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NASA NTRS2011

Space Transportation Infrastructure Supported By Propellant Depots

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.

Smitherman, David↗

NASA NTRS2009

Water Contaminant Mitigation in Ionic Liquid Propellant

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.

Conroy, David↗

NASA NTRS2015

Green Propellant Demonstration with Hydrazine Catalyst of F-16 Emergency Power Unit

Some space vehicle and aircraft Auxiliary Power Units (APUs) use hydrazine propellant for generating power. Hydrazine is a toxic, hazardous fuel which requires special safety equipment and processes for handling and loading. In recent years, there has been development of two green propellants that could enable their use in APU's: the Swedish LMP-103S and the Air Force Research Laboratory (AFRL) AF-M315E. While there has been work on development of these propellants for thruster applications (Prisma and Green Propulsion Infusion Mission, respectively), there has been less focus on the application to power units. Beginning in 2012, an effort was started by the Marshall Space Flight Center (MSFC) on the APU application. The MSFC plan was to demonstrate green propellants with residual Space Shuttle hardware. The principal investigator was able to acquire a Solid Rocket Booster gas generator and an Orbiter APU. Since these test assets were limited in number, an Air Force equivalent asset was identified: the F-16 Emergency Power Unit (EPU). In June 2013, two EPU's were acquired from retired aircraft located at Davis Monthan Air Force Base. A gas generator from one of these EPU's was taken out of an assembly and configured for testing with a version of the USAF propellant with a higher water content (AF-M315EM) to reduce decomposition temperatures. Testing in November 2014 has shown that this green propellant is reactive with the Hydrazine catalyst (Shell 405) generating 300 psi of pressure with the existing F-16 EPU configuration. This paper will highlight the results of MSFC testing in collaboration with AFRL.

Robinson, Joel W.↗

NASA NTRS2016

A Simple Method for High-Lift Propeller Conceptual Design

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.

Patterson, Michael↗

NASA NTRS2018

Low-Noise Operating Mode for Propeller-Driven Electric Airplanes

Mechanical shaft power and shaft speed of reciprocating internal combustion engines are closely coupled. Maximum rated shaft power is typically produced at or near peak shaft speed. If a general aviation airplane equipped with a reciprocating engine and a variable-pitch propeller attempts a low-noise takeoff by reducing propeller tip speed, propeller power and thrust are reduced. Such takeoffs are not tolerated due to punishing performance effects, such as increased field lengths and poor climb rates. Certain electric motors, however, are able to deliver maximum shaft power over a wide range of shaft speed. Electric or hybrid-electric propeller-driven airplanes should be able to exploit this behavior. At low shaft speeds, high shaft power levels and high blade pitch angles could be combined to recover much of the thrust that would otherwise be lost. This could enable a low-noise operating mode for propellers normally designed for performance rather than for noise. The subject of this paper is an analytical investigation into low-noise takeoffs and steady overflights of a notional general aviation airplane equipped with a propeller driven by an electric motor.

Berton, Jeffrey J.↗

NASA NTRS2018

Exploring the Effects of Installation Geometry in High-Lift Propeller Systems

A high-lift propeller system is a distributed electric propulsion technology which dedicates an array of wing-mounted tractor propellers to actively augment wing lift during takeoff and landing. This paper describes the results of a wind tunnel experiment dedicated to investigating the effects of high-lift propeller installation geometry on lift generation. Variables investigated include propeller height, offset, and inclination. Results show that propeller height is the most critical variable and that the height for maximum lift depends highly on the angle of attack and flap deflection. In addition, a relationship between optimal propeller height and the wing’s unblown lift coefficient is discovered.

Fei, Xiaofan↗

NASA NTRS2020

Controller Design for Propeller Phase Synchronization with Aeroacoustic Performance Metrics

Active noise reduction using phase control takes advantage of the propellers used in a distributed electric air vehicle by treating each propeller as an independent acoustic source. These acoustic sources can destructively interfere if the propellers are synchronized. With this method, the radiated sound power around the vehicle can be reduced. The purpose of this paper is to design a controller that regulates propeller positions to reduce the radiated sound power through destructive interference and to demonstrate the performance of the controller through acoustic testing. There are two control requirements for reducing the sound pressure level considered in this paper. The first is accurately regulating the difference between the propeller azimuthal blade positions (phase) relative to their neighbors. By changing the phase at the source, we can control the phase of the sound wave at an arrival position to create destructive interference. The second consideration is maintaining high coherence between the two propeller sources. To achieve significant attenuation, the controller needs to regulate phase error and suppress sources of incoherence. Performance is demonstrated through sound recordings performed in the NASA Langley Structural Acoustic Loads and Transmission (SALT) anechoic chamber. We show that the controller is capable of reducing sound pressure level at a given observer location by 17 dB at the blade passage frequency and that this method can reduce the radiated sound power by 6 dB at the blade passage frequency.

Andrew Patterson↗

NASA NTRSDate not supplied

Aerodynamic and Acoustic Interactions Associated with Inboard Propeller-Wing Configurations

A series of aerodynamic performance and acoustic measurements have been made on a range of inboard propeller-wing interaction configurations in the NASA Langley Low Speed Aeroacoustic Wind Tunnel (LSAWT). The results presented in this paper are part of a more expansive testing campaign encompassing both single propeller-wing and multipropeller-wing interactions, the former of which is discussed in the present work. The primary testing parameters of interest to this study are the axial and vertical positioning of the wing relative to the propeller slipstream under a constant propeller advance ratio. A multi-faceted computational effort was also employed in an effort to identify reflection and scattering effects imposed by both the wing geometry as well as the primary components of the facility test setup. This effort consisted of aerodynamic predictions using high-fidelity computational fluid dynamics (CFD), acoustic predictions using an impermeable Ffowcs Williams and Hawkings (FW-H) solver, and acoustic scattering predictions. Acoustic measurements reveal variations in the acoustic directivity behavior of the propeller blade passage frequency for even modest variations in wing position. CFD-based acoustic predictions reveal discrepancies relative to the experimental data, which is believed to be due to complex acoustic scattering behavior within the test section. Initial attempts at modeling the scattered acoustic field showed functional dependency of the acoustic amplitude variations on the wing position relative to the propeller disk, however discrepancies with experimental data remain.

Nikolas S. Zawodny↗

NASA NTRSDate not supplied

Concept of Operations and Checkout Testing for Green Propellant Loading System

Low toxicity propulsion, or Green Propulsion, is increasingly being considered for use on spacecraft following the successful demonstration during the Green Propellant Infusion Mission in 2020. One of these missions is JPL’s Lunar Flashlight (LF), a polar orbiting 12U CubeSat that uses lasers to investigate water ice deposits on the moon’s surface. For the past 3 years, Marshall Space Flight Center has developed the ASCENT propulsion system for LF, along with a green propellant loading system to service the spacecraft. The Propellant Loading System (PLS) is composed of 3 mobile carts designed to operate at a variety of launch facilities and handle the 2 primary green propellants, LMP-103S and ASCENT. In anticipation of LF propellant loading, the PLS was recently installed into an MSFC facility for checkout operations. In the winter of 2020, a simulant spacecraft was successfully loaded with water, pressurized to flight pressure, and drained back into the system. These checkouts drove the concept of operations for final spacecraft loading and informed minor modifications to the PLS for better operability. Additionally, a parallel effort was started to design the next generation of the PLS based on water flow lessons learned. With water flow testing complete, propellant checkouts and wet dress rehearsals with spacecraft pathfinder hardware are anticipated to begin in late Spring 2021.

Green Propulsion↗

NASA NTRSDate not supplied

Nodal Numerical Modeling of Submerged Helium Injection in a Cryogenic Propellant Tank

Subcooling of cryogenic propellant by helium injection is one of the most effective methods for suppressing bulk boiling and keeping subcooled propellant conditions for pre-launch, launch, and post-launch pressurization applications. For tank pressurization, submerged helium injection can substantially reduce helium consumption by infusing gaseous propellant into the tank ullage. This paper presents a mathematical model of the helium bubbling process in liquid oxygen to estimate the amount of oxygen vapor absorbed by the rising helium bubbles and the amount of subcooling of liquid oxygen due to evaporative heat and mass transfer. This mathematical model was incorporated in a simulation model of tank pressurization built with Generalized Fluid System Simulation Program (GFSSP), a general-purpose flow network code developed at NASA/Marshall Space Flight Center. The numerical predictions of subcooling have been compared with the experimental data of Cho et al. which investigated the propellant subcooling effect as a function of system pressure, helium injection temperature, and flowrate for a non-drained submerged injection system. The numerical predictions of helium consumption have been compared with the test data from a NASA Centaur test vehicle which included both direct and submerged injection with draining of propellants. The numerical model developed with GFSSP has been validated against two sets of experimental data and has been shown to predict both propellant subcooling and helium consumption to within 30% in most cases. The test data used for the model validation were taken in 1-g, but the mass diffusion model was developed to be applied in both 1-g and micro-g environments.

Cryogenics↗

NASA NTRSDate not supplied

Slosh-Induced Ullage Collapse Validation by Computational Fluid Dynamics and Application to In-space Propellant Dynamics

Ullage collapse predictions made using Computational Fluid Dynamics (CFD) were validated then made for an in-space cryogenic propellant unsettling application. Ullage collapse is the decrease in propellant tank pressure due to cooling of the ullage gas, which can be significant for cryogenic propulsion systems. Vehicles may be developed that are robust to temperature and pressure changes caused by heat leaks, e.g., propellant tanks with high maximum operation pressures, technologies yielding active thermal conditioning, and systems with ample pressurant for conditioning. These methods are robust because additional mass, cost, and complexity are used to address cryogenic fluid management (CFM). As vehicles are optimized for increased payload capacity, tracking of events such as ullage collapse and self-pressurization in tanks will become increasingly important for propellant inventory calculations. Some ullage collapse events can be studied through ground experimentation, but propellant unsettling in and around in-space coast mission phases must be studied in the absence of an acceleration field. Computational Fluid Dynamics is a powerful tool to address this need, enabling prediction of complex physics prior to flight. Validation of the chosen CFD tool and modeling methodology was done using data from a previously conducted subscale slosh-induced ullage collapse ground test performed in part by the Japan Aerospace Exploration Agency (JAXA). A small tank was partially filled with liquid nitrogen and pressurized with gaseous nitrogen at various temperatures. The tank was oscillated laterally to induce high amplitude slosh dynamics resulting in wave breakup. Successful qualitative validation of slosh wave dynamics and quantitative validation of ullage collapse promoted confidence in extending ullage collapse predictions to a flight application. Mechanisms of ullage collapse caused by main engine cutoff while in-space were identified. Both slosh dynamics leading to engine shutdown and body dynamics following shutdown were shown to contribute to propellant unsettling and ullage collapse.

J M Brodnick↗

NASA NTRSDate not supplied

Analysis of Cryogenic Propellant Liquefaction Rates in Cooled Constant-Wall-Temperature Tanks

NASA has been focused on developing technology that would allow the production of cryogenic propellants on the Lunar and Martian surfaces. Utilizing Lunar/Martian resources, the produced gaseous propellants must first be liquefied and stored prior to use on the Moon or Mars ascent vehicle. Liquefaction of cryogenic propellant is a necessary technical development to enable NASA’s future spaceflight goals. This paper presents an overview of the effect of spacecraft propellant system parameters on cryogenic propellant liquefaction rates. The propellant system was assumed to be a receiver of a gas from an In-Situ Resource Utilization (ISRU) harvester and condensed the gas by cooling the tank walls. First, the general thermodynamic principles pertinent to condensation are summarized, followed by an overview of the model which performed the analysis, including a summary of the Fortran implementation. Comparisons between the model results and test data are discussed, as well as sensitivity of the condensation rate to various parameters.

Anson Koch↗

NASA NTRSDate not supplied

Analysis of Cryogenic Propellant Liquefaction Rates in Cooled Constant-Wall-Temperature Tanks

NASA has been focused on developing technology that would allow the production of cryogenic propellants on the Lunar and Martian surfaces. Utilizing Lunar/Martian resources, the produced gaseous propellants must first be liquefied and stored prior to use on the Moon or Mars ascent vehicle. Liquefaction of cryogenic propellants is a necessary technical development to enable NASA’s future spaceflight goals. This paper presents an overview of a proposed model for a propellant liquefaction system, and the effect of tank wall temperatures on cryogenic propellant liquefaction rates. The propellant system was assumed to be a receiver of a gas from an In-Situ Resource Utilization (ISRU) harvester and condensed the gas by cooling the tank walls. First, the heat transfer principles pertinent to condensation are summarized, followed by an overview of the model which performed the analysis, including a summary of the Fortran algorithm implementation. The sensitivity of the condensation rate relative to varying tank wall temperatures is also discussed.

Anson Koch↗

NASA NTRSDate not supplied

Analysis of Cryogenic Propellant Liquefaction Rates in Cooled Constant-Wall-TemperatureTanks

NASA has been focused on developing technology that would allow the production of cryogenic propellants on the Lunar and Martian surfaces. Utilizing Lunar/Martian resources, the produced gaseous propellants must first be liquefied and stored prior to use on the Moon or Mars ascent vehicle. Liquefaction of cryogenic propellants is a necessary technical development to enable NASA’s future spaceflight goals. This paper presents an overview of a proposed model for a propellant liquefaction system, and the effect of tank wall temperature and the ullage pressure control band on cryogenic propellant liquefaction rates. The propellant system was assumed to be a receiver of a gas from an In-Situ Resource Utilization (ISRU) harvester and condensed the gas by cooling the tank walls. First, the heat transfer principles pertinent to condensation are summarized, followed by an overview of the model which performed the analysis, including a summary of the Fortran algorithm implementation. The sensitivity of the condensation rate relative to varying tank wall temperatures is also discussed.

cryogenics↗

NASA NTRSDate not supplied

Analysis of Cryogenic Propellant Liquefaction Rates in Cooled Constant-Wall-Temperature Tanks

cryogenics↗

NASA NTRS2023

Analysis of Cryogenic Propellant Liquefaction Rates in Cooled Constant-Wall-TemperatureTanks

cryogenics↗

NASA NTRSDate not supplied

Cryogenic Propellant Tank Drain Simulation with a Stratified Ullage Nodal Model

Cryogenic propellant tank pressurization systems maintain tank pressure within specified limits to comply with propellant thermodynamic and tank structural design requirements. Many pump-fed cryogenic liquid propulsion systems control tank pressure using an autogenous pressurization system. Autogenous pressurization systems vaporize and heat propellant to produce the pressurant gas delivered to the tank. While autogenous pressurization systems eliminate the need for a separate pressurant gas storage and delivery system, they couple pressurant mass requirements to propellant mass requirements. Accurate prediction of autogenous pressurant mass during propellant drain is essential in optimizing cryogenic propellant tank designs.

Richard G Hibbs↗

NASA NTRS1985

Aerodynamics and Acoustics of Propellers

A state of the art review of the aerodynamic design and testing of modern propellers, acoustic and vibration environmental problems and their solutions, and considerations in the integration of the propeller(s) and airframe is presented. Propellers have recently become a focus of attention after having been neglected for many years. The efficient use of propellers as a propulsion medium up to Mach 0.6 was realized at the expense of high noise and vibration. Since the mid-1970s there have been many developments, including the propfan, with an aerodynamic efficiency of 80% at Mach 0.8 now seriously challenges the fanjet. The advent of the supercritical airfoil is another significant development beginning to influence propeller design.

Design↗