon, krypton, argon and a Kr/Xe mix. The discharge initiation tests were undertaken with a view to investigate some of the factors thought to influence the starting potential such as mass flow rate and tip temperature. It was found that, for mass flow rates ranging from 0.2-1.1 mg/s and cathode tip temperatures ranging from 900-1300oC, the breakdown potential was less than 50V for argon, less than 25V for krypton, less than 21V for xenon and less than 35V for the Kr/Xe mix. The discharge initiation results were then compared to those obtained by Fearn et al. with a T5 cathode operating on mercury and with a T6 cathode utilising an open keeper design using xenon propellant. The xenon breakdown potentials were found to be lower than those obtained with an open keeper design by as much as 4V. Steady state discharge behaviour was also investigated in a range of operating conditions. Spot to plume mode transitions were observed in argon, krypton and Kr/Xe discharges for the first time.

NASA's Evolutionary Xenon Thruster (NEXT) Long-Duration Test as of 736 kg of Propellant Throughput

NASA Technical Reports Server (NTRS)

Shastry, Rohit; Herman, Daniel A.; Soulas, George C.; Patterson, Michael J.

2012-01-01

The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation solar-electric ion propulsion system with significant enhancements beyond the state-of-the-art NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) ion propulsion system to provide future NASA science missions with enhanced mission capabilities. A Long-Duration Test (LDT) was initiated in June 2005 to validate the thruster service life modeling and to qualify the thruster propellant throughput capability. The thruster has set electric propulsion records for the longest operating duration, highest propellant throughput, and most total impulse demonstrated. At the time of this publication, the NEXT LDT has surpassed 42,100 h of operation, processed more than 736 kg of xenon propellant, and demonstrated greater than 28.1 MN s total impulse. Thruster performance has been steady with negligible degradation. The NEXT thruster design has mitigated several lifetime limiting mechanisms encountered in the NSTAR design, including the NSTAR first failure mode, thereby drastically improving thruster capabilities. Component erosion rates and the progression of the predicted life-limiting erosion mechanism for the thruster compare favorably to pretest predictions based upon semi-empirical ion thruster models used in the thruster service life assessment. Service life model validation has been accomplished by the NEXT LDT. Assuming full-power operation until test article failure, the models and extrapolated erosion data predict penetration of the accelerator grid grooves after more than 45,000 hours of operation while processing over 800 kg of xenon propellant. Thruster failure due to degradation of the accelerator grid structural integrity is expected after groove penetration.

NASA's Evolutionary Xenon Thruster (NEXT) Long-Duration Test as of 736 kg of Propellant Throughput

NASA Technical Reports Server (NTRS)

Shastry, Rohit; Herman, Daniel A.; Soulas, George C.; Patterson, Michael J.

2012-01-01

The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation solar-electric ion propulsion system with significant enhancements beyond the state-of-the-art NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) ion propulsion system to provide future NASA science missions with enhanced mission capabilities. A Long-Duration Test (LDT) was initiated in June 2005 to validate the thruster service life modeling and to qualify the thruster propellant throughput capability. The thruster has set electric propulsion records for the longest operating duration, highest propellant throughput, and most total impulse demonstrated. At the time of this publication, the NEXT LDT has surpassed 42,100 h of operation, processed more than 736 kg of xenon propellant, and demonstrated greater than 28.1 MN s total impulse. Thruster performance has been steady with negligible degradation. The NEXT thruster design has mitigated several lifetime limiting mechanisms encountered in the NSTAR design, including the NSTAR first failure mode, thereby drastically improving thruster capabilities. Component erosion rates and the progression of the predicted life-limiting erosion mechanism for the thruster compare favorably to pretest predictions based upon semi-empirical ion thruster models used in the thruster service life assessment. Service life model validation has been accomplished by the NEXT LDT. Assuming full-power operation until test article failure, the models and extrapolated erosion data predict penetration of the accelerator grid grooves after more than 45,000 hours of operation while processing over 800 kg of xenon propellant. Thruster failure due to degradation of the accelerator grid structural integrity is expected after

Design and use of a servo-controlled high pressure window bomb in spectroscopic studies of solid propellant combustion

NASA Technical Reports Server (NTRS)

Goetz, F.; Mann, D. M.

1980-01-01

The feasibility of using a high pressure window bomb as a laboratory scale model of actual motor conditions. The design and operation of a modified high pressure window bomb is discussed. An optical servocontrol mechanism has been designed to hold the burning surface of a propellant strand at a fixed position within the bomb chamber. This mechanism permits the recording of visible and infrared emission spectra from various propellants. Preliminary visible emission spectra of a nonmetalized and metalized propellant are compared with spectra recorded using the modified bomb.

En route noise of two turboprop aircraft

NASA Technical Reports Server (NTRS)

Dobrzynski, Werner

1990-01-01

In order to weigh en route noise emissions originating from future propfan powered aircraft, a data base of emission levels from conventional turboprop aircraft is needed. For this reason flyover noise measurements on two twin-engine turboprop aircraft were conducted at flight heights between 17,000 and 21,000 ft. Acoustic data are presented together with propeller operational parameters and environmental meteorological data. Narrowband spectral analyses demonstrate the characteristic features of the measured propeller noise signatures: Noise spectra are dominated by the propeller rotational noise fundamental frequency and pronounced noise beats occur as a consequence of different rotational speeds of the propellers.

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.

A method for calculation of free-space sound pressures near a propeller in flight including considerations of the chordwise blade loading

NASA Technical Reports Server (NTRS)

Watkins, Charles E; Durling, Barbara J

1956-01-01

This report presents tabulated values of certain definite integral that are involved in the calculation of near-field propeller noise when the chordwise forces are assumed to be either uniform or of a Dirac delta type. The tabulations are over a wide range of operating conditions and are useful for estimating propeller noise when either the concept of an effective radius or radial distributions of forces are considered. Use of the tabulations is illustrated by several examples of calculated results for some specific propellers.

An investigation of in-flight near-field propeller noise generation and transmission

NASA Astrophysics Data System (ADS)

Bonneau, H.; Wilford, D. F.; Wood, L. K.

1985-02-01

In flight near field propeller noise measurements, made on a General Aviation turboprop aircraft, are reported for a range of propeller operating conditions, and are shown to be well defined and reproducible. Measurements have been made at 8 exterior microphones, 2 located on a wing mounted boom, and 6 embedded in, and flush with the aircraft fuselage. Interior noise levels are also presented. Measured propeller harmonic levels are compared to first principle calculations of near field noise, using a modified version of the Farassat computer program, in which the blade surface pressure is described using the known aerodynamic properties of the blade (NACA 16) airfoil sections. The first few; i.e., the dominant harmonic levels of propeller noise are shown to be well predicted, while higher harmonic levels are underpredicted. The transmission loss between exterior and interior noise levels is shown to be relatively constant for varying propeller operating conditions and at two different locations along the length of the fuselage. Interior noise levels are also shown for the aircraft in gliding flight at various forward velocities, with both engines at idle and propellers feathered. A method of interpolating these measurements is discussed, which allows the interior noise due only to the forward velocity of the aircraft, to be determined. The transmission loss for this component is also discussed. Finally, interior noise levels are presented for a series of ground static tests with engine mounts of various different stiffnessses.

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.

Investigations Into Tank Venting for Propellant Resupply

NASA Technical Reports Server (NTRS)

Hearn, H. C.; Harrison, Robert A. (Technical Monitor)

2002-01-01

Models and simulations have been developed and applied to the evaluation of propellant tank ullage venting, which is integral to one approach for propellant resupply. The analytical effort was instrumental in identifying issues associated with resupply objectives, and it was used to help develop an operational procedure to accomplish the desired propellant transfer for a particular storable bipropellant system. Work on the project was not completed, and several topics have been identified as requiring further study; these include the potential for liquid entrainment during the low-g and thermal/freezing effects in the vent line and orifice. Verification of the feasibility of this propellant venting and resupply approach still requires additional analyses as well as testing to investigate the fluid and thermodynamic phenomena involved.

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.

Flight demonstration of new thruster and green propellant technology on the PRISMA satellite

NASA Astrophysics Data System (ADS)

Anflo, K.; Möllerberg, R.

2009-11-01

The concept of a storable liquid monopropellant blend for space applications based on ammonium dinitramide (ADN) was invented in 1997, within a co-operation between the Swedish Space Corporation (SSC) and the Swedish Defense Research Agency (FOI). The objective was to develop a propellant which has higher performance and is safer than hydrazine. The work has been performed under contract from the Swedish National Space Board and ESA. The progress of the development has been presented in several papers since 2000. ECAPS, a subsidiary of the Swedish Space Corporation was established in 2000 with the aim to develop and market the novel "high performance green propellant" (HPGP) technology for space applications. The new technology is based on several innovations and patents w.r.t. propellant formulation and thruster design, including a high temperature resistant catalyst and thrust chamber. The first flight demonstration of the HPGP propulsion system will be performed on PRISMA. PRISMA is an international technology demonstration program with Swedish Space Corporation as the Prime Contractor. This paper describes the performance, characteristics, design and verification of the HPGP propulsion system for PRISMA. Compatibility issues related to using a new propellant with COTS components is also discussed. The PRISMA mission includes two satellites in LEO orbit were the focus is on rendezvous and formation flying. One of the satellites will act as a "target" and the main spacecraft performs rendezvous and formation flying maneuvers, where the ECAPS HPGP propulsion system will provide delta-V capability. The PRISMA CDR was held in January 2007. Integration of the flight propulsion system is about to be finalized. The flight opportunity on PRISMA represents a unique opportunity to demonstrate the HPGP propulsion system in space, and thus take a significant step towards its use in future space applications. The launch of PRISMA scheduled to 2009.

Space Transportation Infrastructure Supported By Propellant Depots

NASA Technical Reports Server (NTRS)

Smitherman, David; Woodcock, Gordon

2011-01-01

A space transportation infrastructure is described that utilizes propellant depots to support all foreseeable missions in the Earth-Moon vicinity and deep space out to Mars. The infrastructure utilizes current expendable launch vehicles such as the Delta IV Heavy, Atlas V, and Falcon 9, for all crew, cargo, and propellant launches to orbit. Propellant launches are made to a Low-Earth-Orbit (LEO) Depot and an Earth-Moon Lagrange Point 1 (L1) Depot to support new reusable in-space transportation vehicles. The LEO Depot supports missions to Geosynchronous Earth Orbit (GEO) for satellite servicing, and to L1 for L1 Depot missions. The L1 Depot supports Lunar, Earth-Sun L2 (ESL2), Asteroid, and Mars missions. A Mars Orbital Depot is also described to support ongoing Mars missions. New concepts for vehicle designs are presented that can be launched on current 5-meter diameter expendable launch vehicles. These new reusable vehicle concepts include a LEO Depot, L1 Depot, and Mars Orbital Depot based on International Space Station (ISS) heritage hardware. The high-energy depots at L1 and Mars orbit are compatible with, but do not require, electric propulsion tug use for propellant and/or cargo delivery. New reusable in-space crew transportation vehicles include a Crew Transfer Vehicle (CTV) for crew transportation between the LEO Depot and the L1 Depot, a new reusable Lunar Lander for crew transportation between the L1 Depot and the lunar surface, and a Deep Space Habitat (DSH) to support crew missions from the L1 Depot to ESL2, Asteroid, and Mars destinations. A 6 meter diameter Mars lander concept is presented that can be launched without a fairing based on the Delta IV heavy Payload Planners Guide, which indicates feasibility of a 6.5 meter fairing. This lander would evolve to re-usable operations when propellant production is established on Mars. Figure 1 provides a summary of the possible missions this infrastructure can support. Summary mission profiles are presented for each primary mission capability. These profiles are the basis for propellant loads, numbers of vehicles/stages and launches for each mission capability. Data includes the number of launches required for each mission utilizing current expendable launch vehicle systems, and concluding remarks include ideas for reducing the number of launches through incorporation of heavy-lift launch vehicles, solar electric propulsion, and other transportation support concepts.

RASSOR - Regolith Advanced Surface Systems Operations Robot

NASA Technical Reports Server (NTRS)

Gill, Tracy R.; Mueller, Rob

2015-01-01

The Regolith Advanced Surface Systems Operations Robot (RASSOR) is a lightweight excavator for mining in reduced gravity. RASSOR addresses the need for a lightweight (<100 kg) robot that is able to overcome excavation reaction forces while operating in reduced gravity environments such as the moon or Mars. A nominal mission would send RASSOR to the moon to operate for five years delivering regolith feedstock to a separate chemical plant, which extracts oxygen from the regolith using H2 reduction methods. RASSOR would make 35 trips of 20 kg loads every 24 hours. With four RASSORs operating at one time, the mission would achieve 10 tonnes of oxygen per year (8 t for rocket propellant and 2 t for life support). Accessing craters in space environments may be extremely hard and harsh due to volatile resources - survival is challenging. New technologies and methods are required. RASSOR is a product of KSC Swamp Works which establishes rapid, innovative and cost effective exploration mission solutions by leveraging partnerships across NASA, industry and academia.

Vented Chill / No-Vent Fill of Cryogenic Propellant Tanks

NASA Technical Reports Server (NTRS)

Rhys, Noah O.; Foster, Lee W.; Martin, Adam K.; Stephens, Jonathan R.

2016-01-01

Architectures for extended duration missions often include an on-orbit replenishment of the space vehicle's cryogenic liquid propellants. Such a replenishment could be accomplished via a tank-to-tank transfer from a dedicated tanker or a more permanent propellant depot storage tank. Minimizing the propellant loss associated with transfer line and receiver propellant tank thermal conditioning is essential for mass savings. A new methodology for conducting tank-to-tank transfer while minimizing such losses has been demonstrated. Charge-Hold-Vent is the traditional methodology for conducting a tank-to-tank propellant transfer. A small amount of cryogenic liquid is introduced to chill the transfer line and propellant tank. As the propellant absorbs heat and undergoes a phase change, the tank internal pressure increases. The tank is then vented to relieve pressure prior to another charge of cryogenic liquid being introduced. This cycle is repeated until the transfer lines and tank are sufficiently chilled and the replenishment of the propellant tank is complete. This method suffers inefficiencies due to multiple chill and vent cycles within the transfer lines and associated feed system components. Additionally, this system requires precise measuring of cryogenic fluid delivery for each transfer, multiple valve cycling events, and other complexities associated with cycled operations. To minimize propellant loss and greatly simplify on-orbit operations, an alternate methodology has been designed and demonstrated. The Vented Chill / No Vent Fill method is a simpler, constant flow approach in which the propellant tank and transfer lines are only chilled once. The receiver tank is continuously vented as cryogenic liquid chills the transfer lines, tank mass and ullage space. Once chilled sufficiently, the receiver tank valve is closed and the tank is completely filled. Interestingly, the vent valve can be closed prior to receiver tank components reaching liquid saturation temperature. An incomplete fill results if insufficient energy is removed from the tank's thermal mass and ullage space. The key to successfully conducting the no vent fill is to assure that sufficient energy is removed from the system prior to closing the receiver tank vent valve. This paper will provide a description of the transfer methodology and test article, and will provide a discussion of test results.

Measurements of Free-Space Oscillating Pressures Near Propellers at Flight Mach Numbers to 0.72

NASA Technical Reports Server (NTRS)

Kurbjun, Max C; Vogeley, Arthur W

1958-01-01

In the course of a short flight program initiated to check the theory of Garrick and Watkins (NACA rep. 1198), a series of measurements at three stations were made of the oscillating pressures near a tapered-blade plan-form propeller and rectangular-blade plan form propeller at flight Mach numbers up to 0.72. In contradiction to the results for the propeller studied in NACA rep. 1198, the oscillating pressures in the plane ahead of the propeller were found to be higher than those immediately behind the propeller. Factors such as variation in torque and thrust distribution, since the blades of the present investigation were operating above their design forward speed, may account for this contradiction. The effect of blade plan form shows that a tapered-blade plan-form propeller will produce lower sound-pressure levels than a rectangular-blade plan-form propeller for the low blade-passage harmonics (the frequencies where structural considerations are important) and produce higher sound-pressure levels for the higher blade-passage harmonics (frequencies where passenger comfort is important).

Experimental verification of propeller noise prediction

NASA Technical Reports Server (NTRS)

Succi, G. P.; Munro, D. H.; Zimmer, J. A.

1980-01-01

Results of experimental measurements of the sound fields of 1/4-scale general aviation propellers are presented and experimental wake surveys and pressure signatures obtained are compared with theoretical predictions. Experiments were performed primarily on a 1C160 propeller model mounted in front of a symmetric body in an anechoic wind tunnel, and measured the thrust and torque produced by propeller at different rotation speeds and tunnel velocities, wakes at three axial distances, and sound pressure at various azimuths and tip speeds with advance ratio or tunnel velocity constant. Aerodynamic calculations of blade loading were performed using airfoil section characteristics and a modified strip analysis procedure. The propeller was then modeled as an array of point sound sources with each point characterized by the force and volume of the corresponding propeller section in order to obtain the acoustic characteristics. Measurements are found to agree with predictions over a wide range of operating conditions, tip speeds and propeller nacelle combinations, without the use of adjustable constants.

An assessment of propeller aircraft noise reduction technology

NASA Technical Reports Server (NTRS)

Metzger, F. Bruce

1995-01-01

This report is a review of the literature regarding propeller airplane far-field noise reduction. Near-field and cabin noise reduction are not specifically addressed. However, some of the approaches used to reduce far-field noise produce beneficial effects in the near-field and in the cabin. The emphasis is on propeller noise reduction but engine exhaust noise reduction by muffling is also addressed since the engine noise becomes a significant part of the aircraft noise signature when propeller noise is reduced. It is concluded that there is a substantial body of information available that can be used as the basis to reduce propeller airplane noise. The reason that this information is not often used in airplane design is the associated weight, cost, and performance penalties. It is recommended that the highest priority be given to research for reducing the penalties associated with lower operating RPM and propeller diameter while increasing the number of blades. Research to reduce engine noise and explore innovative propeller concepts is also recommended.

Space tug economic analysis study. Volume 1: Executive summary

NASA Technical Reports Server (NTRS)

1972-01-01

An economic analysis of space tug operations is presented. The space tug is defined as any liquid propulsion stage under 100,000 pounds propellant loading that is flown from the space shuttle cargo bay. Two classes of vehicles are the orbit injection stages and reusable space tugs. The vehicle configurations, propellant combinations, and operating modes used for the study are reported. The summary contains data on the study approach, results, conclusions, and recommendations.

Cruise noise of counterrotation propeller at angle of attack in wind tunnel

NASA Technical Reports Server (NTRS)

Dittmar, J. H.

1986-01-01

The noise of a counterrotation propeller at angle of attack was measured in the NASA Lewis 8- by 6-Foot Supersonic Wind Tunnel at cruise conditions. Noise increases of as much as 4 dB were measured at positive angles of attack on the tunnel side wall, which represented an airplane fuselage. These noise increases could be minimized or eliminated by operating the counterrotation propeller with the front propeller turning up-inboard. This would require oppositely rotating propellers on opposite sides of the airplane. Noise analyses at different bandwidths enabled the separate front- and rear-propeller tones, as well as the total noise, at each harmonic to be determined. A simplified noise model was explored to show how the observed circumferential noise patterns of the separate propeller tones might have occurred. The total noise pattern, which represented the sum of the front- and rear-propeller tones at a particular harmonic, showed trends that would be hard to interpret without the separate-tone results. Therefore it is important that counterrotation angle-of-attack noise data be taken in such a manner that the front- and rear-propeller tones can be separated.

3-D viscous flow CFD analysis of the propeller effect on an advanced ducted propeller subsonic inlet

NASA Technical Reports Server (NTRS)

Iek, Chanthy; Boldman, Donald R.; Ibrahim, Mounir

1993-01-01

The time-marching Navier-Stokes code PARC3D was used to study the 3D viscous flow associated with an advanced ducted propeller subsonic inlet at take-off operating conditions. At a free stream Mach number of 0.2, experimental data for the inlet-with-propeller test model indicated that the airflow was attached on the cowl windward lip at an angle of attack of 25 deg became unstable at 29 deg, and separated at 30 deg. An experimental study with a similar inlet and without propeller (through-flow) indicated that flow separation occurred at an angle of attack a few degrees below the value observed when the inlet was tested with the propeller, indicating the propeller's favorable effect on inlet performance. In the present numerical study, flow blockage analogous to the propeller was modeled via a PARC3D computational boundary condition (BC), the 'screen BC', based on 1-1/2 dimension actuator disk theory. The application of the screen BC in this numerical study provided results similar to those of past experimental efforts in which either the blockage device or the propeller was used.

46 CFR 153.1 - Applicability.

Code of Federal Regulations, 2010 CFR

2010-10-01

...'s Certificate of Inspection is endorsed for a limited short protected coastwise route and the ship..., DEPARTMENT OF HOMELAND SECURITY (CONTINUED) CERTAIN BULK DANGEROUS CARGOES SHIPS CARRYING BULK LIQUID... following: (a) All United States self-propelled ships and those foreign self-propelled ships operating in...

14 CFR 417.417 - Propellants and explosives.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 417.417 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION LICENSING LAUNCH SAFETY Ground Safety § 417.417 Propellants and explosives. (a) A launch operator must comply with the explosive safety criteria in part 420 of this chapter. (b) A...

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.

Cryogenic Fluid Management Technology and Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Taylor, Brian D.; Caffrey, Jarvis; Hedayat, Ali; Stephens, Jonathan; Polsgrove, Robert

2016-01-01

Cryogenic fluid management (CFM) is critical to the success of future nuclear thermal propulsion powered vehicles. While this is an issue for any propulsion system utilizing cryogenic propellants, this is made more challenging by the radiation flux produced by the reactor in a nuclear thermal rocket (NTR). Managing the cryogenic fuel to prevent propellant loss to boil off and leakage is needed to limit the required quantity of propellant to a reasonable level. Analysis shows deposition of energy into liquid hydrogen fuel tanks in the vicinity of the nuclear thermal engine. This is on top of ambient environment sources of heat. Investments in cryogenic/thermal management systems (some of which are ongoing at various organizations) are needed in parallel to nuclear thermal engine development in order to one day see the successful operation of an entire stage. High durability, low thermal conductivity insulation is one developmental need. Light weight cryocoolers capable of removing heat from large fluid volumes at temperatures as low as approx. 20 K are needed to remove heat leak from the propellant of an NTR. Valve leakage is an additional CFM issue of great importance. Leakage rates of state of the art, launch vehicle size valves (which is approximately the size valves needed for a Mars transfer vehicle) are quite high and would result in large quantities of lost propellant over a long duration mission. Additionally, the liquid acquisition system inside the propellant tank must deliver properly conditioned propellant to the feed line for successful engine operation and avoid intake of warm or gaseous propellant. Analysis of the thermal environment and the CFM technology development are discussed in the accompanying presentation.

Numerical Prediction of Magnetic Cryogenic Propellant Storage in Reduced Gravity

NASA Astrophysics Data System (ADS)

Marchetta, J. G.; Hochstein, J. I.

2002-01-01

Numerical Prediction of Magnetic Cryogenic Propellant Storage in Reduced strong evidence that a magnetic positioning system may be a feasible alternative technology for use in the management of cryogenic propellants onboard spacecraft. The results of these preliminary studies have indicated that further investigation of the physical processes and potential reliability of such a system is required. The utility of magnetic fields as an alternative method in cryogenic propellant management is dependent on its reliability and flexibility. Simulations and experiments have previously yielded evidence in support of the magnetic positive positioning (MPP) process to predictably reorient LOX for a variety of initial conditions. Presently, though, insufficient evidence has been established to support the use of magnetic fields with respect to the long-term storage of cryogenic propellants. Current modes of propellant storage have met with a moderate level of success and are well suited for short duration missions using monopropellants. However, the storage of cryogenic propellants warrants additional consideration for long-term missions. For example, propellant loss during storage is due to vaporization by incident solar radiation and the vaporized ullage must be vented to prevent excessive pressurization of the tank. Ideally, positioning the fluid in the center of the tank away from the tank wall will reduce vaporization by minimizing heat transfer through the tank wall to the liquid. A second issue involves the capability of sustaining a stable fluid configuration at tank center under varying g-levels or perturbations propellant storage. Results presented herein include comparisons illustrating the influence of gravity, fluid volume, and the magnetic field on a paramagnetic fluid, LOX. The magnetic Bond number is utilized as predictive correlating parameter for investigating these processes. A dimensionless relationship between the Bom and Bo was sought with the goal of developing a correlation that was independent of fluid volume and tank geometry. Evidence is presented to support the hypothesis that the magnetic Bond number is an effective dimensionless parameter for modeling and understanding such systems. Further, this study supports the conclusion that magnetic storage appears to be a viable emerging technology for cryogenic propellant management systems that merits further computational investigation and space-based experimentation to establish the technology base required for future spacecraft design.

Cryogenic Fluid Technologies for Long Duration In-Space Operations

NASA Technical Reports Server (NTRS)

Motil, Susan M.; Tramel, Terri L.

2008-01-01

Reliable knowledge of low-gravity cryogenic fluid management behavior is lacking and yet is critical in the areas of storage, distribution, and low-gravity propellant management. The Vision for Space Exploration mission objectives will require the use of high performance cryogenic propellants (hydrogen, oxygen, and methane). Additionally, lunar missions will require success in storing and transferring liquid and gas commodities on the surface. The fundamental challenges associated with the in-space use of cryogens are their susceptibility to environmental heat, their complex thermodynamic and fluid dynamic behavior in low gravity and the uncertainty of the position of the liquid-vapor interface if the propellants are not settled. The Cryogenic Fluid Management (CFM) project is addressing these issues through ground testing and analytical model development, and has crosscutting applications and benefits to virtually all missions requiring in-space operations with cryogens. Such knowledge can significantly reduce or even eliminate tank fluid boil-off losses for long term missions, reduce propellant launch mass and on-orbit margins, and simplify vehicle operations. The Cryogenic Fluid Management (CFM) Project is conducting testing and performing analytical evaluation of several areas to enable NASA s Exploration Vision. This paper discusses the content and progress of the technology focus areas within CFM.

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.

Simulating the Composite Propellant Manufacturing Process

NASA Technical Reports Server (NTRS)

Williamson, Suzanne; Love, Gregory

2000-01-01

There is a strategic interest in understanding how the propellant manufacturing process contributes to military capabilities outside the United States. The paper will discuss how system dynamics (SD) has been applied to rapidly assess the capabilities and vulnerabilities of a specific composite propellant production complex. These facilities produce a commonly used solid propellant with military applications. The authors will explain how an SD model can be configured to match a specific production facility followed by a series of scenarios designed to analyze operational vulnerabilities. By using the simulation model to rapidly analyze operational risks, the analyst gains a better understanding of production complexities. There are several benefits of developing SD models to simulate chemical production. SD is an effective tool for characterizing complex problems, especially the production process where the cascading effect of outages quickly taxes common understanding. By programming expert knowledge into an SD application, these tools are transformed into a knowledge management resource that facilitates rapid learning without requiring years of experience in production operations. It also permits the analyst to rapidly respond to crisis situations and other time-sensitive missions. Most importantly, the quantitative understanding gained from applying the SD model lends itself to strategic analysis and planning.

Development of surrogate models for the prediction of the flow around an aircraft propeller

NASA Astrophysics Data System (ADS)

Salpigidou, Christina; Misirlis, Dimitris; Vlahostergios, Zinon; Yakinthos, Kyros

2018-05-01

In the present work, the derivation of two surrogate models (SMs) for modelling the flow around a propeller for small aircrafts is presented. Both methodologies use derived functions based on computations with the detailed propeller geometry. The computations were performed using k-ω shear stress transport for modelling turbulence. In the SMs, the modelling of the propeller was performed in a computational domain of disk-like geometry, where source terms were introduced in the momentum equations. In the first SM, the source terms were polynomial functions of swirl and thrust, mainly related to the propeller radius. In the second SM, regression analysis was used to correlate the source terms with the velocity distribution through the propeller. The proposed SMs achieved faster convergence, in relation to the detail model, by providing also results closer to the available operational data. The regression-based model was the most accurate and required less computational time for convergence.

MAST Propellant and Delivery System Design Methods

NASA Technical Reports Server (NTRS)

Nadeem, Uzair; Mc Cleskey, Carey M.

2015-01-01

A Mars Aerospace Taxi (MAST) concept and propellant storage and delivery case study is undergoing investigation by NASA's Element Design and Architectural Impact (EDAI) design and analysis forum. The MAST lander concept envisions landing with its ascent propellant storage tanks empty and supplying these reusable Mars landers with propellant that is generated and transferred while on the Mars surface. The report provides an overview of the data derived from modeling between different methods of propellant line routing (or "lining") and differentiate the resulting design and operations complexity of fluid and gaseous paths based on a given set of fluid sources and destinations. The EDAI team desires a rough-order-magnitude algorithm for estimating the lining characteristics (i.e., the plumbing mass and complexity) associated different numbers of vehicle propellant sources and destinations. This paper explored the feasibility of preparing a mathematically sound algorithm for this purpose, and offers a method for the EDAI team to implement.

In-situ propellant rocket engines for Mars missions ascent vehicle

NASA Technical Reports Server (NTRS)

Roncace, Elizabeth A.

1991-01-01

When contemplating the human exploration of Mars, many scenarios using various propulsion systems have been considered. One propulsion option among them is a vehicle stage with multiple, pump fed rocket engines capable of operating on propellants available on Mars. This reduces the earth launch mass requirements, resulting in economic and payload benefits. No plentiful sources of hydrogen on Mars have been identified on the surface of Mars, so most commonly used high performance liquid fuels, such as hydrogen and hydrocarbons, can be eliminated as possible in situ propellants. But 95 pct of the Martian atmosphere consists of carbon dioxide, which can be converted into carbon monoxide and oxygen. The carbon monoxide oxygen propellant combination is a candidate for a Martian in situ propellant rocket engine. The feasibility is analyzed of a pump fed engine cycle using the propellant combination of carbon monoxide and oxygen.

In-situ propellant rocket engines for Mars mission ascent vehicle

NASA Technical Reports Server (NTRS)

Roncace, Elizabeth A.

1991-01-01

When comtemplating the human exploration of Mars, many scenarios using various propulsion systems have been considered. One propulsion option among them is a vehicle stage with multiple, pump fed rocket engines capable of operating on propellants available on Mars. This reduces the Earth launch mass requirements, resulting in economic and payload benefits. No plentiful sources of hydrogen on Mars have been identified on the surface of Mars, so most commonly used high performance liquid fuels, such as hydrogen and hydrocarbons, can be eliminated as possible in-situ propellants. But 95 pct. of the Martian atmosphere consists of carbon dioxide, which can be converted into carbon monoxide and oxygen. The carbon monoxide oxygen propellant conbination is a candidate for a Martian in-situ propellant rocket engine. The feasibility is analyzed of a pump fed engine cycle using the propellant combination of carbon monoxide and oxygen.

The Negative Thrust and Torque of Several Full-scale Propellers and Their Application to Various Flight Problems

NASA Technical Reports Server (NTRS)

Hartman, Edwin P; Biermann, David

1938-01-01

Negative thrust and torque data for 2, 3, and 4-blade metal propellers having Clark y and R.A.F. 6 airfoil sections were obtained from tests in the NACA 20-foot tunnel. The propellers were mounted in front of a radial engine nacelle and the blade-angle settings covered in the tests ranged from l5 degrees to 90 degrees. One propeller was also tested at blade-angle settings of 0 degree, 5 degrees, and 10 degrees. A considerable portion of the report deals with the various applications of the negative thrust and torque to flight problems. A controllable propeller is shown to have a number of interesting, and perhaps valuable, uses within the negative thrust and torque range of operation. A small amount of engine-friction data is included to facilitate the application of the propeller data.

Operational Concept Evaluation of Solid Oxide Fuel Cells for Space Vehicle Applications

NASA Technical Reports Server (NTRS)

Poast, Kenneth I.

2011-01-01

With the end of the Space Shuttle Program, NASA is evaluating many different technologies to support future missions. Green propellants, like liquid methane and liquid oxygen, have potential advantages for some applications. A Lander propelled with LOX/methane engines is one such application. When the total vehicle design and infrastructure are considered, the advantages of the integration of propulsion, heat rejection, life support and power generation become attractive for further evaluation. Scavenged residual propellants from the propulsion tanks could be used to generate needed electric power, heat and water with a Solid Oxide Fuel Cell(SOFC). In-Situ Resource Utilization(ISRU) technologies may also generate quantities of green propellants to refill these tanks and/or supply these fuel cells. Technology demonstration projects such as the Morpheus Lander are currently underway to evaluate the practicality of such designs and operational concepts. Tethered tests are currently in progress on this vertical test bed to evaluate the propulsion and avionics systems. Evaluation of the SOFC seeks to determine the feasibility of using these green propellants to supply power and identify the limits to the integration of this technology into a space vehicle prototype.

In-Space Cryogenic Propellant Depot (ISCPD) Architecture Definitions and Systems Studies

NASA Technical Reports Server (NTRS)

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

2006-01-01

The objectives of the ISCPD Architecture Definitions and Systems Studies were to determine high leverage propellant depot architecture concepts, system configuration trades, and related technologies to enable more ambitious and affordable human and robotic exploration of the Earth Neighborhood and beyond. This activity identified architectures and concepts that preposition and store propellants in space for exploration and commercial space activities, consistent with Exploration Systems Research and Technology (ESR&T) objectives. Commonalities across mission scenarios for these architecture definitions, depot concepts, technologies, and operations were identified that also best satisfy the Vision of Space Exploration. Trade studies were conducted, technology development needs identified and assessments performed to drive out the roadmap for obtaining an in-space cryogenic propellant depot capability. The Boeing Company supported the NASA Marshall Space Flight Center (MSFC) by conducting this Depot System Architecture Development Study. The primary objectives of this depot architecture study were: (1) determine high leverage propellant depot concepts and related technologies; (2) identify commonalities across mission scenarios of depot concepts, technologies, and operations; (3) determine the best depot concepts and key technology requirements and (4) identify technology development needs including definition of ground and space test article requirements.

The design and development of a miniature bi-stable latching solenoid valve for low thrust resistojets

NASA Technical Reports Server (NTRS)

Lynch, R. A.

1972-01-01

The design approach and development demonstration are presented for a flightweight, miniature, magnetic latching shutoff valve, suitable for use with the ruggedized H2 and NH3 resistojet and the biowaste resistojet. The design criteria established provided for compatibility with specified ruggedized resistojet propellants as well as the biowaste and other propulsion system propellants.

A Navier-Stokes Solution of Hull-Ring Wing-Thruster Interaction

NASA Technical Reports Server (NTRS)

Yang, C.-I.; Hartwich, P.; Sundaram, P.

1991-01-01

Navier-Stokes simulations of high Reynolds number flow around an axisymmetric body supported in a water tunnel were made. The numerical method is based on a finite-differencing high resolution second-order accurate implicit upwind scheme. Four different configurations were investigated, these are: (1) barebody; (2) body with an operating propeller; (3) body with a ring wing; and (4) body with a ring wing and an operating propeller. Pressure and velocity components near the stern region were obtained computationally and are shown to compare favorably with the experimental data. The method correctly predicts the existence and extent of stern flow separation for the barebody and the absence of flow separation for the three other configurations with ring wing and/or propeller.

Aeroacoustic effects of reduced aft tip speed at constant thrust for a model counterrotation turboprop at takeoff conditions

NASA Technical Reports Server (NTRS)

Woodward, Richard P.; Hughes, Christopher E.

1990-01-01

A model high-speed, advanced counterrotation propeller, F7/A7, was tested in the anechoic wind tunnel at simulated takeoff and approach conditions of Mach 0.2. The propeller was operated in a baseline configuration with the forward and aft rotor blade setting angles and forward and aft rotational speeds essentially equal. Two additional configurations were tested with the aft rotor at increased blade setting angles and the rotational speed reduced to achieve overall performance similar to that of the baseline configuration. Acoustic data were taken with an axially translating microphone probe that was attached to the tunnel floor. Concurrent aerodynamic data were taken to define propeller operating conditions.

A Lifting-Surface Program for Contrarotating Propellers

DTIC Science & Technology

1989-04-01

computer program for a set the force and induced flow field could be determined . of CR propellers has been developed based on a modi- The variations of...and their methods. To determine the forces and induced Nelson’s lifting life programs use the same approach, flow field, they applied lifting-line...Velocimetry (LDV). The propeller set, designed wake should be exactly the same as the hub geometry, to operate in uniform flow , was tested in the DTRC The

The Influence of Forward Flight on Propeller Noise

NASA Technical Reports Server (NTRS)

Magliozzi, B.

1977-01-01

The effect of flight on blade surface pressures and propeller noise was reported. There were significant differences in blade surface pressures and far-field noise between static and flight conditions. The static data showed many high-intensity, tone-like peaks whereas the flight data was generally free from tones. The turbulence ingested by the propeller operating statically was dominated by long, thin eddies. In flight the scale of the turbulence was greately reduced from that observed statically.

Propellant for the NASA Standard Initiator

NASA Technical Reports Server (NTRS)

Hohmann, Carl; Tipton, Bill, Jr.; Dutton, Maureen

2000-01-01

This paper discusses processes employed in manufacturing zirconium-potassium perchlorate propellant for the NASA standard initiator. It provides both a historical background on the NSI device-detailing problem areas and their resolution--and on propellant blending techniques. Emphasis is placed on the precipitation blending method. The findings on mixing equipment, processing, and raw materials are described. Also detailed are findings on the bridgewire slurry operation, one of the critical steps in the production of the NASA standard initiator.

49 CFR 218.93 - Definitions.

Code of Federal Regulations, 2014 CFR

2014-10-01

... button or radio control when such switch is protected by distant switch indicators, switch point... units are connected so that they may be operated from a single control stand. Locomotive means, for... one or more propelling motors designed for moving other equipment; (2) With one or more propelling...

Robust Exploration and Commercial Missions to the Moon Using Nuclear Thermal Rocket Propulsion and Lunar Liquid Oxygen Derived from FeO-Rich Pyroclasitc Deposits

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Ryan, Stephen W.; Burke, Laura M.; McCurdy, David R.; Fittje, James E.; Joyner, Claude R.

2018-01-01

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable access through cislunar space necessary for commercial development and sustained human presence on the Moon. It is a demonstrated technology capable of generating both high thrust and high specific impulse (I(sub sp) approx. 900 s) twice that of today's best chemical rockets. Nuclear lunar transfer vehicles-consisting of a propulsion stage using three approx. 16.5-klb(sub f) small nuclear rocket engines (SNREs), an in-line propellant tank, plus the payload-are reusable, enabling a variety of lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong ''tourism'' missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing a robust in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in-situ propellant production used to supply strategically located propellant depots and transportation nodes. The use of lunar liquid oxygen (LLO2) derived from iron oxide (FeO)-rich volcanic glass beads, found in numerous pyroclastic deposits on the Moon, can significantly reduce the launch mass requirements from Earth by enabling reusable, surface-based lunar landing vehicles (LLVs)that use liquid oxygen and hydrogen (LO2/LH2) chemical rocket engines. Afterwards, a LO2/LH2 propellant depot can be established in lunar equatorial orbit to supply the LTS. At this point a modified version of the conventional NTR-called the LO2-augmented NTR, or LANTR-is introduced into the LTS allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an ''afterburner'' into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engine's choked sonic throat-essentially ''scramjet propulsion in reverse.'' By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and I(sub sp) values while the reactor core power level remains relatively constant. A LANTR-based LTS offers unique mission capabilities including short-transit-time crewed cargo transports. Even a ''commuter'' shuttle service may be possible allowing ''one-way'' trip times to and from the Moon on the order of 36 hours or less. If only 1% of the extracted LLO2 propellant from identified resource sites were available for use in lunar orbit, such a supply could support daily commuter flights to the Moon for many thousands of years! This report outlines an evolutionary architecture and examines a variety of mission types and transfer vehicle designs, along with the increasing demands on LLO2 production as mission complexity and velocity change delta V requirements increase. A comparison of vehicle features and engine operating characteristics, for both NTR and LANTR engines, is also provided along with a discussion of the propellant production and mining requirements associated with using FeO-rich volcanic glass as source material.

Robust Exploration and Commercial Missions to the Moon Using LANTR Propulsion and Lunar Liquid Oxygen Derived from FeO-Rich Pyroclastic Deposits

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Ryan, Stephen W.; Burke, Laura M.; McCurdy, David R.; Fittje, James E.; Joyner, Claude R.

2017-01-01

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable access through cislunar space necessary for commercial development and sustained human presence on the Moon. It is a demonstrated technology capable of generating both high thrust and high specific impulse (Isp approx.900 s) twice that of todays best chemical rockets. Nuclear lunar transfer vehicles consisting of a propulsion stage using three approx.16.5 klbf Small Nuclear Rocket Engines (SNREs), an in-line propellant tank, plus the payload can enable a variety of reusable lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong tourism missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing a robust in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in-situ propellant production used to supply strategically located propellant depots and transportation nodes. The use of lunar liquid oxygen (LLO2) derived from iron oxide (FeO)-rich volcanic glass beads, found in numerous pyroclastic deposits on the Moon, can significantly reduce the launch mass requirements from Earth by enabling reusable, surface-based lunar landing vehicles (LLVs) using liquid oxygen/hydrogen (LO2/H2) chemical rocket engines. Afterwards, a LO2/H2 propellant depot can be established in lunar equatorial orbit to supply the LTS. At this point a modified version of the conventional NTR called the LOX-augmented NTR, or LANTR is introduced into the LTS allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an afterburner into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engines choked sonic throat - essentially scramjet propulsion in reverse. By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and Isp values while the reactor core power level remains relatively constant. A LANTR-based LTS offers unique mission capabilities including short transit time crewed cargo transports. Even a commuter shuttle service may be possible allowing one-way trip times to and from the Moon on the order of 36 hours or less. If only 1 of the extracted LLO2 propellant from identified resource sites were available for use in lunar orbit, such a supply could support daily commuter flights to the Moon for many thousands of years! The proposed paper outlines an evolutionary architecture and examines a variety of mission types and transfer vehicle designs, along with the increasing demands on LLO2 production as mission complexity and (Delta)V requirements increase. A comparison of vehicle features and engine operating characteristics, for both NTR and LANTR engines, is also provided along with a discussion of the propellant production and mining requirements associated with using FeO-rich volcanic glass as source material.

Electrostatic Evaluation of the Propellant Handlers Ensemble

NASA Technical Reports Server (NTRS)

Hogue, Michael D.; Calle, Carlos I.; Buhler, Charles

2006-01-01

The Self-Contained Atmospheric Protective Ensemble (SCAPE) used in propellant handling at NASA's Kennedy Space Center (KSC) has recently completed a series of tests to determine its electrostatic properties of the coverall fabric used in the Propellant Handlers Ensemble (PHE). Understanding these electrostatic properties are fundamental to ensuring safe operations when working with flammable rocket propellants such as hydrazine, methyl hydrazine, and unsymmetrical dimethyl hydrazine. These tests include surface resistivity, charge decay, triboelectric charging, and flame incendivity. In this presentation, we will discuss the results of these tests on the current PHE as well as new fabrics and materials being evaluated for the next generation of PHE.

The Aerodynamic Characteristics of Six Full-Scale Propellers Having Different Airfoil Sections

NASA Technical Reports Server (NTRS)

Biermann, David; Hartman, Edwin P

1939-01-01

Wind-tunnel tests are reported of six 3-blade 10-foot propellers operated in front of a liquid-cooled engine nacelle. The propellers were identical except for blade airfoil sections, which were: Clark y, R.A.F. 6, NACA 4400, NACA 2400-34, NACA 2rsub200, and NACA 6400. The range of blade angles investigated extended for 15 degrees to 40 degrees for all propellers except the Clark y, for which it extended to 45 degrees. The results showed that the range in maximum efficiency between the highest and lowest values was about 3 percent. The highest efficiencies were for the low-camber sections.

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.

Langley Full-Scale Tunnel Investigation of a 1/3-Scale Model of the Chance Vought XF5U-1 Airplane

NASA Technical Reports Server (NTRS)

Lange, Roy H.; Cocke, Bennie W., Jr.; Proterra, Anthony J.

1946-01-01

The results of an investigation of a 1/3-scale model of the Chance Vought XF5U-1 airplane in the Langley full-scale tunnel are presented in this report. The maximum lift and stalling characteristics of several model configurations, the longitudinal stability characteristics of the model, and the effectiveness of the control surfaces were determined with the propellers removed. The propulsive characteristics, the effect of propeller operation on the lift, and the static thrust of the model propellers were determined at several propeller-blade angles. The results with the propellers removed showed that the maximum lift coefficient of the complete model configuration was only 0.97 was compared with the value of 1.31 for the model configuration in which the engine-air ducts and canopy are removed. The model with the propellers removed (normal center-of-gravity position) has a positive static margin, stick fixed, varying from 5 to 13 percent of the mean aerodynamic chord throughout the unstalled range of lift coefficients. The unit horizontal tail is sufficiently powerful to trim the airplane with the propellers removed throughout the unstalled range of lift coefficients. The peak propulsive efficiencies for beta = 20 degrees and beta = 30 degrees were increased 7 percent at C(sub L) congruent to 0.67 and 20 percent at C(sub L) congruent to 0.74, respectively, with the propellers rotating upward in the center than with the propellers rotating downward in the center. Indications are that the minimum forward-flight speed of the airplane for full-power operation at sea level will be about 90 miles per hour. Decreasing the weight and increasing the power reduced this value of minimum speed and there were no indications from the results of a lower limit to the minimum speed.

In-space propellant logistics. Volume 4: Project planning data

NASA Technical Reports Server (NTRS)

1972-01-01

The prephase A conceptual project planning data as it pertains to the development of the selected logistics module configuration transported into earth orbit by the space shuttle orbiter. The data represents the test, implementation, and supporting research and technology requirements for attaining the propellant transfer operational capability for early 1985. The plan is based on a propellant module designed to support the space-based tug with cryogenic oxygen-hydrogen propellants. A logical sequence of activities that is required to define, design, develop, fabricate, test, launch, and flight test the propellant logistics module is described. Included are the facility and ground support equipment requirements. The schedule of activities are based on the evolution and relationship between the R and T, the development issues, and the resultant test program.

Development of advanced inert-gas ion thrusters

NASA Technical Reports Server (NTRS)

Poeschel, R. L.

1983-01-01

Inert gas ion thruster technology offers the greatest potential for providing high specific impulse, low thrust, electric propulsion on large, Earth orbital spacecraft. The development of a thruster module that can be operated on xenon or argon propellant to produce 0.2 N of thrust at a specific impulse of 3000 sec with xenon propellant and at 6000 sec with argon propellant is described. The 30 cm diameter, laboratory model thruster is considered to be scalable to produce 0.5 N thrust. A high efficiency ring cusp discharge chamber was used to achieve an overall thruster efficiency of 77% with xenon propellant and 66% with argon propellant. Measurements were performed to identify ion production and loss processes and to define critical design criteria (at least on a preliminary basis).

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.

Ion-thruster propellant utilization

NASA Technical Reports Server (NTRS)

Kaufman, H. R.

1971-01-01

The evaluation and understanding of maximum propellant utilization, with mercury used as the propellant are presented. The primary-electron region in the ion chamber of a bombardment thruster is analyzed at maximum utilization. The results of this analysis, as well as experimental data from a range of ion-chamber configurations, show a nearly constant loss rate for unionized propellant at maximum utilization over a wide range of total propellant flow rate. The discharge loss level of 1000 eV/ion was used as a definition of maximum utilization, but the exact level of this definition has no effect on the qualitative results and little effect on the quantitative results. There are obvious design applications for the results of this investigation, but the results are particularly significant whenever efficient throttled operation is required.

Space Station fluid management logistics

NASA Technical Reports Server (NTRS)

Dominick, Sam M.

1990-01-01

Viewgraphs and discussion on space station fluid management logistics are presented. Topics covered include: fluid management logistics - issues for Space Station Freedom evolution; current fluid logistics approach; evolution of Space Station Freedom fluid resupply; launch vehicle evolution; ELV logistics system approach; logistics carrier configuration; expendable fluid/propellant carrier description; fluid carrier design concept; logistics carrier orbital operations; carrier operations at space station; summary/status of orbital fluid transfer techniques; Soviet progress tanker system; and Soviet propellant resupply system observations.

Space Storable Rocket Technology (SSRT) basic program

NASA Technical Reports Server (NTRS)

Chazen, M. L.; Mueller, T.; Casillas, A. R.; Huang, D.

1992-01-01

The Space Storable Rocket Technology Program (SSRT) was conducted to establish a technology for a new class of high performance and long life bipropellant engines using space storable propellants. The results are described. Task 1 evaluated several characteristics for a number of fuels to determine the best space storable fuel for use with LO2. The results indicated that LO2-N2H4 is the best propellant combination and provides the maximum mission/system capability maximum payload into GEO of satellites. Task 2 developed two models, performance and thermal. The performance model indicated the performance goal of specific impulse greater than or = 340 seconds (sigma = 204) could be achieved. The thermal model was developed and anchored to hot fire test data. Task 3 consisted of design, fabrication, and testing of a 200 lbf thrust test engine operating at a chamber pressure of 200 psia using LO2-N2H4. A total of 76 hot fire tests were conducted demonstrating performance greater than 340 (sigma = 204) which is a 25 second specific impulse improvement over the existing highest performance flight apogee type engines.

Analysis of a Radioisotope Thermal Rocket Engine

NASA Technical Reports Server (NTRS)

Machado-Rodriguez, Jonathan P.; Landis, Geoffrey A.

2017-01-01

The Triton Hopper is a concept for a vehicle to explore the surface of Neptunes moon Triton, which uses a radioisotope heated rocket engine and in-situ propellant acquisition. The initial Triton Hopper conceptual design stores pressurized Nitrogen in a spherical tank to be used as the propellant. The aim of the research was to investigate the benefits of storing propellant at ambient temperature and heating it through a thermal block during engine operation, as opposed to storing gas at a high temperature.

Investigation of the NACA 4-(3)(8)-045 Two-blade Propellers at Forward Mach Numbers to 0.725 to Determine the Effects of Compressibility and Solidity on Performance

NASA Technical Reports Server (NTRS)

Stack, John; Draley, Eugene C; Delano, James B; Feldman, Lewis

1950-01-01

As part of a general investigation of propellers at high forward speeds, tests of two 2-blade propellers having the NACA 4-(3)(8)-03 and NACA 4-(3)(8)-45 blade designs have been made in the Langley 8-foot high-speed tunnel through a range of blade angle from 20 degrees to 60 degrees for forward Mach numbers from 0.165 to 0.725 to establish in detail the changes in propeller characteristics due to compressibility effects. These propellers differed primarily only in blade solidity, one propeller having 50 percent and more solidity than the other. Serious losses in propeller efficiency were found as the propeller tip Mach number exceeded 0.91, irrespective of forward speed or blade angle. The magnitude of the efficiency losses varied from 9 percent to 22 percent per 0.1 increase in tip Mach number above the critical value. The range of advance ratio for peak efficiency decreased markedly with increase of forward speed. The general form of the changes in thrust and power coefficients was found to be similar to the changes in airfoil lift coefficient with changes in Mach number. Efficiency losses due to compressibility effects decreased with increase of blade width. The results indicated that the high level of propeller efficiency obtained at low speeds could be maintained to forward sea-level speeds exceeding 500 miles per hour.

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.

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.

An Experimental Investigation of the Effect of Propellers Used as Aerodynamic Brakes on Stability and Control

NASA Technical Reports Server (NTRS)

Hanson, Frederick H

1945-01-01

Tests were made of a model representative of a single-engine tractor-type airplane for the purpose of determining the stability and control effects of a propeller used as an aerodynamic brake. The tests were made with single-and dual-rotation propellers to show the effect of type of propeller rotation, and with positive thrust to provide basic data with which to compare the effects of negative thrust. Four configurations of the model were used to give the effects of tilting the propeller thrust axis down 5 deg., raising the horizontal tail, and combining both tilt and raised tail. Results of the tests are reported herein. The effects of negative thrust were found to be significant. The longitudinal stability was increased because of the loss of wing lift and increase of the angle of attack of the tail. Directional stability and both longitudinal and directional control were decreased because of the reduced velocity at the tail. These effects are moderate for moderate braking but become pronounced with full-power braking, particularly at high values of lift coefficient. The effects of model configuration changes were small when compared with the over-all effects of negative-thrust operation; however, improved stability and control characteristics were exhibited by the model with the tilted thrust axis. Raising the horizontal tail improved the longitudinal characteristics, but was detrimental to directional characteristics. The use of dual-rotation propeller reduced the directional trim charges resulting from the braking operation. A prototype airplane was assumed and handling qualities were computed and analyzed for normal (positive thrust) and braking operation with full and partial power. The results of these analyses are presented for the longitudinal characteristics in steady and accelerated flight, and for the directional characteristics in high- and low-speed flight. It was found that by limiting the power output of the engine (assuming the constant-speed propeller will function in the range of blade angles required for negative thrust) the stability and control characteristics may be held within the limits required for safe operation. Braking with full power, particularly at low speeds, is dangerous, but braking with very small power output is satisfactory from the standpoint of control. The amount of braking produced with zero power output is equal to or better than that produced by conventional spoiler-type brakes.

Turboprop Propulsion Mechanic 2-8. Military Curriculum Materials for Vocational and Technical Education.

ERIC Educational Resources Information Center

Ohio State Univ., Columbus. National Center for Research in Vocational Education.

These military-developed curriculum materials for turboprop propulsion mechanics are targeted for use in grades 11-adult. Organized in five instructional blocks, the materials deal with the following topics: fundamentals of turboprop propulsion mechanics; engine and propeller systems operation; propeller maintenance; engine repair; and engine…

14 CFR 21.500 - Approval of engines and propellers.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Approval of engines and propellers. 21.500 Section 21.500 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION... propeller— (a) Conforms to its U.S. type certificate and is in condition for safe operation; and (b) Has...

Space operations center: Shuttle interaction study extension, executive summary

NASA Technical Reports Server (NTRS)

1982-01-01

The Space Operations Center (SOC) is conceived as a permanent facility in low Earth orbit incorporating capabilities for space systems construction; space vehicle assembly, launching, recovery and servicing; and the servicing of co-orbiting satellites. The Shuttle Transportation System is an integral element of the SOC concept. It will transport the various elements of the SOC into space and support the assembly operation. Subsequently, it will regularly service the SOC with crew rotations, crew supplies, construction materials, construction equipment and components, space vehicle elements, and propellants and spare parts. The implications to the SOC as a consequence of the Shuttle supporting operations are analyzed. Programmatic influences associated with propellant deliveries, spacecraft servicing, and total shuttle flight operations are addressed.

Propellant Analysis and Distillation Unit Design

NASA Technical Reports Server (NTRS)

Barragan, Michelle H.; Spangler, Cindy; Barrera, Louis K.

2007-01-01

The NASA White Sands Test Facility (WSTF) routinely operates hypergolic propulsion systems. Some of the onsite activities include performing long duration studies on the operational life of these systems. A few of them have been in use for over twenty years. During this span of time contamination has built up in the propellant and some of the distribution infrastructure. This study investigated the nature of this contamination, the pathology of its generation, and developed a process for removal of the contamination that was cost efficient with minimal waste generation.

Evaluation of Aero Commander sidewall vibration and interior acoustic data: Static operations

NASA Technical Reports Server (NTRS)

Piersol, A. G.; Wilby, E. G.; Wilby, J. F.

1980-01-01

Results for the vibration measured at five locations on the fuselage structure during static operations are presented. The analysis was concerned with the magnitude of the vibration and the relative phase between different locations, the frequency response (inertance) functions between the exterior pressure field and the vibration, and the coherent output power functions at interior microphone locations based on sidewall vibration. Fuselage skin panels near the plane of rotation of the propeller accept propeller noise excitation more efficiently than they do exhaust noise.

A Simple Method for High-Lift Propeller Conceptual Design

NASA Technical Reports Server (NTRS)

Patterson, Michael; Borer, Nick; German, Brian

2016-01-01

In this paper, we present a simple method for designing propellers that are placed upstream of the leading edge of a wing in order to augment lift. Because the primary purpose of these "high-lift propellers" is to increase lift rather than produce thrust, these props are best viewed as a form of high-lift device; consequently, they should be designed differently than traditional propellers. We present a theory that describes how these props can be designed to provide a relatively uniform axial velocity increase, which is hypothesized to be advantageous for lift augmentation based on a literature survey. Computational modeling indicates that such propellers can generate the same average induced axial velocity while consuming less power and producing less thrust than conventional propeller designs. For an example problem based on specifications for NASA's Scalable Convergent Electric Propulsion Technology and Operations Research (SCEPTOR) flight demonstrator, a propeller designed with the new method requires approximately 15% less power and produces approximately 11% less thrust than one designed for minimum induced loss. Higher-order modeling and/or wind tunnel testing are needed to verify the predicted performance.

The Effect of an Operating Propeller on the Aerodynamic Characteristics of a 1/10-Scale Model of the Lockheed XFV-1 Airplane at High Subsonic Speeds (TED No. NACA DE-377)

NASA Technical Reports Server (NTRS)

Sutton, Fred B.; Buell, Donald A.

1952-01-01

An investigation was conducted in the Ames 12-foot pressure wind tunnel to determine the effect of an operating propeller on the aerodynamic characteristics of a l/l9-scale model of the Lockheed XFV-1 airplane, Several full-scale power conditions were simulated at Mach numbers from 0.50 to 0.92; the.Reynolds number was constant at 1,7 million. Lift, longitudinal force, pitch, roll, and yaw characteristics, determined with and without power, are presented for the complete model and for various combinations of model components, Results of an investigation to determine the characteristics of the dual-rotating propeller used on the model are given also,

Research on the time-temperature-damage superposition principle of NEPE propellant

NASA Astrophysics Data System (ADS)

Han, Long; Chen, Xiong; Xu, Jin-sheng; Zhou, Chang-sheng; Yu, Jia-quan

2015-11-01

To describe the relaxation behavior of NEPE (Nitrate Ester Plasticized Polyether) propellant, we analyzed the equivalent relationships between time, temperature, and damage. We conducted a series of uniaxial tensile tests and employed a cumulative damage model to calculate the damage values for relaxation tests at different strain levels. The damage evolution curve of the tensile test at 100 mm/min was obtained through numerical analysis. Relaxation tests were conducted over a range of temperature and strain levels, and the equivalent relationship between time, temperature, and damage was deduced based on free volume theory. The equivalent relationship was then used to generate predictions of the long-term relaxation behavior of the NEPE propellant. Subsequently, the equivalent relationship between time and damage was introduced into the linear viscoelastic model to establish a nonlinear model which is capable of describing the mechanical behavior of composite propellants under a uniaxial tensile load. The comparison between model prediction and experimental data shows that the presented model provides a reliable forecast of the mechanical behavior of propellants.

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.

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.

High-power and 2.5 kW advanced-technology ion thruster

NASA Technical Reports Server (NTRS)

Poeschel, R. L.

1977-01-01

Investigations for improving ion thruster components in the 30 cm engineering model thruster (EMT) resulted in the demonstration of useful techniques for grid short removal and discharge chamber erosion monitoring, establishment of relationships between double ion production and thruster operating parameters, verification of satisfactory specifications on porous tungsten vaporizer material and barium impregnated porous tungsten inserts, demonstration of a new hollow cathode configuration, and specification of magnetic circuit requirements for reproducing desired magnetic mappings. The capacity of a 30 cm EMT to operate at higher beam voltages and currents (higher power) was determined. Operation at 2 A beam current and higher beam voltage is shown to be essentially equivalent to operation at 1.1 kV with regard to efficiency, lifetime and operating conditions. The only additional requirement is an improvement in high voltage insulation and propellant isolator capacity. Operation at minimum voltage and higher beam currents is shown to increase thruster discharge chamber erosion in proportion to beam current. Studies to find alternatives to molybdenum for manufacturing ion optics grids are also reported.

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.

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.

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.

Method and apparatus for duct sealing using a clog-resistant insertable injector

DOEpatents

Wang, Duo; Modera, Mark P.

2007-01-02

A clog-resistant injector spray nozzle allows relatively unobtrusive insertion through a small access aperture into existing ductwork in occupied buildings for atomized particulate sealing of a ductwork. The spray nozzle comprises an easily cleaned and easily replaced straight liquid tube whose liquid contents are principally propelled by a heated propellant gas, such as heated air. Heat transfer is minimized from the heated propellant gas to the liquid tube until they both exit the injector, thereby greatly reducing the likelihood of nozzle clogging. A method of duct sealing using particles driven by heated propellant gas is described, whereby duct-sealing operations become both faster, and commercially practicable in inhabited commercial and residential buildings.

Optimization design of submerged propeller in oxidation ditch by computational fluid dynamics and comparison with experiments.

PubMed

Zhang, Yuquan; Zheng, Yuan; Fernandez-Rodriguez, E; Yang, Chunxia; Zhu, Yantao; Liu, Huiwen; Jiang, Hao

The operating condition of a submerged propeller has a significant impact on flow field and energy consumption of the oxidation ditch. An experimentally validated numerical model, based on the computational fluid dynamics (CFD) tool, is presented to optimize the operating condition by considering two important factors: flow field and energy consumption. Performance demonstration and comparison of different operating conditions were carried out in a Carrousel oxidation ditch at the Yingtang wastewater treatment plants in Anhui Province, China. By adjusting the position and rotating speed together with the number of submerged propellers, problems of sludge deposit and the low velocity in the bend could be solved in a most cost-effective way. The simulated results were acceptable compared with the experimental data and the following results were obtained. The CFD model characterized flow pattern and energy consumption in the full-scale oxidation ditch. The predicted flow field values were within -1.28 ± 7.14% difference from the measured values. By determining three sets of propellers under the rotating speed of 6.50 rad/s with one located 5 m from the first curved wall, after numerical simulation and actual measurement, not only the least power density but also the requirement of the flow pattern could be realized.

Electromagnetic propulsion for spacecraft

NASA Technical Reports Server (NTRS)

Myers, Roger M.

1993-01-01

Three electromagnetic propulsion technologies, solid propellant pulsed plasma thrusters (PPT), magnetoplasmadynamic (MPD) thrusters, and pulsed inductive thrusters (PIT), were developed for application to auxiliary and primary spacecraft propulsion. Both the PPT and MPD thrusters were flown in space, though only PPT's were used on operational satellites. The performance of operational PPT's is quite poor, providing only approximately 8 percent efficiency at approximately 1000 s specific impulse. However, laboratory PPT's yielding 34 percent efficiency at 2000 s specific impulse were extensively tested, and peak performance levels of 53 percent efficiency at 5170 s specific impulse were demonstrated. MPD thrusters were flown as experiments on the Japanese MS-T4 spacecraft and the Space Shuttle and were qualified for a flight in 1994. The flight MPD thrusters were pulsed, with a peak performance of 22 percent efficiency at 2500 s specific impulse using ammonia propellant. Laboratory MPD thrusters were demonstrated with up to 70 percent efficiency and 700 s specific impulse using lithium propellant. While the PIT thruster has never been flown, recent performance measurements using ammonia and hydrazine propellants are extremely encouraging, reaching 50 percent efficiency for specific impulses between 4000 to 8000 s. The fundamental operating principles, performance measurements, and system level design for the three types of electromagnetic thrusters are reviewed, and available data on flight tests are discussed for the PPT and MPD thrusters.

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.

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.

Cryogenic Fluid Management Technology for Moon and Mars Missions

NASA Technical Reports Server (NTRS)

Doherty, Michael P.; Gaby, Joseph D.; Salerno, Louis J.; Sutherlin, Steven G.

2010-01-01

In support of the U.S. Space Exploration Policy, focused cryogenic fluid management technology efforts are underway within the National Aeronautics and Space Administration. Under the auspices of the Exploration Technology Development Program, cryogenic fluid management technology efforts are being conducted by the Cryogenic Fluid Management Project. Cryogenic Fluid Management Project objectives are to develop storage, transfer, and handling technologies for cryogens to support high performance demands of lunar, and ultimately, Mars missions in the application areas of propulsion, surface systems, and Earth-based ground operations. The targeted use of cryogens and cryogenic technologies for these application areas is anticipated to significantly reduce propellant launch mass and required on-orbit margins, to reduce and even eliminate storage tank boil-off losses for long term missions, to economize ground pad storage and transfer operations, and to expand operational and architectural operations at destination. This paper organizes Cryogenic Fluid Management Project technology efforts according to Exploration Architecture target areas, and discusses the scope of trade studies, analytical modeling, and test efforts presently underway, as well as future plans, to address those target areas. The target areas are: liquid methane/liquid oxygen for propelling the Altair Lander Ascent Stage, liquid hydrogen/liquid oxygen for propelling the Altair Lander Descent Stage and Ares V Earth Departure Stage, liquefaction, zero boil-off, and propellant scavenging for Lunar Surface Systems, cold helium and zero boil-off technologies for Earth-Based Ground Operations, and architecture definition studies for long term storage and on-orbit transfer and pressurization of LH2, cryogenic Mars landing and ascent vehicles, and cryogenic production via in situ resource utilization on Mars.

Advanced APS impacts on vehicle payloads

NASA Technical Reports Server (NTRS)

Schneider, Steven J.; Reed, Brian D.

1989-01-01

Advanced auxiliary propulsion system (APS) technology has the potential to both, increase the payload capability of earth-to-orbit (ETO) vehicles by reducing APS propellant mass, and simplify ground operations and logistics by reducing the number of fluids on the vehicle and eliminating toxic, corrosive propellants. The impact of integrated cryogenic APS on vehicle payloads is addressed. In this system, launch propulsion system residuals are scavenged from integral launch propulsion tanks for use in the APS. Sufficient propellant is preloaded into the APS to return to earth with margin and noncomplete scavenging assumed. No propellant conditioning is required by the APS, but ambient heat soak is accommodated. High temperature rocket materials enable the use of the unconditioned hydrogen/oxygen in the APS and are estimated to give APS rockets specific impulse of up to about 444 sec. The payload benefits are quantified and compared with an uprated monomethylhydrazine/nitrogen tetroxide system in a conservative fashion, by assuming a 25.5 percent weight growth for the hydrogen/oxygen system and a 0 percent weight growth for the uprated system. The combination of scavenging and high performance gives payload impacts which are highly mission specific. A payload benefit of 861 kg (1898 lbm) was estimated for a Space Station Freedom rendezvous mission and 2099 kg (4626 lbm) for a sortie mission, with payload impacts varying with the amount of launch propulsion residual propellants. Missions without liquid propellant scavenging were estimated to have payload penalties, however, operational benefits were still possible.

Advanced APS Impacts on Vehicle Payloads

NASA Technical Reports Server (NTRS)

Schneider, Steven J.; Reed, Brian D.

1989-01-01

Advanced auxiliary propulsion system (APS) technology has the potential to both, increase the payload capability of earth-to-orbit (ETO) vehicles by reducing APS propellant mass, and simplify ground operations and logistics by reducing the number of fluids on the vehicle and eliminating toxic, corrosive propellants. The impact of integrated cryogenic APS on vehicle payloads is addressed. In this system, launch propulsion system residuals are scavenged from integral launch propulsion tanks for use in the APS. Sufficient propellant is preloaded into the APS to return to earth with margin and noncomplete scavenging assumed. No propellant conditioning is required by the APS, but ambient heat soak is accommodated. High temperature rocket materials enable the use of the unconditioned hydrogen/oxygen in the APS and are estimated to give APS rockets specific impulse of up to about 444 sec. The payload benefits are quantified and compared with an uprated monomethyl hydrazine/nitrogen tetroxide system in a conservative fashion, by assuming a 25.5 percent weight growth for the hydrogen/oxygen system and a 0 percent weight growth for the uprated system. The combination and scavenging and high performance gives payload impacts which are highly mission specific. A payload benefit of 861 kg (1898 lbm) was estimated for a Space Station Freedom rendezvous mission and 2099 kg (4626 lbm) for a sortie mission, with payload impacts varying with the amount of launch propulsion residual propellants. Missions without liquid propellant scavenging were estimated to have payload penalties, however, operational benefits were still possible.

Pressure-Volume-Temperature (PVT) Gauging of an Isothermal Cryogenic Propellant Tank Pressurized with Gaseous Helium

NASA Technical Reports Server (NTRS)

VanDresar, Neil T.; Zimmerli, Gregory A.

2014-01-01

Results are presented for pressure-volume-temperature (PVT) gauging of a liquid oxygen/liquid nitrogen tank pressurized with gaseous helium that was supplied by a high-pressure cryogenic tank simulating a cold helium supply bottle on a spacecraft. The fluid inside the test tank was kept isothermal by frequent operation of a liquid circulation pump and spray system, and the propellant tank was suspended from load cells to obtain a high-accuracy reference standard for the gauging measurements. Liquid quantity gauging errors of less than 2 percent of the tank volume were obtained when quasi-steady-state conditions existed in the propellant and helium supply tanks. Accurate gauging required careful attention to, and corrections for, second-order effects of helium solubility in the liquid propellant plus differences in the propellant/helium composition and temperature in the various plumbing lines attached to the tanks. On the basis of results from a helium solubility test, a model was developed to predict the amount of helium dissolved in the liquid as a function of cumulative pump operation time. Use of this model allowed correction of the basic PVT gauging calculations and attainment of the reported gauging accuracy. This helium solubility model is system specific, but it may be adaptable to other hardware systems.

An Investigation into the Potential Benefits of Distributed Electric Propulsion on Small UAVs at Low Reynolds Numbers

NASA Astrophysics Data System (ADS)

Baris, Engin

Distributed electric propulsion systems benefit from the inherent scale independence of electric propulsion. This property allows the designer to place multiple small electric motors along the wing of an aircraft instead of using a single or several internal combustion motors with gear boxes or other power train components. Aircraft operating at low Reynolds numbers are ideal candidates for benefiting from increased local flow velocities as provided by distributed propulsion systems. In this study, a distributed electric propulsion system made up of eight motor/propellers was integrated into the leading edge of a small fixed wing-body model to investigate the expected improvements on the aerodynamics available to small UAVs operating at low Reynolds numbers. Wind tunnel tests featuring a Design of Experiments (DOE) methodology were used for aerodynamic characterization. Experiments were performed in four modes: all-propellers-on, wing-tip-propellers-alone-on, wing-alone mode, and two-inboard-propellers-on-alone mode. In addition, the all-propeller-on, wing-alone, and a single-tractor configuration were analyzed using VSPAERO, a vortex lattice code, to make comparisons between these different configurations. Results show that the distributed propulsion system has higher normal force, endurance, and range features, despite a potential weight penalty.

Studies on an aerial propellant transfer space plane (APTSP)

NASA Astrophysics Data System (ADS)

Jayan, N.; Biju Kumar, K. S.; Gupta, Anish Kumar; Kashyap, Akhilesh Kumar; Venkatraman, Kartik; Mathew, Joseph; Mukunda, H. S.

2004-04-01

This paper presents a study of a fully reusable earth-to-orbit launch vehicle concept with horizontal take-off and landing, employing a turbojet engine for low speed, and a rocket for high-speed acceleration and space operations. This concept uses existing technology to the maximum possible extent, thereby reducing development time, cost and effort. It uses the experience in aerial filling of military aircrafts for propellant filling at an altitude of 13 km at a flight speed of M=0.85. Aerial filling of propellant reduces the take-off weight significantly thereby minimizing the structural weight of the vehicle. The vehicle takes off horizontally and uses turbojet engines till the end of the propellant filling operation. The rocket engines provide thrust for the next phase till the injection of a satellite at LEO. A sensitivity analysis of the mission with respect to rocket engine specific impulse and overall vehicle structural factor is also presented in this paper. A conceptual design of space plane with a payload capability of 10 ton to LEO is carried out. The study shows that the realization of an aerial propellant transfer space plane is possible with limited development of new technology thus reducing the demands on the finances required for achieving the objectives.

Low noise propeller design using numerical analysis

NASA Astrophysics Data System (ADS)

Humpert, Bryce

The purpose of this study is to explore methods for reducing aircraft propeller noise with minimal losses in performance. Using numerical analysis, a standard two blade propeller configuration was taken from experiments conducted by Patrick, Finn, and Stich, and implemented into the numerical code XROTOR. The blade design modifications that were investigated to achieve the proposed goals include: increasing the number of blades, adjusting the chord length, beta distribution, radius of the blade, airfoil shape, and operating RPM. In order to determine the optimal blade design, a baseline case is first developed and the parameters listed earlier are then varied to create a new propeller design that reduces the sound pressure level (SPL) while maintaining performance levels within a predetermined range of the original specifications. From the analysis, the most significant improvements observed in lowering the acoustic signature are dominated by operating rpm and blade radius. A three-, four-, and five-blade configuration was developed that reduced the SPL generated by the propeller during cruise flight conditions. The optimum configuration that produced the greatest SPL reduction was the five-blade configuration. The resulting sound pressure level was reduced from the original 77 dB at 1000' ft above ground level (AGL), to 54 dB at 1000' AGL while remaining within 1.4% of the original thrust and efficiency.

A review of the supply of liquid propellants and other fluids in support of the Space Shuttle Program

NASA Technical Reports Server (NTRS)

Churchwell, Stacy E.; Bain, A. L.

1989-01-01

In this study, over twenty significant liquid propellants and other fluids were reviewed as to their supply in support of the Space Shuttle Program (SSP), primarily at KSC. The uniqueness of most of the products, either by their application or production characteristics, present a variety of supply issues to contend with. Each, however, is critical to the success of the SSP. It becomes necessary to formulate, and maintain, a logistic approach to assure a continued availability of each product. For convenience, two categories were established. One, labeled limited-availability, represents those products wherein they are single sourced, have production restrictions and/or there has been a history of supply problems. The other, labeled universally-available, is characteristic of those having several sources and/or having little, if any, historical supply problems. This last category was not examined in depth. Through concepts of establishing stockpile inventories, multiple supply contracts, or other arrangements, the supply of liquid propellants and other fluids can be assured.

Effect of Body Nose Shape on the Propulsive Efficiency of a Propeller

NASA Technical Reports Server (NTRS)

Stickle, George W; Crigler, John L; Naiman, Irven

1941-01-01

Report presents the results of an investigation of the propulsive efficiency of three adjustable propellers of 10-foot diameter operated in front of four body nose shapes, varying from streamline nose that continued through the propeller plane in the form of a large spinner to a conventional open-nose radial-engine cowling. One propeller had airfoil sections close to the hub, the second had conventional round blade shanks, and the third differed from the second only in pitch distribution. The blade-angle settings ranged from 20 degrees to 55 degrees at the 0.75 radius. The effect of the body nose shape on propulsive efficiency may be divided into two parts: (1) the change in the body drag due to the propeller slipstream and (2) the change in propeller load distribution due to the change in velocity caused by the body. For the nose shape tested in the report, the first effect is shown to be very small; therefore, the chief emphasis of the report is confined to the second effect.

Numerical simulations in the development of propellant management devices

NASA Astrophysics Data System (ADS)

Gaulke, Diana; Winkelmann, Yvonne; Dreyer, Michael

Propellant management devices (PMDs) are used for positioning the propellant at the propel-lant port. It is important to provide propellant without gas bubbles. Gas bubbles can inflict cavitation and may lead to system failures in the worst case. Therefore, the reliable operation of such devices must be guaranteed. Testing these complex systems is a very intricate process. Furthermore, in most cases only tests with downscaled geometries are possible. Numerical sim-ulations are used here as an aid to optimize the tests and to predict certain results. Based on these simulations, parameters can be determined in advance and parts of the equipment can be adjusted in order to minimize the number of experiments. In return, the simulations are validated regarding the test results. Furthermore, if the accuracy of the numerical prediction is verified, then numerical simulations can be used for validating the scaling of the experiments. This presentation demonstrates some selected numerical simulations for the development of PMDs at ZARM.

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.

Prediction of added noise due to the effect of unsteady flow on pusher propellers

NASA Technical Reports Server (NTRS)

Takallu, M. A.; Block, P. J. W.

1987-01-01

An analytical/computational study has been conducted to predict the effect of an upstream wing or pylon on the noise of an operating propeller. The wing trailing edge was placed at variable distances (0.1 and 0.3 chord) upstream of a scaled model propeller (SR-2). The wake was modeled using a similarity formulation. The instantaneous pressure distribution on the propeller blades during the passage through the wake was formulated in terms of a time-dependent variation of each blade section's angle of attack and in terms of the shed vortices from the blade trailing edge. It was found that the final expressions for the unsteady loads considerably altered the radiated noise pattern. Predicted noise for various observer positions, rotational speeds and propeller/pylon distances were computed and are presented in terms of the pressure time history. It has been shown that the positioning of a pylon upstream of a propeller indeed increases the noise. Some comparisons with experimental results are also given.

Performance characteristics of ring-cusp thrusters with xenon propellant

NASA Technical Reports Server (NTRS)

Patterson, M. J.

1986-01-01

The performance characteristics and operating envelope of several 30-cm ring-cusp ion thrusters with xenon propellant were investigated. Results indicate a strong performance dependence on the discharge chamber boundary magnetic fields and resultant distribution of electron currents. Significant improvements in discharge performance over J-series divergent-field thrusters were achieved for large throttling ranges, which translate into reduced cathode emission currents and reduced power dissipation which should be of significant benefit for operation at thruster power levels in excess of 10 kW. Mass spectrometry of the ion beam was documented for both the ring-cusp and J-series thrusters with xenon propellant for determination of overall thruster efficiency, and lifetime. Based on the lower centerline values of doubly charged ions in the ion beam and the lower operating discharge voltage, the screen grid erosion rate of the ring-cusp thruster is expected to be lower than the divergent-field J-series thruster by a factor of 2.

Performance characteristics of ring-cusp thrusters with xenon propellant

NASA Technical Reports Server (NTRS)

Patterson, M. J.

1986-01-01

The performance characteristics and operating envelope of several 30-cm ring-cusp ion thrusters with xenon propellant were investigated. Results indicate a strong performance dependence on the discharge chamber boundary magnetic fields and resultant distribution of electron currents. Significant improvements in discharge performance over J-series divergent-field thrusters were achieved for large throttling ranges, which translate into reduced cathode emission currents and reduced power dissipation which should be of significant benefit for operation at thruster power levels in excess of 10 kW. Mass spectrometer of the ion beam was documented for both the ring-cusp and J-series thrusters with xenon propellant for determination of overall thruster efficiency, and lifetime. Based on the lower centerline values of doubly charged ions in the ion beam and the lower operating discharge voltage, the screen grid erosion rate of the ring-cusp thruster is expected to be lower than the divergent-field J-series thruster by a factor of 2.

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.;

High variable mixture ratio oxygen/hydrogen engine

NASA Technical Reports Server (NTRS)

Erickson, C. M.; Tu, W. H.; Weiss, A. H.

1988-01-01

The ability of an O2/H2 engine to operate over a range of high-propellant mixture ratios was previously shown to be advantageous in single stage to orbit (SSTO) vehicles. The results are presented for the analysis of high-performance engine power cycles operating over propellant mixture ratio ranges of 12 to 6 and 9 to 6. A requirement to throttle up to 60 percent of nominal thrust was superimposed as a typical throttle range to limit vehicle acceleration as propellant is expended. The object of the analysis was to determine areas of concern relative to component and engine operability or potential hazards resulting from the operating requirements and ranges of conditions that derive from the overall engine requirements. The SSTO mission necessitates a high-performance, lightweight engine. Therefore, staged combustion power cycles employing either dual fuel-rich preburners or dual mixed (fuel-rich and oxygen-rich) preburners were examined. Engine mass flow and power balances were made and major component operating ranges were defined. Component size and arrangement were determined through engine layouts for one of the configurations evaluated. Each component is being examined to determine if there are areas of concern with respect to component efficiency, operability, reliability, or hazard. The effects of reducing the maximum chamber pressure were investigated for one of the cycles.

30 CFR 57.9316 - Notifying the equipment operator.

Code of Federal Regulations, 2010 CFR

2010-07-01

..., Hauling, and Dumping Safety Devices, Provisions, and Procedures for Roadways, Railroads, and Loading and Dumping Sites § 57.9316 Notifying the equipment operator. When an operator of self-propelled mobile...

30 CFR 56.9316 - Notifying the equipment operator.

Code of Federal Regulations, 2010 CFR

2010-07-01

..., Hauling, and Dumping Safety Devices, Provisions, and Procedures for Roadways, Railroads, and Loading and Dumping Sites § 56.9316 Notifying the equipment operator. When an operator of self-propelled mobile...

High performance alloy electroforming

NASA Technical Reports Server (NTRS)

Malone, G. A.; Winkelman, D. M.

1989-01-01

Electroformed copper and nickel are used in structural applications for advanced propellant combustion chambers. An improved process has been developed by Bell Aerospace Textron, Inc. wherein electroformed nickel-manganese alloy has demonstrated superior mechanical and thermal stability when compared to previously reported deposits from known nickel plating processes. Solution chemistry and parametric operating procedures are now established and material property data is established for deposition of thick, large complex shapes such as the Space Shuttle Main Engine. The critical operating variables are those governing the ratio of codeposited nickel and manganese. The deposition uniformity which in turn affects the manganese concentration distribution is affected by solution resistance and geometric effects as well as solution agitation. The manganese concentration in the deposit must be between 2000 and 3000 ppm for optimum physical properties to be realized. The study also includes data regarding deposition procedures for achieving excellent bond strength at an interface with copper, nickel-manganese or INCONEL 718. Applications for this electroformed material include fabrication of complex or re-entry shapes which would be difficult or impossible to form from high strength alloys such as INCONEL 718.

Experimental and Numerical Investigation of Reduced Gravity Fluid Slosh Dynamics for the Characterization of Cryogenic Launch and Space Vehicle Propellants

NASA Technical Reports Server (NTRS)

Walls, Laurie K.; Kirk, Daniel; deLuis, Kavier; Haberbusch, Mark S.

2011-01-01

As space programs increasingly investigate various options for long duration space missions the accurate prediction of propellant behavior over long periods of time in microgravity environment has become increasingly imperative. This has driven the development of a detailed, physics-based understanding of slosh behavior of cryogenic propellants over a range of conditions and environments that are relevant for rocket and space storage applications. Recent advancements in computational fluid dynamics (CFD) models and hardware capabilities have enabled the modeling of complex fluid behavior in microgravity environment. Historically, launch vehicles with moderate duration upper stage coast periods have contained very limited instrumentation to quantify propellant stratification and boil-off in these environments, thus the ability to benchmark these complex computational models is of great consequence. To benchmark enhanced CFD models, recent work focuses on establishing an extensive experimental database of liquid slosh under a wide range of relevant conditions. In addition, a mass gauging system specifically designed to provide high fidelity measurements for both liquid stratification and liquid/ullage position in a micro-gravity environment has been developed. This pUblication will summarize the various experimental programs established to produce this comprehensive database and unique flight measurement techniques.

Establishing Decision Trees for Predicting Successful Postpyloric Nasoenteric Tube Placement in Critically Ill Patients.

PubMed

Chen, Weisheng; Sun, Cheng; Wei, Ru; Zhang, Yanlin; Ye, Heng; Chi, Ruibin; Zhang, Yichen; Hu, Bei; Lv, Bo; Chen, Lifang; Zhang, Xiunong; Lan, Huilan; Chen, Chunbo

2016-08-31

Despite the use of prokinetic agents, the overall success rate for postpyloric placement via a self-propelled spiral nasoenteric tube is quite low. This retrospective study was conducted in the intensive care units of 11 university hospitals from 2006 to 2016 among adult patients who underwent self-propelled spiral nasoenteric tube insertion. Success was defined as postpyloric nasoenteric tube placement confirmed by abdominal x-ray scan 24 hours after tube insertion. Chi-square automatic interaction detection (CHAID), simple classification and regression trees (SimpleCart), and J48 methodologies were used to develop decision tree models, and multiple logistic regression (LR) methodology was used to develop an LR model for predicting successful postpyloric nasoenteric tube placement. The area under the receiver operating characteristic curve (AUC) was used to evaluate the performance of these models. Successful postpyloric nasoenteric tube placement was confirmed in 427 of 939 patients enrolled. For predicting successful postpyloric nasoenteric tube placement, the performance of the 3 decision trees was similar in terms of the AUCs: 0.715 for the CHAID model, 0.682 for the SimpleCart model, and 0.671 for the J48 model. The AUC of the LR model was 0.729, which outperformed the J48 model. Both the CHAID and LR models achieved an acceptable discrimination for predicting successful postpyloric nasoenteric tube placement and were useful for intensivists in the setting of self-propelled spiral nasoenteric tube insertion. © 2016 American Society for Parenteral and Enteral Nutrition.

Establishing Decision Trees for Predicting Successful Postpyloric Nasoenteric Tube Placement in Critically Ill Patients.

PubMed

Chen, Weisheng; Sun, Cheng; Wei, Ru; Zhang, Yanlin; Ye, Heng; Chi, Ruibin; Zhang, Yichen; Hu, Bei; Lv, Bo; Chen, Lifang; Zhang, Xiunong; Lan, Huilan; Chen, Chunbo

2018-01-01

Despite the use of prokinetic agents, the overall success rate for postpyloric placement via a self-propelled spiral nasoenteric tube is quite low. This retrospective study was conducted in the intensive care units of 11 university hospitals from 2006 to 2016 among adult patients who underwent self-propelled spiral nasoenteric tube insertion. Success was defined as postpyloric nasoenteric tube placement confirmed by abdominal x-ray scan 24 hours after tube insertion. Chi-square automatic interaction detection (CHAID), simple classification and regression trees (SimpleCart), and J48 methodologies were used to develop decision tree models, and multiple logistic regression (LR) methodology was used to develop an LR model for predicting successful postpyloric nasoenteric tube placement. The area under the receiver operating characteristic curve (AUC) was used to evaluate the performance of these models. Successful postpyloric nasoenteric tube placement was confirmed in 427 of 939 patients enrolled. For predicting successful postpyloric nasoenteric tube placement, the performance of the 3 decision trees was similar in terms of the AUCs: 0.715 for the CHAID model, 0.682 for the SimpleCart model, and 0.671 for the J48 model. The AUC of the LR model was 0.729, which outperformed the J48 model. Both the CHAID and LR models achieved an acceptable discrimination for predicting successful postpyloric nasoenteric tube placement and were useful for intensivists in the setting of self-propelled spiral nasoenteric tube insertion. © 2016 American Society for Parenteral and Enteral Nutrition.

76 FR 9495 - Feathering Propeller Systems for Light-Sport Aircraft Powered Gliders

Federal Register 2010, 2011, 2012, 2013, 2014

2011-02-18

.... No. 1-66] RIN 2120-AJ81 Feathering Propeller Systems for Light-Sport Aircraft Powered Gliders AGENCY... rule amends the definition of light-sport aircraft by removing ``auto'' from the term ``autofeathering... operation for powered gliders that qualify as light-sport aircraft. DATES: The effective date for the final...

Evaluating shrinkage of wood propellers in a high-temperature environment

Treesearch

Richard Bergman; Robert J. Ross

2008-01-01

Minimizing wood shrinkage is a priority for many wood products in use, particularly engineered products manufactured to close tolerances, such as wood propellers for unmanned surveillance aircraft used in military operations. Those currently in service in the Middle East are experiencing performance problems as a consequence of wood shrinking during long-term storage...

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.

PHM Enabled Autonomous Propellant Loading Operations

NASA Technical Reports Server (NTRS)

Walker, Mark; Figueroa, Fernando

2017-01-01

The utility of Prognostics and Health Management (PHM) software capability applied to Autonomous Operations (AO) remains an active research area within aerospace applications. The ability to gain insight into which assets and subsystems are functioning properly, along with the derivation of confident predictions concerning future ability, reliability, and availability, are important enablers for making sound mission planning decisions. When coupled with software that fully supports mission planning and execution, an integrated solution can be developed that leverages state assessment and estimation for the purposes of delivering autonomous operations. The authors have been applying this integrated, model-based approach to the autonomous loading of cryogenic spacecraft propellants at Kennedy Space Center.

Performance of 10-kW class xenon ion thrusters

NASA Technical Reports Server (NTRS)

Patterson, Michael J.; Rawlin, Vincent K.

1988-01-01

Presented are performance data for laboratory and engineering model 30 cm-diameter ion thrusters operated with xenon propellant over a range of input power levels from approximately 2 to 20 kW. Also presented are preliminary performance results obtained from laboratory model 50 cm-diameter cusp- and divergent-field ion thrusters operating with both 30 cm- amd 50 cm-diameter ion optics up to a 20 kW input power. These data include values of discharge chamber propellant and power efficiencies, as well as values of specific impulse, thruster efficiency, thrust and power. The operation of the 30 cm- and 50 cm-diameter ion optics are also discussed.

50 KW Class Krypton Hall Thruster Performance

NASA Technical Reports Server (NTRS)

Jacobson, David T.; Manzella, David H.

2003-01-01

The performance of a 50-kilowatt-class Hall thruster designed for operation on xenon propellant was measured using kryton propellant. The thruster was operated at discharge power levels ranging from 6.4 to 72.5 kilowatts. The device produced thrust ranging from 0.3 to 2.5 newtons. The thruster was operated at discharge voltages between 250 and 1000 volts. At the highest anode mass flow rate and discharge voltage and assuming a 100 percent singly charged condition, the discharge specific impulse approached the theoretical value. Discharge specific impulse of 4500 seconds was demonstrated at a discharge voltage of 1000 volts. The peak discharge efficiency was 64 percent at 650 volts.

NASA's Evolutionary Xenon Thruster (NEXT) Component Verification Testing

NASA Technical Reports Server (NTRS)

Herman, Daniel A.; Pinero, Luis R.; Sovey, James S.

2009-01-01

Component testing is a critical facet of the comprehensive thruster life validation strategy devised by the NASA s Evolutionary Xenon Thruster (NEXT) program. Component testing to-date has consisted of long-duration high voltage propellant isolator and high-cycle heater life validation testing. The high voltage propellant isolator, a heritage design, will be operated under different environmental condition in the NEXT ion thruster requiring verification testing. The life test of two NEXT isolators was initiated with comparable voltage and pressure conditions with a higher temperature than measured for the NEXT prototype-model thruster. To date the NEXT isolators have accumulated 18,300 h of operation. Measurements indicate a negligible increase in leakage current over the testing duration to date. NEXT 1/2 in. heaters, whose manufacturing and control processes have heritage, were selected for verification testing based upon the change in physical dimensions resulting in a higher operating voltage as well as potential differences in thermal environment. The heater fabrication processes, developed for the International Space Station (ISS) plasma contactor hollow cathode assembly, were utilized with modification of heater dimensions to accommodate a larger cathode. Cyclic testing of five 1/22 in. diameter heaters was initiated to validate these modified fabrication processes while retaining high reliability heaters. To date two of the heaters have been cycled to 10,000 cycles and suspended to preserve hardware. Three of the heaters have been cycled to failure giving a B10 life of 12,615 cycles, approximately 6,000 more cycles than the established qualification B10 life of the ISS plasma contactor heaters.

Moon-Based Advanced Reusable Transportation Architecture: The MARTA Project

NASA Astrophysics Data System (ADS)

Alexander, R.; Bechtel, R.; Chen, T.; Cormier, T.; Kalaver, S.; Kirtas, M.; Lewe, J.-H.; Marcus, L.; Marshall, D.; Medlin, M.; McIntire, J.; Nelson, D.; Remolina, D.; Scott, A.; Weglian, J.; Olds, J.

2000-01-01

The Moon-based Advanced Reusable Transportation Architecture (MARTA) Project conducted an in-depth investigation of possible Low Earth Orbit (LEO) to lunar surface transportation systems capable of sending both astronauts and large masses of cargo to the Moon and back. This investigation was conducted from the perspective of a private company operating the transportation system for a profit. The goal of this company was to provide an Internal Rate of Return (IRR) of 25% to its shareholders. The technical aspect of the study began with a wide open design space that included nuclear rockets and tether systems as possible propulsion systems. Based on technical, political, and business considerations, the architecture was quickly narrowed down to a traditional chemical rocket using liquid oxygen and liquid hydrogen. However, three additional technologies were identified for further investigation: aerobraking, in-situ resource utilization (ISRU), and a mass driver on the lunar surface. These three technologies were identified because they reduce the mass of propellant used. Operational costs are the largest expense with propellant cost the largest contributor. ISRU, the production of materials using resources on the Moon, was considered because an Earth to Orbit (ETO) launch cost of 1600 per kilogram made taking propellant from the Earth's surface an expensive proposition. The use of an aerobrake to circularize the orbit of a vehicle coming from the Moon towards Earth eliminated 3, 100 meters per second of velocity change (Delta V), eliminating almost 30% of the 11,200 m/s required for one complete round trip. The use of a mass driver on the lunar surface, in conjunction with an ISRU production facility, would reduce the amount of propellant required by eliminating using propellant to take additional propellant from the lunar surface to Low Lunar Orbit (LLO). However, developing and operating such a system required further study to identify if it was cost effective. The vehicle was modeled using the Simulated Probabilistic Parametric Lunar Architecture Tool (SPPLAT), which incorporated the disciplines of Weights and Sizing, Trajectories, and Cost. This tool used ISRU propellant cost, Technology Reduction Factor (a dry weight reduction due to improved technology), and vehicle engine specific impulse as inputs. Outputs were vehicle dry weight, total propellant used per trip, and cost to charge the customer in order to guarantee an IRR of 25%. SPPLAT also incorporated cost estimation error, weight estimation error, market growth, and ETO launch cost as uncertainty variables. Based on the stipulation that the venture be profitable, the price to charge the customer was highly dependent on ISRU propellant cost and relatively insensitive to the other inputs. The best estimate of ISRU cost is 1000/kg, and results in a price to charge the customer of 2600/kg of payload. If ISRU cost can be reduced to 160/kg, the price to the customer is reduced to just 800/kg of payload. Additionally, the mass driver was only cost effective at an ISRU propellant cost greater than 250/kg, although it reduced total propellant used by 35%. In conclusion, this mission is achievable with current technology, but is only profitable with greater research into the enabling technology of ISRU propellant production.

Fiber-optic sensing in cryogenic environments. [for rocket propellant tank monitoring

NASA Technical Reports Server (NTRS)

Sharma, M.; Brooks, R. E.

1980-01-01

Passive optical sensors using fiber-optic signal transmission to a remote monitoring station are explored as an alternative to electrical sensors used to monitor the status of explosive propellants. The designs of passive optical sensors measuring liquid level, pressure, and temperature in cryogenic propellant tanks are discussed. Test results for an experimental system incorporating these sensors and operating in liquid nitrogen demonstrate the feasibility of passive sensor techniques and indicate that they can serve as non-hazardous replacements for more conventional measuring equipment in explosive environments.

A Preliminary Study of a Propeller Powered by Gas Jets Issuing from the Blade Tips

DTIC Science & Technology

1946-11-01

ISSUING FROM THE BLADE TIPS By J. C. Sanders and N. D. Sanders Aircraft Engine Research Laboratory Cleveland, Ohio icaflit w<• w &£N •^5$" jm "^o*6w...propeller powered by Jets in the blade tips made by Roy in 1930 (reference 3) showed that this engine would be less efficient than;a reciprocating...development of the turbojet engine , which is .now of outstanding interest. The possibilities of the jet -operated propeller are re-exeroined and the

Solid-propellant motors for high-incremental-velocity low-acceleration maneuvers in space

NASA Technical Reports Server (NTRS)

Shafer, J. I.

1972-01-01

The applicability of solid-propellant rockets into a regime of high-performance long-burning tasks beyond the capability of existing motors is discussed. Successful static test firings have demonstrated the feasibility of: (1) utilizing fully case-bonded end-burning propellant charges without mechanical stress relief; (2) using an all-carbon radiative nozzle markedly lighter than the flight-weight ablative nozzle it replaces, and (3) producing low spacecraft acceleration rates during the thrust transient through a controlled-flow igniter that promotes operation below the previous combustion limit.

Multiple resonant railgun power supply

DOEpatents

Honig, E.M.; Nunnally, W.C.

1985-06-19

A multiple repetitive resonant railgun power supply provides energy for repetitively propelling projectiles from a pair of parallel rails. A plurality of serially connected paired parallel rails are powered by similar power supplies. Each supply comprises an energy storage capacitor, a storage inductor to form a resonant circuit with the energy storage capacitor and a magnetic switch to transfer energy between the resonant circuit and the pair of parallel rails for the propelling of projectiles. The multiple serial operation permits relatively small energy components to deliver overall relatively large amounts of energy to the projectiles being propelled.

Multiple resonant railgun power supply

DOEpatents

Honig, Emanuel M.; Nunnally, William C.

1988-01-01

A multiple repetitive resonant railgun power supply provides energy for repetitively propelling projectiles from a pair of parallel rails. A plurality of serially connected paired parallel rails are powered by similar power supplies. Each supply comprises an energy storage capacitor, a storage inductor to form a resonant circuit with the energy storage capacitor and a magnetic switch to transfer energy between the resonant circuit and the pair of parallel rails for the propelling of projectiles. The multiple serial operation permits relatively small energy components to deliver overall relatively large amounts of energy to the projectiles being propelled.

Aerodynamic Characteristics at High Speeds of Full-Scale Propellers having Different Shank Designs

NASA Technical Reports Server (NTRS)

Maynard, Julian D.

1947-01-01

Tests of two 10-foot-diameter two-blade propellers which differed only in shank design have been made in the Langley 16-foot high-speed tunnel. The propellers are designated by their blade design numbers, NACA 10-(5)(08)-03, which had aerodynamically efficient airfoil shank sections, and NACA l0-(5)(08)-03R which had thick cylindrical shank sections typical of conventiona1 blades, The propellers mere tested on a 2000-horsepower dynamometer through a range of blade-angles from 20deg to 55deg at various rotational speeds and at airspeeds up to 496 miles per hour. The resultant tip speeds obtained simulate actual flight conditions, and the variation of air-stream Mach number with advance ratio is within the range of full-scale constant-speed propeller operation. Both propellers were very efficient, the maximum envelope efficiency being approximately 0,95 for the NACA 10-(5)(08)-03 propeller and about 5 percent less for the NACA 10-(5)(08)-03R propeller. Based on constant power and rotational speed, the efficiency of the NACA 10-(05)(08)-03 propeller was from 2.8 to 12 percent higher than that of the NACA 10-(5)(08)-03R propeller over a range of airspeeds from 225 to 450 miles per hour. The loss in maximum efficiency at the design blade angle for the NACA 10-(5)(08)-03 and 10-(5)(08)-03R propellers vas about 22 and 25 percent, respectively, for an increase in helical tip Mach number from 0.70 to 1.14.

Experimental Investigation of a Direct-drive Hall Thruster and Solar Array System at Power Levels up to 10 kW

NASA Technical Reports Server (NTRS)

Snyder, John S.; Brophy, John R.; Hofer, Richard R.; Goebel, Dan M.; Katz, Ira

2012-01-01

As NASA considers future exploration missions, high-power solar-electric propulsion (SEP) plays a prominent role in achieving many mission goals. Studies of high-power SEP systems (i.e. tens to hundreds of kilowatts) suggest that significant mass savings may be realized by implementing a direct-drive power system, so NASA recently established the National Direct-Drive Testbed to examine technical issues identified by previous investigations. The testbed includes a 12-kW solar array and power control station designed to power single and multiple Hall thrusters over a wide range of voltages and currents. In this paper, single Hall thruster operation directly from solar array output at discharge voltages of 200 to 450 V and discharge powers of 1 to 10 kW is reported. Hall thruster control and operation is shown to be simple and no different than for operation on conventional power supplies. Thruster and power system electrical oscillations were investigated over a large range of operating conditions and with different filter capacitances. Thruster oscillations were the same as for conventional power supplies, did not adversely affect solar array operation, and were independent of filter capacitance from 8 to 80 ?F. Solar array current and voltage oscillations were very small compared to their mean values and showed a modest dependence on capacitor size. No instabilities or anomalous behavior were observed in the thruster or power system at any operating condition investigated, including near and at the array peak power point. Thruster startup using the anode propellant flow as the power 'switch' was shown to be simple and reliable with system transients mitigated by the proper selection of filter capacitance size. Shutdown via cutoff of propellant flow was also demonstrated. A simple electrical circuit model was developed and is shown to have good agreement with the experimental data.

Counter-rotating propeller noise directivity and trends

NASA Technical Reports Server (NTRS)

Block, P. J. W.; Klatte, R. J.; Druez, P. M.

1986-01-01

The effects of power loading on the far field noise spectra and directivity of counter-rotating propellers (CRP) were studied using a model scale SR-2 propeller in a low-speed anechoic wind tunnel. Approximately 264 far field noise measurements were obtained for each CRP configuration (pusher and tractor) and operating conditions covering from 30 to 140 deg to the flight direction and up to 340 deg circumferentially. Data indicated that the CRP tractor produced higher levels in the second and third harmonics which propagated axially; in effect, the noise exposure time increased over that of a single single-rotation propeller. The effects of pylon-to-propeller spacing, type of pylon attachment and reduced rear-blade row radius are considered and it is found that the 0.3 chord radial pylon produces less additional noise than the 0.1 chord radial pylon and that the 0.2 chord tangential pylon is the quietest pusher configuration.

Supersonic propeller noise in a uniform flow

NASA Technical Reports Server (NTRS)

Jou, Wen-Huei

1989-01-01

The sound field produced by a supersonic propeller operating in a uniform flow is investigated. The main interest is the effect of the finite forward flight speed on the directivity of the sound field as seen by an observer on the aircraft. It is found that there are cones of silence on the axis of the propeller. The semiapex angles on these cones are equal fore and aft of the propeller plane, and depend on the tip Mach number only. The Fourier coefficients of the acoustic pressure contain the Doppler amplification factor. The sound field weakens in the upstream direction and strengthen downstream. Kinematic considerations of the emitted Mach waves not only confirm these results, but also provide physical insight into the sound generation mechanism. The predicted zone of silence and the Doppler amplification factor are compared to the theoretical prediction of shock wave formation and the flight test of the SR3 propeller.

Drag and Propulsive Characteristics of Air-Cooled Engine-Nacelle Installations for Large Airplane

NASA Technical Reports Server (NTRS)

Silverstein, Abe; Wilson, Herbert A , Jr

1942-01-01

An investigation was conducted in the NACA full-scale wind tunnel to determine the drag and the propulsive efficiency of nacelle-propeller arrangements for a large range of nacelle sizes. In contrast with usual tests with a single nacelle, these tests were conducted with nacelle-propeller installations on a large model of a four-engine airplane. Data are presented on the first part of the investigation, covering seven nacelle arrangements with nacelle diameters from 0.53 to 1.5 times the wing thickness. These ratios are similar to those occurring on airplanes weighing from about 20 to 100 tons. The results show the drag, the propulsive efficiency, and the over-all efficiency of the various nacelle arrangements as functions of the nacelle size, the propeller position, and the airplane lift coefficient. The effect of the nacelles on the aerodynamic characteristics of the model is shown for both propeller-removed and propeller-operating conditions.

Simulation Analysis of Computer-Controlled pressurization for Mixture Ratio Control

NASA Technical Reports Server (NTRS)

Alexander, Leslie A.; Bishop-Behel, Karen; Benfield, Michael P. J.; Kelley, Anthony; Woodcock, Gordon R.

2005-01-01

A procedural code (C++) simulation was developed to investigate potentials for mixture ratio control of pressure-fed spacecraft rocket propulsion systems by measuring propellant flows, tank liquid quantities, or both, and using feedback from these measurements to adjust propellant tank pressures to set the correct operating mixture ratio for minimum propellant residuals. The pressurization system eliminated mechanical regulators in favor of a computer-controlled, servo- driven throttling valve. We found that a quasi-steady state simulation (pressure and flow transients in the pressurization systems resulting from changes in flow control valve position are ignored) is adequate for this purpose. Monte-Carlo methods are used to obtain simulated statistics on propellant depletion. Mixture ratio control algorithms based on proportional-integral-differential (PID) controller methods were developed. These algorithms actually set target tank pressures; the tank pressures are controlled by another PID controller. Simulation indicates this approach can provide reductions in residual propellants.

Simulated propeller slipstream effects on a supercritical wing

NASA Technical Reports Server (NTRS)

Welge, H. R.; Crowder, J. P.

1978-01-01

To quantify the installed performance of high speed (M = 0.8) turboprop propulsion systems, an experimental program designed to assess the magnitude of the aerodynamic interference of a propeller slipstream on a supercritical wing has been conducted. The test was conducted in the NASA Ames 14-foot wind tunnel. An ejector-nacelle propeller slipstream simulator was used to produce a slipstream with characteristics typical of advanced propellers presently being investigated. A supercritical wing-body configuration was used to evaluate the interference effects. A traversing total pressure rake was used to make flow field measurements behind the wing and to calibrate the slipstream simulator. The force results indicated that the interference drag amounted to an increase of ten counts or about 3% of the wing-body drag for a two engine configuration at the nominal propeller operating conditions. However, at the higher swirl angles (11 deg vs. 7 deg nominally) the interference drag was favorable by about the same magnitude.

Study of monopropellants for electrothermal thrusters

NASA Technical Reports Server (NTRS)

Kuenzly, J. D.

1974-01-01

A 333 mN electrothermal thruster designed to use MIL-grade hydrazine was demonstrated to be suitable for operation with low freezing point monopropellants containing hydrazine azide, monomethylhydrazine, unsymmetrical-dimethylhydrazine and ammonia. The steady-state specific impulse was greater than 200 sec for all propellants. The pulsed-mode specific impulse for an azide blend exceeded 175 sec for pulse widths greater than 50 msec; propellants containing carbonaceous species delivered 175 sec pulsed-mode specific impulses for pulse widths greater than 100 msec. Longer thrust chamber residence times were required for the carbonaceous propellants; the original thruster design was modified by increasing the characteristic chamber length and screen packing density. Specific recommendations were made for the work required to design and develop flight worthy thrusters, including methods to increase propellant dispersal at injection, thruster geometry changes to reduce holding power levels and methods to initiate the rapid decomposition of the carbonaceous propellants.

Sea trials of a ducted tip propeller designed for improved cavitation performance

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

Hordnes, I.; Bidaud, A.; Green, S.I.

1994-12-31

Studies have shown that ``ring-wing`` or ``ducted`` tip devices reduce substantially the inception index of trailing vortices generated by a hydrofoil (Green et al. 1988). It has also been shown that these devices improve the lift/drag ratio of an airfoil at high angle of incidence (Duan et al. 1992). These finding indicate that there may be a marine application for the ducted tip. Experimental equipment has been designed and manufactured in preparation for upcoming tests of a propeller with ducted tips. The tips are tubes aligned with the propeller blade tips that will replace a radial fraction of the originalmore » blade tips equal to the diameter of the tubes. The tube dimensions have been chosen according to the span/tip diameter and chord/tip length ratios used by Duan et al. (1992), and the tubes will be given a curvature equal to the propeller tip radius. Field trials will be given a curvature equal to the propeller tip radius. Field trials will be conducted on a 36 inch diameter propeller that is used to propel a 45 ft. fishing (seine) boat operating in the coastal waters outside Vancouver. The performance of the propeller will be measured in terms of the propeller efficiency as a function of advance ratio. A special force transducer has been designed that is capable of recording both torque and thrust on the propeller shaft even though these are expected to produce shaft strains of different orders of magnitude. As a supplementary means of monitoring the propeller performance, a hydrophone will be located near the propeller wake in order to measure the tip vortex cavitation noise.« less

Thermo-physical performance prediction of the KSC Ground Operation Demonstration Unit for liquid hydrogen

NASA Astrophysics Data System (ADS)

Baik, J. H.; Notardonato, W. U.; Karng, S. W.; Oh, I.

2015-12-01

NASA Kennedy Space Center (KSC) researchers have been working on enhanced and modernized cryogenic liquid propellant handling techniques to reduce life cycle costs of propellant management system for the unique KSC application. The KSC Ground Operation Demonstration Unit (GODU) for liquid hydrogen (LH2) plans to demonstrate integrated refrigeration, zero-loss flexible term storage of LH2, and densified hydrogen handling techniques. The Florida Solar Energy Center (FSEC) has partnered with the KSC researchers to develop thermal performance prediction model of the GODU for LH2. The model includes integrated refrigeration cooling performance, thermal losses in the tank and distribution lines, transient system characteristics during chilling and loading, and long term steady-state propellant storage. This paper will discuss recent experimental data of the GODU for LH2 system and modeling results.

Fractional Consumption of Liquid Hydrogen and Liquid Oxygen During the Space Shuttle Program

NASA Technical Reports Server (NTRS)

Partridge, Jonathan K.

2011-01-01

The Space Shuttle uses the propellants, liquid hydrogen and liquid oxygen, to meet part of the propulsion requirements from ground to orbit. The Kennedy Space Center procured over 25 million kilograms of liquid hydrogen and over 250 million kilograms of liquid oxygen during the 3D-year Space Shuttle Program. Because of the cryogenic nature of the propellants, approximately 55% of the total purchased liquid hydrogen and 30% of the total purchased liquid oxygen were used in the Space Shuttle Main Engines. The balance of the propellants were vaporized during operations for various purposes. This paper dissects the total consumption of liqUid hydrogen and liqUid oxygen and determines the fraction attributable to each of the various processing and launch operations that occurred during the entire Space Shuttle Program at the Kennedy Space Center.

Study on Endurance and Performance of Impregnated Ruthenium Catalyst for Thruster System.

PubMed

Kim, Jincheol; Kim, Taegyu

2018-02-01

Performance and endurance of the Ru catalyst were studied for nitrous oxide monopropellant thruster system. The thermal decomposition of N2O requires a considerably high temperature, which make it difficult to be utilized as a thruster propellant, while the propellant decomposition temperature can be reduced by using the catalyst through the decomposition reaction with the propellant. However, the catalyst used for the thruster was frequently exposed to high temperature and high-pressure environment. Therefore, the state change of the catalyst according to the thruster operation was analyzed. Characterization of catalyst used in the operation condition of the thruster was performed using FE-SEM and EDS. As a result, performance degradation was occurred due to the volatilization of Ru catalyst and reduction of the specific surface area according to the phase change of Al2O3.

Technology Development for Hydrogen Propellant Storage and Transfer at the Kennedy Space Center (KSC)

NASA Technical Reports Server (NTRS)

Youngquist, Robert; Starr, Stanley; Krenn, Angela; Captain, Janine; Williams, Martha

2016-01-01

The National Aeronautics and Space Administration (NASA) is a major user of liquid hydrogen. In particular, NASA's John F. Kennedy (KSC) Space Center has operated facilities for handling and storing very large quantities of liquid hydrogen (LH2) since the early 1960s. Safe operations pose unique challenges and as a result NASA has invested in technology development to improve operational efficiency and safety. This paper reviews recent innovations including methods of leak and fire detection and aspects of large storage tank health and integrity. We also discuss the use of liquid hydrogen in space and issues we are addressing to ensure safe and efficient operations should hydrogen be used as a propellant derived from in-situ volatiles.

Velocity Fluctuations in Helical Propulsion: How Small Can a Propeller Be.

PubMed

Ghosh, Arijit; Paria, Debadrita; Rangarajan, Govindan; Ghosh, Ambarish

2014-01-02

Helical propulsion is at the heart of locomotion strategies utilized by various natural and artificial swimmers. We used experimental observations and a numerical model to study the various fluctuation mechanisms that determine the performance of an externally driven helical propeller as the size of the helix is reduced. From causality analysis, an overwhelming effect of orientational noise at low length scales is observed, which strongly affects the average velocity and direction of motion of a propeller. For length scales smaller than a few micrometers in aqueous media, the operational frequency for the propulsion system would have to increase as the inverse cube of the size, which can be the limiting factor for a helical propeller to achieve locomotion in the desired direction.

Propellers in yaw

NASA Technical Reports Server (NTRS)

Ribner, Herbert S

1945-01-01

It was realized as early as 1909 that a propeller in yaw develops a side force like that of a fin. In 1917, R. G. Harris expressed this force in terms of the torque coefficient for the unyawed propeller. Of several attempts to express the side force directly in terms of the shape of the blades, however, none has been completely satisfactory. An analysis that incorporates induction effects not adequately covered in previous work and that gives good agreement with experiment over a wide range of operating conditions is presented. The present analysis shows that the fin analogy may be extended to the form of the side-force expression and that the effective fin area may be taken as the projected side area of the propeller.

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

Analysis of Experimental Sea-level Transient Data and Analog Method of Obtaining Altitude Response for Turbine-propeller Engine with Relay-type Speed Control

NASA Technical Reports Server (NTRS)

Vasu, George; Pack, George J

1951-01-01

Correlation has been established between transient engine and control data obtained experimentally and data obtained by simulating the engine and control with an analog computer. This correlation was established at sea-level conditions for a turbine-propeller engine with a relay-type speed control. The behavior of the controlled engine at altitudes of 20,000 and 35,000 feet was determined with an analog computer using the altitude pressure and temperature generalization factors to calculate the new engine constants for these altitudes. Because the engine response varies considerably at altitude some type of compensation appears desirable and four methods of compensation are discussed.

Effect of a rotating propeller on the separation angle of attack and distortion in ducted propeller inlets

NASA Technical Reports Server (NTRS)

Boldman, D. R.; Iek, C.; Hwang, D. P.; Larkin, M.; Schweiger, P.

1993-01-01

The present study represents an extension of an earlier wind tunnel experiment performed with the P&W 17-in. Advanced Ducted Propeller (ADP) Simulator operating at Mach 0.2. In order to study the effects of a rotating propeller on the inlet flow, data were obtained in the UTRC 10- by 15-Foot Large Subsonic Wind Tunnel with the same hardware and instrumentation, but with the propeller removed. These new tests were performed over a range of flow rates which duplicated flow rates in the powered simulator program. The flow through the inlet was provided by a remotely located vacuum source. A comparison of the results of this flow-through study with the previous data from the powered simulator indicated that in the conventional inlet the propeller produced an increase in the separation angle of attack between 4.0 deg at a specific flow of 22.4 lb/sec-sq ft to 2.7 deg at a higher specific flow of 33.8 lb/sec-sq ft. A similar effect on separation angle of attack was obtained by using stationary blockage rather than a propeller.

The acoustics of ducted propellers

NASA Astrophysics Data System (ADS)

Ali, Sherif F.

The return of the propeller to the long haul commercial service may be rapidly approaching in the form of advanced "prop fans". It is believed that the advanced turboprop will considerably reduce the operational cost. However, such aircraft will come into general use only if their noise levels meet the standards of community acceptability currently applied to existing aircraft. In this work a time-marching boundary-element technique is developed, and used to study the acoustics of ducted propeller. The numerical technique is developed in this work eliminated the inherent instability suffered by conventional approaches. The methodology is validated against other numerical and analytical results. The results show excellent agreement with the analytical solution and show no indication of unstable behavior. For the ducted propeller problem, the propeller is modeled by a rotating source-sink pairs, and the duct is modeled by rigid annular body of elliptical cross-section. Using the model and the developed technique, the effect of different parameters on the acoustic field is predicted and analyzed. This includes the effect of duct length, propeller axial location, and source Mach number. The results of this study show that installing a short duct around the propeller can reduce the noise that reaches an observer on a side line.

Three-Dimensional Simulation of Base Bleed Unit with AP/HTPB Propellant in Fast Cook-off Conditions

NASA Astrophysics Data System (ADS)

Li, Wen-feng; Yu, Yong-gang; Ye, Rui; Yang, Hou-wen

2017-07-01

In this work, a three-dimensional unsteady heat transfer model of base bleed unit with trilobite ammonium perchlorate (AP)/hydroxyl-terminated polybutadiene (HTPB) composite solid propellant is presented to analyze the cook-off characteristics. According to the two-step chemical reaction of AP/HTPB propellant, a small-scale cook-off test is established. A comparison of the experimental and calculated results is made to verify the rationality of the computation model. On this basis, a cook-off numerical simulation of the base bleed unit at the heating rates of 0.33, 0.58 and 0.83 K/s is presented to investigate the ignition and initiation characteristics. The results show that the ignitions occur on the head face of the AP/HTPB propellant and near the internal gas chamber in these conditions. As the heating rate increases, the runaway time decreases and the ignition temperature rises.

Experimental and Theoretical Study of Propeller Spinner/Shank Interference. M.S. Thesis

NASA Technical Reports Server (NTRS)

Cornell, C. C.

1986-01-01

A fundamental experimental and theoretical investigation into the aerodynamic interference associated with propeller spinner and shank regions was conducted. The research program involved a theoretical assessment of solutions previously proposed, followed by a systematic experimental study to supplement the existing data base. As a result, a refined computational procedure was established for prediction of interference effects in terms of interference drag and resolved into propeller thrust and torque components. These quantities were examined with attention to engineering parameters such as two spinner finess ratios, three blade shank forms, and two/three/four/six/eight blades. Consideration of the physics of the phenomena aided in the logical deduction of two individual interference quantities (cascade effects and spinner/shank juncture interference). These interference effects were semi-empirically modeled using existing theories and placed into a compatible form with an existing propeller performance scheme which provided the basis for examples of application.

Propellant/material compatibility program and results: Ten-year milestones

NASA Technical Reports Server (NTRS)

Moran, C.; Bjorkland, R.

1982-01-01

The analyses and results of a test program to establish the effects of long term (10 years or more) contact of materials with earth-storable propellants for the purpose of designing chemical propulsion system components which are used for current as well as future planetary spacecraft are described. The period from the publication of JPL TM 33-779 IN 1976 through the testing accomplished in 1981 is covered. The following propellants are reported herein: hydrazine, monomethylhydrazine and nitrogen tetroxide. Materials included the following: aluminum alloys, corrosion resistant steels and a titanium alloy. The results of the testing of more than 80 specimens are included. Material ratings relative to the ten year milepost were assigned. Some evidence of propellant decomposition was found. Titanium is rated as acceptable for ten year applications. Aluminum and stainless steel alloys are also rated as acceptable with few restrictions.

Design and Testing of Non-Toxic RCS Thrusters for Second Generation Reusable Launch Vehicle

NASA Technical Reports Server (NTRS)

Calvignac, Jacky; Tramel, Terri

2003-01-01

The current NASA Space Shuttle auxiliary propulsion system utilizes nitrogen tetroxide (NTO) and monomethylhydrazine (MMH), hypergolic propellants. This use of these propellants has resulted in high levels of maintenance and precautions that contribute to costly launch operations. By employing alternate propellant combinations, those less toxic to humans, the hazards and time required between missions can be significantly reduced. Use of alternate propellants can thereby increase the efficiency and lower the cost in launch operations. In support of NASA's Space Launch Initiative (SLI), TRW proposed a three-phase project structured to significantly increase the technology readiness of a high-performance reaction control subsystem (RCS) thruster using non-toxic propellant for an operationally efficient and reusable auxiliary propulsion system (APS). The project enables the development of an integrated primary/vernier thruster capable of providing dual-thrust levels of both 1000-lbf class thrust and 25-lbf thrust. The intent of the project is to reduce the risk associated with the development of an improved RCS flight design that meets the primary NASA objectives of improved safety and reliability while reducing systems operations and maintenance costs. TRW proposed two non-toxic auxiliary propulsion engine designs, one using liquid oxygen and liquid hydrogen and the other using liquid oxygen and liquid ethanol, as candidates to meet the goals of reliability and affordability at the RCS level. Both of these propellant combinations offer the advantage of a safe environment for maintenance, while at the same time providing adequate to excellent performance for a conventional liquid propulsion systems. The key enabling technology incorporated in both TRW thrusters is the coaxial liquid on liquid pintle injector. This paper will concentrate on only the design and testing of one of the thrusters, the liquid oxygen (LOX) and liquid hydrogen (LH2) thruster. The LOX/LH2 thruster design includes a LOX-centered pintle injector, consisting of two rows of slots that create a radial spoke spray pattern in the combustion chamber. The main fuel injector creates a continuous sheet of LH2 originating upstream of the LOX pintle injector. The two propellants impinge at the pintle slots, where the resulting momentum ratio and spray pattern determines the combustion efficiency and thermal effects on the hardware. Another enabling technology used in the design of this thruster is fuel film cooling through a duct, lining the inner wall of the combustion chamber barrel section. The duct is also acts as a secondary fuel injection point. The variation in the amount of LH2 used for the duct allows for adjustments in the cooling capacity for the thruster. The Non-Toxic LOX-LH2 RCS Workhorse Thruster was tested at the NASA Marshall Space Flight Center's Test Stand 500. Hot-fire tests were conducted between March 08, 2002 and April 05, 2002. All testing during the program base period were performed at sea-level conditions. During the test program, 7 configurations were tested, including 2 combustion chambers, 3 LOX injector pintle tips, and 4 LH2 injector stroke settings. The operating conditions that were surveyed varied thrust levels, mixture ratio and LH2 duct cooling flow. The copper heat sink chamber was used for 16 burns, each burn lasting from 0.4 to 10 seconds, totaling 51.4 seconds, followed by Haynes chamber testing ranging from 0.9 to 120 seconds, totaling 300.9 seconds. The total accumulated burn time for the test program is 352.3 seconds. C* efficiency was calculated and found to be within expectable limits for most operating conditions. The temperature on the Haynes combustion chamber remained below established material limits, with the exception of one localized hot spot. The test results demonstrate that both the coaxial liquid-on-liquid pintle injector design and fuel duct concepts are viable for the intended application. The thruster head-e design maintained cryogenic injection temperatures while firing, which validates the concept for minimal heat soak back. By injecting fuel into the duct, the throat temperatures were manageable, yet the split of fuel through the cooling duct does not compromise the overall combustion efficiency, which indicates that, provided proper design refinement, such a concept can be applied to a high-performance version of the thruster. These hot fire tests demonstrate the robustness of the duct design concept and good capability to withstand off-nominal operating conditions without adversely impacting the thermal response of the engine, a key design feature for a cryogenic thruster.

46 CFR 154.1840 - Protective clothing.

Code of Federal Regulations, 2011 CFR

2011-10-01

... FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Operations § 154.1840 Protective clothing... operation, except those assigned to gas-safe cargo control rooms, wears protective clothing. ...

46 CFR 154.1840 - Protective clothing.

Code of Federal Regulations, 2010 CFR

2010-10-01

... FOR SELF-PROPELLED VESSELS CARRYING BULK LIQUEFIED GASES Operations § 154.1840 Protective clothing... operation, except those assigned to gas-safe cargo control rooms, wears protective clothing. ...

Experimental research on air propellers

NASA Technical Reports Server (NTRS)

Durand, William F

1918-01-01

The purposes of the experimental investigation on the performance of air propellers described in this report are as follows: (1) the development of a series of design factors and coefficients drawn from model forms distributed with some regularity over the field of air-propeller design and intended to furnish a basis of check with similar work done in other aerodynamic laboratories, and as a point of departure for the further study of special or individual types and forms; (2) the establishment of a series of experimental values derived from models and intended for later use as a basis for comparison with similar results drawn from certain selected full-sized forms and tested in free flight.

NEXT Propellant Management System Integration With Multiple Ion Thrusters

NASA Technical Reports Server (NTRS)

Sovey, James S.; Soulas, George C.; Herman, Daniel A.

2011-01-01

As a critical part of the NEXT test validation process, a multiple-string integration test was performed on the NEXT propellant management system and ion thrusters. The objectives of this test were to verify that the PMS is capable of providing stable flow control to multiple thrusters operating over the NEXT system throttling range and to demonstrate to potential users that the NEXT PMS is ready for transition to flight. A test plan was developed for the sub-system integration test for verification of PMS and thruster system performance and functionality requirements. Propellant management system calibrations were checked during the single and multi-thruster testing. The low pressure assembly total flow rates to the thruster(s) were within 1.4 percent of the calibrated support equipment flow rates. The inlet pressures to the main, cathode, and neutralizer ports of Thruster PM1R were measured as the PMS operated in 1-thruster, 2-thruster, and 3-thruster configurations. It was found that the inlet pressures to Thruster PM1R for 2-thruster and 3-thruster operation as well as single thruster operation with the PMS compare very favorably indicating that flow rates to Thruster PM1R were similar in all cases. Characterizations of discharge losses, accelerator grid current, and neutralizer performance were performed as more operating thrusters were added to the PMS. There were no variations in these parameters as thrusters were throttled and single and multiple thruster operations were conducted. The propellant management system power consumption was at a fixed voltage to the DCIU and a fixed thermal throttle temperature of 75 C. The total power consumed by the PMS was 10.0, 17.9, and 25.2 W, respectively, for single, 2-thruster, and 3-thruster operation with the PMS. These sub-system integration tests of the PMS, the DCIU Simulator, and multiple thrusters addressed, in part, the NEXT PMS and propulsion system performance and functionality requirements.

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).

Installation noise measurements of model SR and CR propellers

NASA Technical Reports Server (NTRS)

Block, P. J. W.

1984-01-01

Noise measurements on a 0.1 scale SR-2 propeller in a single and counter rotation mode, in a pusher and tractor configuration, and operating at non-zero angles of attack are summarized. A measurement scheme which permitted 143 measurements of each of these configurations in the Langley 4- by 7-meter low speed tunnel is also described.

Underwater vehicle propulsion and power generation

NASA Technical Reports Server (NTRS)

Jones, Jack A. (Inventor); Chao, Yi (Inventor)

2008-01-01

An underwater vehicle includes a shaft with a propeller disposed thereon; a generator/motor having a stator and a rotor, the rotor being operable to rotate with the propeller; at least one energy storage device connected to the generator/motor; and a controller for setting the generator/motor in a charge mode, a propulsion mode and an idle mode.

76 FR 61558 - Airworthiness Directives; Dowty Propellers Type R212/4-30-4/22 and R251/4-30-4/49 Propeller...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-10-05

... the AD docket shortly after receipt. List of Subjects in 14 CFR Part 39 Air transportation, Aircraft... DEPARTMENT OF TRANSPORTATION Federal Aviation Administration 14 CFR Part 39 [Docket No. FAA-2011... Docket Operations office is located at Docket Management Facility, U.S. Department of Transportation...

The Green Propellant Infusion Mission Thruster Performance Testing for Plume Diagnostics

NASA Technical Reports Server (NTRS)

Deans, Matthew C.; Reed, Brian D.; Arrington, Lynn A.; Williams, George J.; Kojima, Jun J.; Kinzbach, McKenzie I.; McLean, Christopher H.

2014-01-01

The Green Propellant Infusion Mission (GPIM) is sponsored by NASA's Space Technology Mission Directorate (STMD) Technology Demonstration Mission (TDM) office. The goal of GPIM is to advance the technology readiness level of a green propulsion system, specifically, one using the monopropellant, AF-M315E, by demonstrating ground handling, spacecraft processing, and on-orbit operations. One of the risks identified for GPIM is potential contamination of sensitive spacecraft surfaces from the effluents in the plumes of AF-M315E thrusters. NASA Glenn Research Center (GRC) is conducting activities to characterize the effects of AF-M315E plume impingement and deposition. GRC has established individual plume models of the 22-N and 1-N thrusters that will be used on the GPIM spacecraft. The model simulations will be correlated with plume measurement data from Laboratory and Engineering Model 22-N, AF-M315E thrusters. The thrusters are currently being tested in a small rocket, altitude facility at NASA GRC. A suite of diagnostics, including Raman spectroscopy, Rayleigh spectroscopy, and Schlieren imaging are being used to acquire plume measurements of AF-M315E thrusters. Plume data will include temperature, velocity, relative density, and species concentration. The plume measurement data will be compared to the corresponding simulations of the plume model. The GRC effort will establish a data set of AF-M315E plume measurements and a plume model that can be used for future AF-M315E applications.

Performance of a low-power subsonic-arc-attachment arcjet thruster

NASA Technical Reports Server (NTRS)

Sankovic, John M.; Berns, Darren H.

1993-01-01

A subsonic-arc-attachment thruster design was scaled from a 30 kW 1960's vintage thruster to operate at nominally 3 kW. Performance measurements were obtained over a 1-4 kW power range using hydrogen as the propellant. Several modes of operation were identified and were characterized by varying degrees of voltage instability. A stability map was developed showing that the voltage oscillations were brought upon by elevated current or propellant levels. At a given specific energy level the specific impulse increased asymptotically with increased flow rates. Comparisons of performance were made between radial and tangential propellant injection. When the vortex flow was eliminated using radial injection, the operating voltages were lower at a given current, and the specific impulse and efficiency decreased. Tests were also conducted to determine the effects of background pressure on operation, and performance data were obtained at pressures of 0.047 Pa and 18 Pa. For a given specific energy level, the performance increased with a decrease in facility background pressure. Lowering the background pressure also caused a dramatic change in the voltage-current characteristic and the voltage stability, a phenomenon not previously reported with conventional supersonic-arc-attachment thrusters.

Hypervelocity Launching and Frozen Fuels as a Major Contribution to Spaceflight

NASA Astrophysics Data System (ADS)

Cocks, F. H.; Harman, C. M.; Klenk, P. A.; Simmons, W. N.

Acting as a virtual first stage, a hypervelocity launch together with the use of frozen hydrogen/frozen oxygen propellant, offers a Single-Stage-To-Orbit (SSTO) system that promises an enormous increase in SSTO mass-ratio. Ram acceleration provides hypervelocity (2 km/sec) to the orbital vehicle with a gas gun supplying the initial velocity required for ram operation. The vehicle itself acts as the center body of a ramjet inside a launch tube, filled with gaseous fuel and oxidizer, acting as an engine cowling. The high acceleration needed to achieve hypervelocity precludes a crew, and it would require greatly increased liquid fuel tank structural mass if a liquid propellant is used for post-launch vehicle propulsion. Solid propellants do not require as much fuel- chamber strengthening to withstand a hypervelocity launch as do liquid propellants, but traditional solid fuels have lower exhaust velocities than liquid hydrogen/liquid oxygen. The shock-stability of frozen hydrogen/frozen oxygen propellant has been experimentally demonstrated. A hypervelocity launch system using frozen hydrogen/frozen oxygen propellant would be a revolutionary new development in spaceflight.

Numerical Simulation of Cylindrical, Self-field MPD Thrusters with Multiple Propellants

NASA Technical Reports Server (NTRS)

Lapointe, Michael R.

1994-01-01

A two-dimensional, two-temperature, single fluid MHD code was used to predict the performance of cylindrical, self-field magnetoplasmadynamic (MPD) thrusters operated with argon, lithium, and hydrogen propellants. A thruster stability equation was determined relating maximum stable J(sup 2)/m values to cylindrical thruster geometry and propellant species. The maximum value of J(sup 2)/m was found to scale as the inverse of the propellant molecular weight to the 0.57 power, in rough agreement with limited experimental data which scales as the inverse square root of the propellant molecular weight. A general equation which relates total thrust to electromagnetic thrust, propellant molecular weight, and J(sup 2)/m was determined using reported thrust values for argon and hydrogen and calculated thrust values for lithium. In addition to argon, lithium, and hydrogen, the equation accurately predicted thrust for ammonia at sufficiently high J(sup 2)/m values. A simple algorithm is suggested to aid in the preliminary design of cylindrical, self-field MPD thrusters. A brief example is presented to illustrate the use of the algorithm in the design of a low power MPD thruster.

Summary of recent NASA propeller research

NASA Technical Reports Server (NTRS)

Mikkelson, D. C.; Mitchell, G. A.; Bober, L. J.

1984-01-01

Advanced high-speed propellers offer large performance improvements for aircraft that cruise in the Mach 0.7 to 0.8 speed regime. At these speeds, studies indicate that there is a 15 to near 40 percent block fuel savings and associated operating cost benefits for advanced turboprops compared to equivalent technology turbofan powered aircraft. Recent wind tunnel results for five eight to ten blade advanced models are compared with analytical predictions. Test results show that blade sweep was important in achieving net efficiencies near 80 percent at Mach 0.8 and reducing nearfield cruise noise by about 6 dB. Lifting line and lifting surface aerodynamic analysis codes are under development and some results are compared with propeller force and probe data. Also, analytical predictions are compared with some initial laser velocimeter measurements of the flow field velocities of an eightbladed 45 swept propeller. Experimental aeroelastic results indicate that cascade effects and blade sweep strongly affect propeller aeroelastic characteristics. Comparisons of propeller near-field noise data with linear acoustic theory indicate that the theory adequately predicts near-field noise for subsonic tip speeds but overpredicts the noise for supersonic tip speeds.

Summary of recent NASA propeller research

NASA Technical Reports Server (NTRS)

Mikkelson, D. C.; Mitchell, G. A.; Bober, L. J.

1985-01-01

Advanced high speed propellers offer large performance improvements for aircraft that cruise in the Mach 0.7 to 0.8 speed regime. At these speeds, studies indicate that there is a 15 to near 40 percent block fuel savings and associated operating cost benefits for advanced turboprops compared to equivalent technology turbofan powered aircraft. Recent wind tunnel results for five eight to ten blade advanced models are compared with analytical predictions. Test results show that blade sweep was important in achieving net efficiencies near 80 percent at Mach 0.8 and reducing nearfield cruise noise about 6 dB. Lifting line and lifting surface aerodynamic analysis codes are under development and some results are compared with propeller force and probe data. Also, analytical predictions are compared with some initial laser velocimeter measurements of the flow field velocities of an eight bladed 45 swept propeller. Experimental aeroelastic results indicate that cascade effects and blade sweep strongly affect propeller aeroelastic characteristics. Comparisons of propeller nearfield noise data with linear acoustic theory indicate that the theory adequately predicts nearfield noise for subsonic tip speeds, but overpredicts the noise for supersonic tip speeds.

The 3-D viscous flow CFD analysis of the propeller effect on an advanced ducted propeller subsonic inlet

NASA Technical Reports Server (NTRS)

Iek, Chanthy; Boldman, Donald R.; Ibrahim, Mounir

1993-01-01

A time marching Navier-Stokes code called PARC3D was used to study the 3-D viscous flow associated with an advanced ducted propeller (ADP) subsonic inlet at take-off operating conditions. At a free stream Mach number of 0.2, experimental data for the inlet-with-propeller test model indicated that the airflow was attached on the cowl windward lip at an angle of attack of 25 degrees became unstable at 29 degrees, and separated at 30 degrees. An experimental study with a similar inlet and with no propeller (through-flow) indicated that flow separation occurred at an angle of attack a few degrees below the value observed when the inlet was tested with the propeller. This tends to indicate that the propeller exerts a favorable effect on the inlet performance. During the through-flow experiment a stationary blockage device was used to successfully simulate the propeller effect on the inlet flow field at angles of attack. In the present numerical study, this flow blockage was modeled via a PARC3D computational boundary condition (BC) called the screen BC. The principle formulation of this BC was based on the one-and-half dimension actuator disk theory. This screen BC was applied at the inlet propeller face station of the computational grid. Numerical results were obtained with and without the screen BC. The application of the screen BC in this numerical study provided results which are similar to the results of past experimental efforts in which either the blockage device or the propeller was used.

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.

Geometric effects in applied-field MPD thrusters

NASA Technical Reports Server (NTRS)

Myers, R. M.; Mantenieks, M.; Sovey, J.

1990-01-01

Three applied-field magnetoplasmadynamic (MPD) thruster geometries were tested with argon propellant to establish the influence of electrode geometry on thruster performance. The thrust increased approximately linearly with anode radius, while the discharge and electrode fall voltages increased quadratically with anode radius. All these parameters increased linearly with applied-field strength. Thrust efficiency, on the other hand, was not significantly influenced by changes in geometry over the operating range studied, though both thrust and thermal efficiencies increased monotonically with applied field strength. The best performance, 1820 sec I (sub sp) at 20 percent efficiency, was obtained with the largest radius anode at the highest discharge current (1500 amps) and applied field strength (0.4 Tesla).

Geometric effects in applied-field MPD thrusters

NASA Technical Reports Server (NTRS)

Myers, R. M.; Mantenieks, M.; Sovey, James S.

1990-01-01

Three applied-field magnetoplasmadynamic (MPD) thruster geometries were tested with argon propellant to establish the influence of electrode geometry on thruster performance. The thrust increased approximately linearly with anode radius, while the discharge and electrode fall voltages increased quadratically with anode radius. All these parameters increased linearly with applied-field strength. Thrust efficiency, on the other hand, was not significantly influenced by changes in geometry over the operating range studied, though both thrust and thermal efficiencies increased monotonically with applied field strength. The best performance, 1820 sec I(sub sp) at 20 percent efficiency, was obtained with the largest radius anode at the highest discharge current (1500 amps) and applied field strength (0.4 Tesla).

14 CFR 33.51 - Operation test.

Code of Federal Regulations, 2012 CFR

2012-01-01

... STANDARDS: AIRCRAFT ENGINES Block Tests; Reciprocating Aircraft Engines § 33.51 Operation test. The... characteristics, starting, idling, acceleration, overspeeding, functioning of propeller and ignition, and any other operational characteristic of the engine. If the engine incorporates a multispeed supercharger...

14 CFR 33.51 - Operation test.

Code of Federal Regulations, 2011 CFR

2011-01-01

... STANDARDS: AIRCRAFT ENGINES Block Tests; Reciprocating Aircraft Engines § 33.51 Operation test. The... characteristics, starting, idling, acceleration, overspeeding, functioning of propeller and ignition, and any other operational characteristic of the engine. If the engine incorporates a multispeed supercharger...

Space shuttle main engine: Interactive design challenges

NASA Technical Reports Server (NTRS)

Mccarty, J. P.; Wood, B. K.

1985-01-01

The operating requirements established by NASA for the SSME were considerably more demanding than those for earlier rocket engines used in the military launch vehicles or Apollo program. The SSME, in order to achieve the high performance, low weight, long life, reusable objectives, embodied technical demands far in excess of its predecessor rocket engines. The requirements dictated the use of high combustion pressure and the staged combustion cycle which maximizes performance through total use of all propellants in the main combustion process. This approach presented a myriad of technical challenges for maximization of performance within attainable state of the art capabilities for operating pressures, operating temperatures and rotating machinery efficiencies. Controlling uniformity of the high pressure turbomachinery turbine temperature environment was a key challenge for thrust level and life capability demanding innovative engineering. New approaches in the design of the components were necessary to accommodate the multiple use, minimum maintenance objectives. Included were the use of line replaceable units to facilitate field maintenance automatic checkout and internal inspection capabilities.

Cryogenic, high speed, turbopump bearing cooling requirements

NASA Technical Reports Server (NTRS)

Dolan, Fred J.; Gibson, Howard G.; Cannon, James L.; Cody, Joe C.

1988-01-01

Although the Space Shuttle Main Engine (SSME) has repeatedly demonstrated the capability to perform during launch, the High Pressure Oxidizer Turbopump (HPOTP) main shaft bearings have not met their 7.5 hour life requirement. A tester is being employed to provide the capability of subjecting full scale bearings and seals to speeds, loads, propellants, temperatures, and pressures which simulate engine operating conditions. The tester design permits much more elaborate instrumentation and diagnostics than could be accommodated in an SSME turbopump. Tests were made to demonstrate the facilities; and the devices' capabilities, to verify the instruments in its operating environment and to establish a performance baseline for the flight type SSME HPOTP Turbine Bearing design. Bearing performance data from tests are being utilized to generate: (1) a high speed, cryogenic turbopump bearing computer mechanical model, and (2) a much improved, very detailed thermal model to better understand bearing internal operating conditions. Parametric tests were also made to determine the effects of speed, axial loads, coolant flow rate, and surface finish degradation on bearing performance.

Investigation of the part-load performance of two 1.12 MW regenerative marine gas turbines

NASA Astrophysics Data System (ADS)

Korakianitis, T.; Beier, K. J.

1994-04-01

Regenerative and intercooled-regenerative gas turbine engines with low pressure ratio have significant efficiency advantages over traditional aero-derivative engines of higher pressure ratios, and can compete with modern diesel engines for marine propulsion. Their performance is extremely sensitive to thermodynamic-cycle parameter choices and the type of components. The performances of two 1.12 MW (1500 hp) regenerative gas turbines are predicted with computer simulations. One engine has a single-shaft configuration, and the other has a gas-generator/power-turbine combination. The latter arrangement is essential for wide off-design operating regime. The performance of each engine driving fixed-pitch and controllable-pitch propellers, or an AC electric bus (for electric-motor-driven propellers) is investigated. For commercial applications the controllable-pitch propeller may have efficiency advantages (depending on engine type and shaft arrangements). For military applications the electric drive provides better operational flexibility.

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.

Linear aerospike engine study. [for reusable launch vehicles

NASA Technical Reports Server (NTRS)

Diem, H. G.; Kirby, F. M.

1977-01-01

Parametric data on split-combustor linear engine propulsion systems are presented for use in mixed-mode single-stage-to-orbit (SSTO) vehicle studies. Preliminary design data for two selected engine systems are included. The split combustor was investigated for mixed-mode operations with oxygen/hydrogen propellants used in the inner combustor in Mode 2, and in conjunction with either oxygen/RP-1, oxygen/RJ-5, O2/CH4, or O2/H2 propellants in the outer combustor for Mode 1. Both gas generator and staged combustion power cycles were analyzed for providing power to the turbopumps of the inner and outer combustors. Numerous cooling circuits and cooling fluids (propellants) were analyzed and hydrogen was selected as the preferred coolant for both combustors and the linear aerospike nozzle. The maximum operating chamber pressure was determined to be limited by the availability of hydrogen coolant pressure drop in the coolant circuit.

The Propeller and Cooling-Air-Flow Characteristics of a Twin-Engine Airplane Model Equipped with NACA D sub s -Type Cowlings and with Propellers of NACA 16-Series Airfoil Sections

DTIC Science & Technology

1944-09-01

with the cowling flaps neutral, did not in any case exceed T] = ±0.03. Drag and Cowling-Air Plow with Propeller Removed The effects, on the lift...cowling flaps. Effect of internal flow on drar.- For convenience in studying the drf.g oharaoterlstio a of the two cowling arrangement•, values of the...operation and take-off. Influence of Cooling Hequireir;ent3 on Airplane Performance In the case of many conventional radial elr-ccoled engine

Aerodynamic design of the Cal Poly Da Vinci Human-Powered Helicopter

NASA Technical Reports Server (NTRS)

Larwood, Scott; Saiki, Neal

1990-01-01

This paper will discuss the methodology used in designing the rotor and drive propellers for the third generation Cal Poly Da Vinci Human-Powered Helicopter. The rotor was designed using a lifting surface, uniform inflow hover analysis code and the propeller was designed using a minimum induced-loss method. Construction, geometry, and operating considerations are discussed as they impact the designs. Optimization of the design performance is also explained. The propellers were tested in a wind tunnel and results are compared with theoretical data. Successful flight tests of the Da Vinci III are discussed.

Aero Propulsion and Power Directorate The McCook Field Years (1917-1927)

DTIC Science & Technology

1995-02-01

34 All of the works cited refer to organizational highlights to provide perspective, but four other works supply detailed analysis. Royal D. Frey...offered higher speed and greater power. Late in WWI, the Royal Aircraft Establishment in England began research on air-cooled cylinders. In 1915 and...obtain with wooden ones. Steel propellers did not tear out at the hub or split at the 113 weld. The Royal Air Force used steel propellers extensively

Numerical Simulation of Flow in Fluidic Valves in Rotating Detonation Engines

NASA Astrophysics Data System (ADS)

Gopalakrishnan, Nandini

Rotating detonation engines (RDE) have received considerable research attention in recent times for use in propulsion systems. The cycle frequency of operation of an RDE can be as high as 10,000 Hz. Conventional mechanical valves cannot operate at such high frequencies, leading to the need for propellant injectors or valves with no moving parts. A fluidic valve is such a valve and is the focus of this study. The valve consists of an orifice connected to a constant area plenum cavity which operates at constant pressure. The fluidic valve supplies propellants to the detonation tube through the orifice. Hydrogen - oxygen detonation is studied in a tube with fluidic valves. A detailed 19-step chemical reaction mechanism has been used to model detonation and the flow simulated in ANSYS Fluent. This research aims to determine the location of contact surface in the cavity and the time taken for the contact surface to leave the valve after a shock wave has passed through it. This will help us understand if the steady-state flow in the cavity is comprised of detonation products or fresh propellants.

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.

High speed propeller performance and noise predictions at takeoff/landing conditions

NASA Technical Reports Server (NTRS)

Nallasamy, M.; Woodward, R. P.; Groeneweg, J. F.

1988-01-01

The performance and noise of a high speed SR-7A model propeller under takeoff/landing conditions are considered. The blade loading distributions are obtained by solving the three-dimensional Euler equations and the sound pressure levels are computed using a time domain approach. At the nominal takeoff operating point, the blade sections near the hub are lightly or negatively loaded. The chordwise loading distributions are distinctly different from those of cruise conditions. The noise of the SR-7A model propeller at takeoff is dominated by the loading noise, similar to that at cruise conditions. The waveforms of the acoustic pressure signature are nearly sinusoidal in the plane of the propeller. The computed directivity of the blade passing frequency tone agrees fairly well with the data at nominal takeoff blade angle.

High speed propeller performance and noise predictions at takeoff/landing conditions

NASA Technical Reports Server (NTRS)

Nallasamy, M.; Woodward, R. P.; Groeneweg, J. F.

1987-01-01

The performance and noise of a high speed SR-7A model propeller under takeoff/landing conditions are considered. The blade loading distributions are obtained by solving the three-dimensional Euler equations and the sound pressure levels are computed using a time domain approach. At the nominal takeoff operating point, the blade sections near the hub are lightly or negatively loaded. The chordwise loading distributions are distinctly different from those of cruise conditions. The noise of the SR-7A model propeller at takeoff is dominated by the loading noise, similar to that at cruise conditions. The waveforms of the acoustic pressure signature are nearly sinusoidal in the plane of the propeller. The computed directivity of the blade passing frequency tone agrees fairly well with the data at nominal takeoff blade angle.

RSRM-13 (360Q013) ballistics mass properties flight designation STS-41

NASA Technical Reports Server (NTRS)

Laubacher, Brian A.; Richards, M. C.

1990-01-01

The propulsion performance and reconstructed mass properties data from Thiokol's RSRM-13 motors which were assigned to the STS-41 launch are presented. The SRM propellant, TP-H1148, is a composite type solid propellant, formulated of polybutadiene acrylic acid acryonitrile terpolymer binder, epoxy curing agent, ammonium perchlorate oxidizer, and aluminum powder fuel. A small amount of burning rate catalyst (iron oxide) was added to achieve the desired propellant burn rate. The propellant evaluation and raw material information are also presented. The presented ballistic performance was based on the Operational Flight Instrumentation. The adjustments made to the raw data on this flight include biasing the data to correct ambient pressure before liftoff. The performance from each motor as well as matched pair performance values were well within the CEI Specification requirements.

Effect of the Thruster Configurations on a Laser Ignition Microthruster

NASA Astrophysics Data System (ADS)

Koizumi, Hiroyuki; Hamasaki, Kyoichi; Kondo, Ryo; Okada, Keisuke; Nakano, Masakatsu; Arakawa, Yoshihiro

Research and development of small spacecraft have advanced extensively throughout the world and propulsion devices suitable for the small spacecraft, microthruster, is eagerly anticipated. The authors proposed a microthruster using 1—10-mm-size solid propellant. Small pellets of solid propellant are installed in small combustion chambers and ignited by the irradiation of diode laser beam. This thruster is referred as to a laser ignition microthruster. Solid propellant enables large thrust capability and compact propulsion system. To date theories of a solid-propellant rocket have been well established. However, those theories are for a large-size solid propellant and there are a few theories and experiments for a micro-solid rocket of 1—10mm class. This causes the difficulty of the optimum design of a micro-solid rocket. In this study, we have experimentally investigated the effect of thruster configurations on a laser ignition microthruster. The examined parameters are aperture ratio of the nozzle, length of the combustion chamber, area of the nozzle throat, and divergence angle of the nozzle. Specific impulse dependences on those parameters were evaluated. It was found that large fraction of the uncombusted propellant was the main cause of the degrading performance. Decreasing the orifice diameter in the nozzle with a constant open aperture ratio was an effective method to improve this degradation.

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.

Adiabatic Compression Sensitivity of Liquid Fuels and Monopropellants

NASA Technical Reports Server (NTRS)

Ismail, Ismail M. K.; Hawkins, Tom W.

2000-01-01

Liquid rocket propellants can be sensitive to rapid compression. Such liquids may undergo decomposition and their handling may be accompanied with risk. Decomposition produces small gas bubbles in the liquid, which upon rapid compression may cause catastrophic explosions. The rapid compression can result from mechanical shocks applied on the tank containing the liquid or from rapid closure of the valves installed on the lines. It is desirable to determine the conditions that may promote explosive reactions. At Air Force Research Laboratory (AFRL), we constructed an apparatus and established a safe procedure for estimating the sensitivity of propellant materials towards mechanical shocks (Adiabatic Compression Tester). A sample is placed on a stainless steel U-tube, held isothermally at a temperature between 20 and 150 C then exposed to an abrupt mechanical shock of nitrogen gas at a pressure between 6.9 and 20.7 MPa (1000 to 3000 psi). The apparatus is computer interfaced and is driven with LABTECH NOTEBOOK-pro (registered) Software. In this presentation, the design of the apparatus is shown, the operating procedure is outlined, and the safety issues are addressed. The results obtained on different energetic materials are presented.

Modeling of Multi-Tube Pulse Detonation Engine Operation

NASA Technical Reports Server (NTRS)

Ebrahimi, Houshang B.; Mohanraj, Rajendran; Merkle, Charles L.

2001-01-01

The present paper explores some preliminary issues concerning the operational characteristics of multiple-tube pulsed detonation engines (PDEs). The study is based on a two-dimensional analysis of the first-pulse operation of two detonation tubes exhausting through a common nozzle. Computations are first performed to assess isolated tube behavior followed by results for multi-tube flow phenomena. The computations are based on an eight-species, finite-rate transient flow-field model. The results serve as an important precursor to understanding appropriate propellant fill procedures and shock wave propagation in multi-tube, multi-dimensional simulations. Differences in behavior between single and multi-tube PDE models are discussed, The influence of multi-tube geometry and the preferred times for injecting the fresh propellant mixture during multi-tube PDE operation are studied.

Thrust Breakdown Characteristics of Conventional Propellers

DTIC Science & Technology

2007-09-01

extends beyond the trailing edge of the blade . These sheets violently collapse as the blade moves out of the wake deficit produced by the hull. This...thrust breakdown, vibration, noise , erosion and blade damage. Propellers operating with enough cavitation to cause thrust breakdown can experience...7 Figure 5. Sensitivity of thrust reduction to harmonic content in wake (Prop 5491) .................. 8 Figure 6. Comparison of

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.

Investigation of Engine Oil-cooling Problem during Idle Conditions on Pusher Type Turbo Prop Aircraft

NASA Astrophysics Data System (ADS)

Premkumar, P. S.; Chakravarthy, S. Bhaskar; Jayagopal, S.; Radhakrishnan, P.; Pillai, S. Nadaraja; Senthil Kumar, C.

2017-11-01

Aircraft engines need a cooling system to keep the engine oil well within the temperature limits for continuous operation. The aircraft selected for this study is a typical pusher type Light Transport Aircraft (LTA) having twin turbo prop engines mounted at the aft end of the fuselage. Due to the pusher propeller configuration, effective oil cooling is a critical issue, especially during low-speed ground operations like engine idling and also in taxiing and initial climb. However, the possibility of utilizing the inflow induced by the propeller for oil cooling is the subject matter of investigation in this work. The oil cooler duct was designed to accommodate the required mass flow, estimated using the oil cooler performance graph. A series of experiments were carried out with and without oil cooler duct attached to the nacelle, in order to investigate the mass flow induced by the propeller and its adequacy to cool the engine oil. Experimental results show that the oil cooler positioned at roughly 25 % of the propeller radius from the nacelle center line leads to adequate cooling, without incorporating additional means. Furthermore, it is suggested to install a NACA scoop to minimize spillage drag by increasing pressure recovery.

A Flight Demonstration of Plasma Rocket Propulsion

NASA Technical Reports Server (NTRS)

Petro, Andrew; Chang-Diaz, Franklin; Schwenterly, WIlliam; Hitt, Michael; Lepore, Joseph

2000-01-01

The Advanced Space Propulsion Laboratory at the NASA Johnson Space Center has been engaged in the development of a variable specific impulse magnetoplasma rocket (V ASIMR) for several years. This type of rocket could be used in the future to propel interplanetary spacecraft and has the potential to open the entire solar system to human exploration. One feature of this propulsion technology is the ability to vary its specific impulse so that it can be operated in a mode that maximizes propellant efficiency or a mode that maximizes thrust. Variation of specific impulse and thrust enhances the ability to optimize interplanetary trajectories and results in shorter trip times and lower propellant requirements than with a fixed specific impulse. In its ultimate application for interplanetary travel, the VASIMR would be a multi-megawatt device. A much lower power system is being designed for demonstration in the 2004 timeframe. This first space demonstration would employ a lO-kilowatt thruster aboard a solar powered spacecraft in Earth orbit. The 1O-kilowatt V ASIMR demonstration unit would operate for a period of several months with hydrogen or deuterium propellant with a specific impulse of 10,000 seconds.

14 CFR 25.1521 - Powerplant limitations.

Code of Federal Regulations, 2010 CFR

2010-01-01

... propellers are type certificated and do not exceed the values on which compliance with any other requirement... following must be established for reciprocating engine installations: (1) Horsepower or torque, r.p.m...) Any other parameter for which a limitation has been established as part of the engine type certificate...

14 CFR 91.815 - Agricultural and fire fighting airplanes: Noise operating limitations.

Code of Federal Regulations, 2014 CFR

2014-01-01

...: Noise operating limitations. 91.815 Section 91.815 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION... RULES Operating Noise Limits § 91.815 Agricultural and fire fighting airplanes: Noise operating limitations. (a) This section applies to propeller-driven, small airplanes having standard airworthiness...

Development and Optimized Design of Propeller Pump System & Structure with VFD in Low-head Pumping Station

NASA Astrophysics Data System (ADS)

Rentian, Zhang; Honggeng, Zhu; Arnold, Jaap; Linbi, Yao

2010-06-01

Compared with vertical-installed pumps, the propeller (bulb tubular) pump systems can achieve higher hydraulic efficiencies, which are particularly suitable for low-head pumping stations. More than four propeller pumping stations are being, or will be built in the first stage of the S-to-N Water Diversion Project in China, diverting water from Yangtze River to the northern part of China to alleviate water-shortage problems and develop the economy. New structures of propeller pump have been developed for specified pumping stations in Jiangsu and Shandong Provinces respectively and Variable Frequency Drives (VFDs) are used in those pumping stations to regulate operating conditions. Based on the Navier-Stokes equations and the standard k-e turbulent model, numerical simulations of the flow field and performance prediction in the propeller pump system were conducted on the platform of commercial software CFX by using the SIMPLEC algorithm. Through optimal design of bulb dimensions and diffuser channel shape, the hydraulic system efficiency has improved evidently. Furthermore, the structures of propeller pumps have been optimized to for the introduction of conventional as well as permanent magnet motors. In order to improve the hydraulic efficiency of pumping systems, both the pump discharge and the motor diameter were optimized respectively. If a conventional motor is used, the diameter of the pump casing has to be increased to accommodate the motor installed inside. If using a permanent magnet motor, the diameter of motor casing can be decreased effectively without decreasing its output power, thus the cross-sectional area is enlarged and the velocity of flowing water decreased favorably to reduce hydraulic loss of discharge channel and thereby raising the pumping system efficiency. Witness model tests were conducted after numerical optimization on specific propeller pump systems, indicating that the model system hydraulic efficiencies can be improved by 0.5%˜3.7% in different specified operational conditions.

TNT equivalency study for space shuttle (EOS). Volume 1: Management summary report

NASA Technical Reports Server (NTRS)

Wolfe, R. R.

1971-01-01

The existing TNT equivalency criterion for LO2/LH2 propellant is reevaluated. It addresses the static, on-pad phase of the space shuttle launch operations and was performed to determine whether the use of a TNT equivalency criterion lower than that presently used (60%) could be substantiated. The large quantity of propellant on-board the space shuttle, 4 million pounds, was considered of prime importance to the study. A qualitative failure analysis of the space shuttle (EOS) on the launch pad was made because it was concluded that available test data on the explosive yield of LO2/LH2 propellant was insufficient to support a reduction in the present TNT equivalency value, considering the large quantity of propellant used in the space shuttle. The failure analysis had two objectives. The first was to determine whether a failure resulting in the total release of propellant could occur. The second was to determine whether, if such a failure did occur, ignition could be delayed long enough to allow the degree of propellant mixing required to produce an explosion of 60% TNT equivalency since the explosive yield of this propellant is directly related to the quantities of LH2 and LO2 mixed at the time of the explosion.

Performance Tests of a Liquid Hydrogen Propellant Densification Ground System for the X33/RLV

NASA Technical Reports Server (NTRS)

Tomsik, Thomas M.

1997-01-01

A concept for improving the performance of propulsion systems in expendable and single-stage-to-orbit (SSTO) launch vehicles much like the X33/RLV has been identified. The approach is to utilize densified cryogenic liquid hydrogen (LH2) and liquid oxygen (LOX) propellants to fuel the propulsion stage. The primary benefit for using this relatively high specific impulse densified propellant mixture is the subsequent reduction of the launch vehicle gross lift-off weight. Production of densified propellants however requires specialized equipment to actively subcool both the liquid oxygen and liquid hydrogen to temperatures below their normal boiling point. A propellant densification unit based on an external thermodynamic vent principle which operates at subatmospheric pressure and supercold temperatures provides a means for the LH2 and LOX densification process to occur. To demonstrate the production concept for the densification of the liquid hydrogen propellant, a system comprised of a multistage gaseous hydrogen compressor, LH2 recirculation pumps and a cryogenic LH2 heat exchanger was designed, built and tested at the NASA Lewis Research Center (LeRC). This paper presents the design configuration of the LH2 propellant densification production hardware, analytical details and results of performance testing conducted with the hydrogen densifier Ground Support Equipment (GSE).