Results for “Propellant” — Page 47

NASA NTRSDate not supplied

Effect of Engine Thrust and Isp Tradeoffs and Alternate Propellants on ΔV Budget and Architecture Mass for 1st Generation Nuclear Thermal Propulsion Flight Test Systems

Following the first Nuclear Thermal Propulsion (NTP) system test, also known as DRACO, the next NTP system to be developed would be the 1st Generation NTP system. An analysis was conducted to determine the performance of different vehicle configurations utilizing hydrogen (H-NTP), ammonia (A-NTP), and methane (M-NTP) as propellants launched onboard commercial launch vehicles. This was enabled by a quasi-steady-state power balance engine model and vehicle component physics that sized the vehicle system using Master Equipment List (MEL) parameters. The analysis considered configuration cases outlined by a matrix of different mission classes and vehicle configurations that covered the design space of the 1st Generation NTP system to explore various propellant options, engine architectures, and mission scopes. Four mission classes were evaluated which included single burn missions performing maneuvers having a ΔV of 1 km/s and 2.5 km/s followed by 2-burn and 4-burn missions that aimed to exhaust the launch vehicles volume or mass limitations. In all cases, the NTP systems with the lowest thrust class had the longest burn time of which A-NTP and M-NTP systems provided the longest and shortest burn times depending on the launch vehicle used while H-NTP systems tended to cluster together in the middle. Longer burn times could be useful as a testing platform to increase the time for data accumulation. Across the multi-burn cases, H-NTP systems were found to be volume limited while A-NTP and M-NTP systems were mass limited. Both single burn missions showed that A-NTP configurations provided the lowest dry mass given that ammonia had the highest density with comparable performance to M-NTP systems and no requirement for cryocoolers. The results showed that beyond the propellant selection type, the launch vehicle selection, which included Starship, New Glenn, Vulcan, Falcon 9 (recoverable), and Falcon Heavy (recoverable), was a primary driving factor in the test vehicle’s capabilities. Trends were determined based on a set of dimensionless parameters that included the mass ratio of inert mass to initial wetted mass, ratio of specific impulse to burn time, and a dimensionless engine parameter (ratio of impulse to engine momentum). These relationships found the “knees-in-the-curves” that could be a significant point of reference for the designer as they indicate a change in the trend which is located at a specific impulse to burn time ratio of 1, a mass ratio of 0.6, and an engine performance parameter of 4. This study did not attempt to make a recommendation rather provide the tools for the reader to select their own configuration based on their needs.

Propellant↗

NASA NTRSDate not supplied

A Comparative Assessment of Evolvable Martian ISRU Propellant Production for Future Human Missions

In 2022, NASA released its “Moon to Mars Strategy and Objectives Development” document, containing top-level goals and objectives designed to create a blueprint for sustained human presence and exploration throughout the solar system. One objective in this “blueprint” calls for the demonstration of “…Mars ISRU capabilities to support an initial human Mars exploration campaign.” NASA is looking at both the benefits and liens of this concept, as one of several trade space aspects under consideration, for inclusion in the overall architecture of the human exploration of Mars. This paper summarizes results from a series of Martian ISRU-focused trade studies from 2023 and 2024. These studies cover five different approaches to providing propellants for returning crew and equipment from the surface of Mars, using several different combinations of propellants delivered (from Earth) and produced from Martian feedstock material. For comparison purposes, the same ISRU infrastructure elements were used across all options wherever possible. The equipment mass, power, operational complexity, and time needed to produce some or all of these propellants have been quantified and compared when used in a common mission scenario. Analysis results for each option are discussed. Performance parameters for these options are compared side-by-side and observations made, including the positive and negative aspects of each case. The possibility of evolving from a simpler ISRU case to more complex cases that would potentially lead to more sustainable operations are also discussed. While no single case emerged as “better” in all aspects when compared to the others, there are distinct advantages and disadvantages that did emerge when comparing the results generated from a common set of assumptions for all cases. Consequently, specific mission objectives may determine which of these advantages or disadvantages are important for choosing among the cases that could benefit any particular mission or campaign.

Mars↗

NASA NTRSDate not supplied

New Solid Propellant Additives: Evaluation for NASA Mission Use Cases

Researchers and small companies have been investigating and developing new solid rocket motor propellant ingredients with renewed vigor. Lithium-aluminum alloy is an ingredient of particular interest, but most of the advertisement about its benefits are focused on defense tactical rocket applications. This paper develops a framework to evaluate this and other propellants against standard aluminized composite propellant for mass-limited missions like many NASA in-space applications. It shows a way to estimate density’s effect with an exponent less than 1 depending on mission and technology parameters, and an algorithm for directly calculating delta-velocity or payload performance improvement.

Lithium-aluminum↗

NASA NTRSDate not supplied

A Comparative Assessment of Evolvable Martian ISRU Propellant Production for Future Human Missions

In 2022, NASA released its “Moon to Mars Strategy and Objectives Development” document, containing top-level goals and objectives designed to create a blueprint for sustained human presence and exploration throughout the solar system. One objective in this “blueprint” calls for the demonstration of “…Mars ISRU capabilities to support an initial human Mars exploration campaign.” NASA is looking at both the benefits and liens of this concept, as one of several trade space aspects under consideration, for inclusion in the overall architecture of the human exploration of Mars. This paper summarizes results from a series of Martian ISRU-focused trade studies from 2023 and 2024. These studies cover five different approaches to providing propellants for returning crew and equipment from the surface of Mars, using several different combinations of propellants delivered (from Earth) and produced from Martian feedstock material. For comparison purposes, the same ISRU infrastructure elements were used across all options wherever possible. The equipment mass, power, operational complexity, and time needed to produce some or all of these propellants have been quantified and compared when used in a common mission scenario. Analysis results for each option are discussed. Performance parameters for these options are compared side-by-side and observations made, including the positive and negative aspects of each case. The possibility of evolving from a simpler ISRU case to more complex cases that would potentially lead to more sustainable operations are also discussed. While no single case emerged as “better” in all aspects when compared to the others, there are distinct advantages and disadvantages that did emerge when comparing the results generated from a common set of assumptions for all cases. Consequently, specific mission objectives may determine which of these advantages or disadvantages are important for choosing among the cases that could benefit any particular mission or campaign.

Propellant↗

NASA NTRSDate not supplied

Effect of Engine Thrust and Isp Tradeoffs and Alternate Propellants on ΔV Budget and Architecture Mass for 1st Generation Nuclear Thermal Propulsion Flight Test Systems

Following the first Nuclear Thermal Propulsion (NTP) system test, also known as DRACO, the next NTP system to be developed would be the 1st Generation NTP system. An analysis was conducted to determine the performance of different vehicle configurations utilizing hydrogen (H-NTP), ammonia (A-NTP), and methane (M-NTP) as propellants launched onboard commercial launch vehicles. This was enabled by a quasi-steady-state power balance engine model and vehicle component physics that sized the vehicle system using Master Equipment List (MEL) parameters. The analysis considered configuration cases outlined by a matrix of different mission classes and vehicle configurations that covered the design space of the 1st Generation NTP system to explore various propellant options, engine architectures, and mission scopes. Four mission classes were evaluated which included single burn missions performing maneuvers having a ΔV of 1 km/s and 2.5 km/s followed by 2-burn and 4-burn missions that aimed to exhaust the launch vehicles volume or mass limitations. In all cases, the NTP systems with the lowest thrust class had the longest burn time of which A-NTP and M-NTP systems provided the longest and shortest burn times depending on the launch vehicle used while H-NTP systems tended to cluster together in the middle. Longer burn times could be useful as a testing platform to increase the time for data accumulation. Across the multi-burn cases, H-NTP systems were found to be volume limited while A-NTP and M-NTP systems were mass limited. Both single burn missions showed that A-NTP configurations provided the lowest dry mass given that ammonia had the highest density with comparable performance to M-NTP systems and no requirement for cryocoolers. The results showed that beyond the propellant selection type, the launch vehicle selection, which included Starship, New Glenn, Vulcan, Falcon 9 (recoverable), and Falcon Heavy (recoverable), was a primary driving factor in the test vehicle’s capabilities. Trends were determined based on a set of dimensionless parameters that included the mass ratio of inert mass to initial wetted mass, ratio of specific impulse to burn time, and a dimensionless engine parameter (ratio of impulse to engine momentum). These relationships found the “knees-in-the-curves” that could be a significant point of reference for the designer as they indicate a change in the trend which is located at a specific impulse to burn time ratio of 1, a mass ratio of 0.6, and an engine performance parameter of 4. This study did not attempt to make a recommendation rather provide the tools for the reader to select their own configuration based on their needs.

Propellant↗

NASA NTRSDate not supplied

Effect of Engine Thrust and Isp Tradeoffs and Alternate Propellants on ΔV Budget and Architecture Mass for 1st Generation Nuclear Thermal Propulsion Flight Test Systems

Following the first Nuclear Thermal Propulsion (NTP) system test, also known as DRACO, the next NTP system to be developed would be the 1st Generation NTP system. An analysis was conducted to determine the performance of different vehicle configurations utilizing hydrogen (H-NTP), ammonia (A-NTP), and methane (M-NTP) as propellants launched onboard commercial launch vehicles. This was enabled by a quasi-steady-state power balance engine model and vehicle component physics that sized the vehicle system using Master Equipment List (MEL) parameters. The analysis considered configuration cases outlined by a matrix of different mission classes and vehicle configurations that covered the design space of the 1st Generation NTP system to explore various propellant options, engine architectures, and mission scopes. Four mission classes were evaluated which included single burn missions performing maneuvers having a ΔV of 1 km/s and 2.5 km/s followed by 2-burn and 4-burn missions that aimed to exhaust the launch vehicles volume or mass limitations. In all cases, the NTP systems with the lowest thrust class had the longest burn time of which A-NTP and M-NTP systems provided the longest and shortest burn times depending on the launch vehicle used while H-NTP systems tended to cluster together in the middle. Longer burn times could be useful as a testing platform to increase the time for data accumulation. Across the multi-burn cases, H-NTP systems were found to be volume limited while A-NTP and M-NTP systems were mass limited. Both single burn missions showed that A-NTP configurations provided the lowest dry mass given that ammonia had the highest density with comparable performance to M-NTP systems and no requirement for cryocoolers. The results showed that beyond the propellant selection type, the launch vehicle selection, which included Starship, New Glenn, Vulcan, Falcon 9 (recoverable), and Falcon Heavy (recoverable), was a primary driving factor in the test vehicle’s capabilities. Trends were determined based on a set of dimensionless parameters that included the mass ratio of inert mass to initial wetted mass, ratio of specific impulse to burn time, and a dimensionless engine parameter (ratio of impulse to engine momentum). These relationships found the “knees-in-the-curves” that could be a significant point of reference for the designer as they indicate a change in the trend which is located at a specific impulse to burn time ratio of 1, a mass ratio of 0.6, and an engine performance parameter of 4. This study did not attempt to make a recommendation rather provide the tools for the reader to select their own configuration based on their needs.

Propellant↗

NASA NTRS1926

Test of a model propeller with symmetrical blade sections

This report, prepared at the request of NACA, gives the results of tests on a model propeller having blade sections with form of Gottingen airfoil no. 409. The model is shown to have a dynamic pitch practically equal to the nominal or geometrical pitch, and a somewhat higher efficiency but lower coefficient than would be expected of a propeller of more conventional sections.

PROPELLERS - CHARACTERISTICS - WIND TUNNEL TESTS↗

NASA NTRS1944

Tests of Thermal-electric De-icing Equipment for Propellers

Flights were made in natural icing conditions at the NACA Ice Research Project, Minneapolis, Minn. to test several designs of thermal-electric propeller de-icing blade shoes and a hub-generator design. It was found that a minimum average unit power of 2.5 watts per square inch of blade-shoe area would protect the propeller blades at the test conditions. The most satisfactory blade shoe of the three designs tested extended to the 20-percent-chord point and to 90 percent of the blade radius. A concentration of heat in the leading-edge region of this shoe was found to reduce the power input necessary for satisfactory de-icing. A satisfactory thermal design of blade shoe and a hub generator of sufficient capacity were developed.

PROPELLERS - HAMILTON STANDARD↗

DOE OSTI2023

Self‐Propelling Macroscale Sheets Powered by Enzyme Pumps

Abstract Nanoscale enzymes anchored to surfaces act as chemical pumps by converting chemical energy released from enzymatic reactions into spontaneous fluid flow that propels entrained nano‐ and microparticles. Enzymatic pumps are biocompatible, highly selective, and display unique substrate specificity. Utilizing these pumps to trigger self‐propelled motion on the macroscale has, however, constituted a significant challenge and thus prevented their adaptation in macroscopic fluidic devices and soft robotics. Using experiments and simulations, we herein show that enzymatic pumps can drive centimeter‐scale polymer sheets along directed linear paths and rotational trajectories. In these studies, the sheets are confined to the air/water interface. With the addition of appropriate substrate, the asymmetric enzymatic coating on the sheets induces chemically driven, buoyancy flows that controllably propel the sheet's motion on the air/water interface. The directionality and speed of the motion can be tailored by changing the pattern of the enzymatic coating, type of enzyme, and nature and concentration of the substrate. This work highlights the utility of biocompatible enzymes for generating motion in macroscale fluidic devices and robotics and indicates their potential utility for in vivo applications.

Song, Jiaqi↗

DOE OSTI2018

Agglomerate Sizing in Aluminized Propellants Using Digital Inline Holography and Traditional Diagnostics

Aluminized ammonium perchlorate composite propellants can form large molten agglomerated particles that may result in poor combustion performance, slag accumulation, and increased two-phase flow losses. Quantifying agglomerate size distributions are needed to gain an understanding of agglomeration dynamics and ultimately design new propellants for improved performance. Due to complexities of the reacting multiphase environment, agglomerate size diagnostics are difficult and measurement accuracies are poorly understood. To address this, the current paper compares three agglomerate sizing techniques applied to two propellant formulations. Particle collection on a quench plate and backlit videography are two relatively common techniques, whereas digital inline holography is an emerging alternative for three-dimensional measurements. Atmospheric pressure combustion results show that all three techniques are able to capture the qualitative trends; however, significant differences exist in the quantitative size distributions and mean diameters. For digital inline holography, methods are proposed that combine temporally resolved high-speed recording with lower-speed but higher spatial resolution measurements to correct for size–velocity correlation biases while extending the measurable size dynamic range. Finally, the results from this work provide new guidance for improved agglomerate size measurements along with statistically resolved datasets for validation of agglomerate models.

42 ENGINEERING↗

NASA NTRS1966

Descent capability of two-propeller tilt-wing configurations

The wing stall problem encountered with tilt-wing V/STOL designs during low-powered descent flight conditions has led to buffeting which adversely affects both performance and handling qualities. The results of tests conducted in the Langley full-scale wind tunnel with a large semispan model of a two-propeller tilt-wing configuration have indicated that three important factors provide substantial improvement in in the wing stall characteristics with consequent improvement in descent capability: down-at-center propeller rotation resulted in delayed inboard stalling and provided far better descent capability than the up-at-center rotation, moderate lowering of the propeller position relative to the wing chord provided further improvement in descent capability, and some flap deflection was absolutely essential in order to have any descent capability for low-powered flight conditions. Use of all three factors should provide good descent capability even without the complexity of other sophisticated stall control devices.

WING STALL↗

NASA NTRS1970

Volumetric calibration of a propellant utilization system

Method of calibrating capacitance-type propellant mass sensors provides meaningful mass units accurate to within one percent of the total load. Neither special propellant loading nor test firing is required. Recalibration of the system is possible without special propellant loading. This method provides increased flexibility for vehicle operation.

Lenning, J. J.↗

NASA NTRS1971

CSM programs SM RCS propellant quantity gaging systems program

Computer program calculates actual and useable remaining propellant quantities as required in positive expulsion rocket engine propellant feed system. Program establishes relationship between helium system pressures and temperatures and propellant weight remaining in tanks. Program is written in FORTRAN 4 for IBM-360 computer.

Cox, G. R.↗

NASA NTRS1971

Low acceleration solid propellant rocket ignition study

A study was conducted to develop a solid-propellant rocket igniter system that would build up thrust at a controlled rate of less than 0.2 G/sec. The system consisted of a long burning, regressive burning, controlled flow igniter and an inhibited progressive burning surface in the main rocket motor. The igniter performed the dual role of igniting, under vacuum backpressure and low L* (motor free volume/nozzle throat area ratio) conditions, the nonrestricted portion of the propellant and providing the mass addition necessary to sustain combustion until the propellant burning area had increased sufficiently to provide a stable motor-chamber pressure. Two series of tests were conducted with existing small test motor hardware to: (1) demonstrate the feasibility of the concept, (2) determine the important parameters governing the system, and (3) obtain design guidelines for future scaled-up motor tests. A quasi-steady-state mass balance for the ignition system was written and programmed for use as a motor design tool.

Strand, L. D.↗

NASA NTRS1971

Shuttle APS propellant thermal conditioner study

A study program was performed to allow selection of thermal conditioner assemblies for superheating O2 and H2 at supercritical pressures. The application was the auxiliary propulsion system (APS) for the space shuttle vehicle. The O2/H2 APS propellant feed system included propellant conditioners, of which the thermal conditioner assemblies were a part. Cryogens, pumped to pressures above critical, were directed to the thermal conditioner assembly included: (1) a gas generator assembly with ignition system and bipropellant valves, which burned superheated O2 and H2 at rich conditions; (2) a heat exchanger assembly for thermal conditioning of the cryogenic propellant; and (3) a dump nozzle for heat exchanger exhaust.

Pearson, W. E.↗

NASA NTRS1972

Propellant sidefeed-short pulse discharge thruster studies

The technique of feeding a solid propellant into the discharge from the sides of the discharge was evaluated. The thrust/power ratio could be significantly effected by the included angle of V-shaped propellants and by the electrode length. This result implies that when results are compared at the same specific impulse it is possible to obtain higher thrust efficiencies. In particular, it was found that for a given discharge energy the thrust/power ratio correlated with propellant mass. Increasing the integral simultaneously increases both the gas dynamic and electromagnetic thrust. An analytic expression was formulated for ablated mass which comprehensively describes experimental data in terms of geometry and electrical parameters. The correlation of the product impulse x specific impulse with discharge energy was also described. It is suggested that the reliability of dry energy storage capacitors does not equal the reliability of liquid impregnated units when the comparison is made at the same joules/Kg rating.

Palumbo, D. J.↗

NASA NTRS1972

Space storable propellant acquisition system

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

Tegart, J. R.↗

NASA NTRS1971

Effects of the acceleration vector on transient burning rate of an aluminized solid propellant.

Experimental results concerning the transient burning-rate augmentation of a 16% aluminum polybutadiene acrylic acid (PBAA) propellant burned in a 2-in. web motor at pressure levels from 300 to 1200 psia with centrifugal accelerations from 0 to 140 g. The orientation of the acceleration vector was varied to determine its effect on the transient burning rate. The burning-rate augmentation was strongly dependent on (1) acceleration level, (2) propellant distance burned (or burn time), and (3) orientation of the acceleration vector with respect to the burning surface. This transient rate augmentation resulted from the retention of molten metallic residue on the burning surface by the normal acceleration loading. The presence of the residue altered the combustion zone heat transfer and caused increased localized burning rates, as evidenced by the pitted propellant surfaces that were observed from extinction tests conducted at various acceleration levels.

Northam, G. B.↗