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Propulsion

Explore propulsion reports from NASA NTRS and DOE OSTI. Search rocket engines, electric propulsion and thrusters; verify designs and performance in the source documents.

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NASA NTRS2008

NASA's Evolutionary Xenon Thruster: The NEXT Ion Propulsion System for Solar System Exploration

This viewgraph presentation reviews NASA s Evolutionary Xenon Thruster (NEXT) Ion Propulsion system. The NEXT project is developing a solar electric ion propulsion system. The NEXT project is advancing the capability of ion propulsion to meet NASA robotic science mission needs. The NEXT system is planned to significantly improve performance over the state of the art electric propulsion systems, such as NASA Solar Electric Propulsion Technology Application Readiness (NSTAR). The status of NEXT development is reviewed, including information on the NEXT Thruster, the power processing unit, the propellant management system (PMS), the digital control interface unit, and the gimbal. Block diagrams NEXT system are presented. Also a review of the lessons learned from the Dawn and NSTAR systems is provided. In summary the NEXT project activities through 2007 have brought next-generation ion propulsion technology to a sufficient maturity level.

Pencil, Eric J.↗

NASA NTRS2009

Electric Propulsion Platforms at DFRC

NASA Dryden Flight Research Center is a world-class flight research facility located at Edwards AFB, CA. With access to a 44 sq. mile dry lakebed and 350 testable days per year, it is the ideal location for flight research. DFRC has been undertaking aircraft research for approximately six decades including the famous X-aircraft (X-1 through X-48) and many science and exploration platforms. As part of this impressive heritage, DFRC has garnered more hours of full-sized electric aircraft testing than any other facility in the US, and possibly the world. Throughout the 80 s and 90 s Dryden was the home of the Pathfinder, Pathfinder Plus, and Helios prototype solar-electric aircraft. As part of the ERAST program, these electric aircraft achieved a world record 97,000 feet altitude for propeller-driven aircraft. As a result of these programs, Dryden s staff has collected thousands of man-hours of electric aircraft research and testing. In order to better answer the needs of the US in providing aircraft technologies with lower fuel consumption, lower toxic emissions (NOx, CO, VOCs, etc.), lower greenhouse gas (GHG) emissions, and lower noise emissions, NASA has engaged in cross-discipline research under the Aeronautics Research Mission Directorate (ARMD). As a part of this overall effort, Mark Moore of LaRC has initiated a cross-NASA-center electric propulsion working group (EPWG) to focus on electric propulsion technologies as applied to aircraft. Electric propulsion technologies are ideally suited to overcome all of the obstacles mentioned above, and are at a sufficiently advanced state of development component-wise to warrant serious R&D and testing (TRL 3+). The EPWG includes participation from NASA Langley Research Center (LaRC), Glenn Research Center (GRC), Ames Research Center (ARC), and Dryden Flight Research Center (DFRC). Each of the center participants provides their own unique expertise to support the overall goal of advancing the state-of-the-art in aircraft electric propulsion technologies. DFRC will leverage its vast experience in flight test to assist in the integration and flight test phases of any electric propulsion program. DFRC s core competencies, that have particular relevance to the goals of the EPWG, include flight research planning and execution and providing aircraft test beds for researching and testing electric propulsion concepts and equipment. There are three flight regimes that the EPWG is focusing on: subsonic small GA and UAV, subsonic transport class, and supersonic. DFRC proposes two classes of test bed aircraft, to answer the early- and mid-phase testing requirements of all flight regimes the EPWG is concerned with. First, a highly efficient PIK motor glider will be used to test concepts and equipment associated with the subsonic GA and UAV aircraft regime (N+1). Second, a small fleet of subscale remotely-piloted aircraft test beds, similar to the X48B Blended Wing Body aircraft tested at Dryden, will be developed to answer the unique testing requirements of the subsonic GA and UAV, subsonic transport and possibly the supersonic class of aircraft (N+2, N+3). These aircraft can be tested in either serial stages or concurrent stages, depending on the actual test requirements and program schedules. Both classes of test bed aircraft are described below.

Baraaclough, Jonathan↗

NASA NTRS2008

Ion Beam Propulsion Study

The Ion Beam Propulsion Study was a joint high-level study between the Applied Physics Laboratory operated by NASA and ASRC Aerospace at Kennedy Space Center, Florida, and Berkeley Scientific, Berkeley, California. The results were promising and suggested that work should continue if future funding becomes available. The application of ion thrusters for spacecraft propulsion is limited to quite modest ion sources with similarly modest ion beam parameters because of the mass penalty associated with the ion source and its power supply system. Also, the ion source technology has not been able to provide very high-power ion beams. Small ion beam propulsion systems were used with considerable success. Ion propulsion systems brought into practice use an onboard ion source to form an energetic ion beam, typically Xe+ ions, as the propellant. Such systems were used for steering and correction of telecommunication satellites and as the main thruster for the Deep Space 1 demonstration mission. In recent years, "giant" ion sources were developed for the controlled-fusion research effort worldwide, with beam parameters many orders of magnitude greater than the tiny ones of conventional space thruster application. The advent of such huge ion beam sources and the need for advanced propulsion systems for exploration of the solar system suggest a fresh look at ion beam propulsion, now with the giant fusion sources in mind.

Source record↗

NASA NTRS2009

Propulsion Investigation for Zero and Near-Zero Emissions Aircraft

As world emissions are further scrutinized to identify areas for improvement, aviation s contribution to the problem can no longer be ignored. Previous studies for zero or near-zero emissions aircraft suggest aircraft and propulsion system sizes that would perform propulsion system and subsystems layout and propellant tankage analyses to verify the weight-scaling relationships. These efforts could be used to identify and guide subsequent work on systems and subsystems to achieve viable aircraft system emissions goals. Previous work quickly focused these efforts on propulsion systems for 70- and 100-passenger aircraft. Propulsion systems modeled included hydrogen-fueled gas turbines and fuel cells; some preliminary estimates combined these two systems. Hydrogen gas-turbine engines, with advanced combustor technology, could realize significant reductions in nitrogen emissions. Hydrogen fuel cell propulsion systems were further laid out, and more detailed analysis identified systems needed and weight goals for a viable overall system weight. Results show significant, necessary reductions in overall weight, predominantly on the fuel cell stack, and power management and distribution subsystems to achieve reasonable overall aircraft sizes and weights. Preliminary conceptual analyses for a combination of gas-turbine and fuel cell systems were also performed, and further studies were recommended. Using gas-turbine engines combined with fuel cell systems can reduce the fuel cell propulsion system weight, but at higher fuel usage than using the fuel cell only.

Snyder, Christopher A.↗

NASA NTRS2009

A3 Subscale Rocket Hot Fire Testing

This paper gives a description of the methodology and results of J2-X Subscale Simulator (JSS) hot fire testing supporting the A3 Subscale Diffuser Test (SDT) project at the E3 test facility at Stennis Space Center, MS (SSC). The A3 subscale diffuser is a geometrically accurate scale model of the A3 altitude simulating rocket test facility. This paper focuses on the methods used to operate the facility and obtain the data to support the aerodynamic verification of the A3 rocket diffuser design and experimental data quantifying the heat flux throughout the facility. The JSS was operated at both 80% and 100% power levels and at gimbal angle from 0 to 7 degrees to verify the simulated altitude produced by the rocket-rocket diffuser combination. This was done with various secondary GN purge loads to quantify the pumping performance of the rocket diffuser. Also, special tests were conducted to obtain detailed heat flux measurements in the rocket diffuser at various gimbal angles and in the facility elbow where the flow turns from vertical to horizontal upstream of the 2nd stage steam ejector.

Saunders, G. P.↗

NASA NTRS2009

Implementation of an Online Database for Chemical Propulsion Systems

The Johns Hopkins University, Chemical Propulsion Information Analysis Center (CPIAC) has been working closely with NASA Goddard Space Flight Center (GSFC); NASA Marshall Space Flight Center (MSFC); the University of Alabama at Huntsville (UAH); The Johns Hopkins University, Applied Physics Laboratory (APL); and NASA Jet Propulsion Laboratory (JPL) to capture satellite and spacecraft propulsion system information for an online database tool. The Spacecraft Chemical Propulsion Database (SCPD) is a new online central repository containing general and detailed system and component information on a variety of spacecraft propulsion systems. This paper only uses data that have been approved for public release with unlimited distribution. The data, supporting documentation, and ability to produce reports on demand, enable a researcher using SCPD to compare spacecraft easily, generate information for trade studies and mass estimates, and learn from the experiences of others through what has already been done. This paper outlines the layout and advantages of SCPD, including a simple example application with a few chemical propulsion systems from various NASA spacecraft.

David B. Owen, II↗

NASA NTRS2010

Nitrous Oxide/Paraffin Hybrid Rocket Engines

Nitrous oxide/paraffin (N2OP) hybrid rocket engines have been invented as alternatives to other rocket engines especially those that burn granular, rubbery solid fuels consisting largely of hydroxyl- terminated polybutadiene (HTPB). Originally intended for use in launching spacecraft, these engines would also be suitable for terrestrial use in rocket-assisted takeoff of small airplanes. The main novel features of these engines are (1) the use of reinforced paraffin as the fuel and (2) the use of nitrous oxide as the oxidizer. Hybrid (solid-fuel/fluid-oxidizer) rocket engines offer advantages of safety and simplicity over fluid-bipropellant (fluid-fuel/fluid-oxidizer) rocket en - gines, but the thrusts of HTPB-based hybrid rocket engines are limited by the low regression rates of the fuel grains. Paraffin used as a solid fuel has a regression rate about 4 times that of HTPB, but pure paraffin fuel grains soften when heated; hence, paraffin fuel grains can, potentially, slump during firing. In a hybrid engine of the present type, the paraffin is molded into a 3-volume-percent graphite sponge or similar carbon matrix, which supports the paraffin against slumping during firing. In addition, because the carbon matrix material burns along with the paraffin, engine performance is not appreciably degraded by use of the matrix.

Zubrin, Robert↗

NASA NTRS2007

Rocket Science 101 Interactive Educational Program

To better educate the public on the basic design of NASA s current mission rockets, Rocket Science 101 software has been developed as an interactive program designed to retain a user s attention and to teach about basic rocket parts. This program also has helped to expand NASA's presence on the Web regarding educating the public about the Agency s goals and accomplishments. The software was designed using Macromedia s Flash 8. It allows the user to select which type of rocket they want to learn about, interact with the basic parts, assemble the parts to create the whole rocket, and then review the basic flight profile of the rocket they have built.

Armstrong, Dennis↗

NASA NTRS2010

Propulsion System Choices and Their Implications

In defining a space vehicle architecture, the propulsion system and related subsystem choices will have a major influence on achieving the goals and objectives desired. There are many alternatives and the choices made must produce a system that meets the performance requirements, but at the same time also provide the greatest opportunity of reaching all of the required objectives. Recognizing the above, the SPST Functional Requirements subteam has drawn on the knowledge, expertise, and experience of its members, to develop insight that wiIJ effectively aid the architectural concept developer in making the appropriate choices consistent with the architecture goals. This data not only identifies many selected choices, but also, more importantly, presents the collective assessment of this subteam on the "pros" and the "cons" of these choices. The propulsion system choices with their pros and cons are presented in five major groups. A. System Integration Approach. Focused on the requirement for safety, reliability, dependability, maintainability, and low cost. B. Non-Chemical Propulsion. Focused on choice of propulsion type. C. Chemical Propulsion. Focused on propellant choice implications. D. Functional Integration. Focused on the degree of integration of the many propulsive and closely associated functions, and on the choice of the engine combustion power cycle. E. Thermal Management. Focused on propellant tank insulation and integration. Each of these groups is further broken down into subgroups, and at that level the consensus pros and cons are presented. The intended use of this paper is to provide a resource of focused material for architectural concept developers to use in designing new advanced systems including college design classes. It is also a possible source of input material for developing a model for designing and analyzing advanced concepts to help identify focused technology needs and their priorities.

Joyner, Claude R., II↗

NASA NTRS2011

Turboelectric Distributed Propulsion in a Hybrid Wing Body Aircraft

The performance of the N3-X, a 300 passenger hybrid wing body (HWB) aircraft with turboelectric distributed propulsion (TeDP), has been analyzed to see if it can meet the 70% fuel burn reduction goal of the NASA Subsonic Fixed Wing project for N+3 generation aircraft. The TeDP system utilizes superconducting electric generators, motors and transmission lines to allow the power producing and thrust producing portions of the system to be widely separated. It also allows a small number of large turboshaft engines to drive any number of propulsors. On the N3-X these new degrees of freedom were used to (1) place two large turboshaft engines driving generators in freestream conditions to maximize thermal efficiency and (2) to embed a broad continuous array of 15 motor driven propulsors on the upper surface of the aircraft near the trailing edge. That location maximizes the amount of the boundary layer ingested and thus maximizes propulsive efficiency. The Boeing B777-200LR flying 7500 nm (13890 km) with a cruise speed of Mach 0.84 and an 118100 lb payload was selected as the reference aircraft and mission for this study. In order to distinguish between improvements due to technology and aircraft configuration changes from those due to the propulsion configuration changes, an intermediate configuration was included in this study. In this configuration a pylon mounted, ultra high bypass (UHB) geared turbofan engine with identical propulsion technology was integrated into the same hybrid wing body airframe. That aircraft achieved a 52% reduction in mission fuel burn relative to the reference aircraft. The N3-X was able to achieve a reduction of 70% and 72% (depending on the cooling system) relative to the reference aircraft. The additional 18% - 20% reduction in the mission fuel burn can therefore be attributed to the additional degrees of freedom in the propulsion system configuration afforded by the TeDP system that eliminates nacelle and pylon drag, maximizes boundary layer ingestion (BLI) to reduce inlet drag on the propulsion system, and reduces the wake drag of the vehicle.

Felder, James L.↗

NASA NTRS2011

Waterhammer Transient Simulation and Model Anchoring for the Robotic Lunar Lander Propulsion System

Waterhammer transients have the potential to adversely impact propulsion system design if not properly addressed. Waterhammer can potentially lead to system plumbing, and component damage. Multi-thruster propulsion systems also develop constructive/destructive wave interference which becomes difficult to predict without detailed models. Therefore, it is important to sufficiently characterize propulsion system waterhammer in order to develop a robust design with minimal impact to other systems. A risk reduction activity was performed at Marshall Space Flight Center to develop a tool for estimating waterhammer through the use of anchored simulation for the Robotic Lunar Lander (RLL) propulsion system design. Testing was performed to simulate waterhammer surges due to rapid valve closure and consisted of twenty-two series of waterhammer tests, resulting in more than 300 valve actuations. These tests were performed using different valve actuation schemes and three system pressures. Data from the valve characterization tests were used to anchor the models that employed MSCSoftware.EASY5 v.2010 to model transient fluid phenomena by using transient forms of mass and energy conservation. The anchoring process was performed by comparing initial model results to experimental data and then iterating the model input to match the simulation results with the experimental data. The models provide good correlation with experimental results, supporting the use of EASY5 as a tool to model fluid transients and provide a baseline for future RLL system modeling. This paper addresses tasks performed during the waterhammer risk reduction activity for the RLL propulsion system. The problem of waterhammer simulation anchoring as applied to the RLL system is discussed with results from the corresponding experimental valve tests. Important factors for waterhammer mitigation are discussed along with potential design impacts to the RLL propulsion system.

Stein, William B.↗

NASA NTRS2011

Catalytic Microtube Rocket Igniter

Devices that generate both high energy and high temperature are required to ignite reliably the propellant mixtures in combustion chambers like those present in rockets and other combustion systems. This catalytic microtube rocket igniter generates these conditions with a small, catalysis-based torch. While traditional spark plug systems can require anywhere from 50 W to multiple kW of power in different applications, this system has demonstrated ignition at less than 25 W. Reactants are fed to the igniter from the same tanks that feed the reactants to the rest of the rocket or combustion system. While this specific igniter was originally designed for liquid methane and liquid oxygen rockets, it can be easily operated with gaseous propellants or modified for hydrogen use in commercial combustion devices. For the present cryogenic propellant rocket case, the main propellant tanks liquid oxygen and liquid methane, respectively are regulated and split into different systems for the individual stages of the rocket and igniter. As the catalyst requires a gas phase for reaction, either the stored boil-off of the tanks can be used directly or one stream each of fuel and oxidizer can go through a heat exchanger/vaporizer that turns the liquid propellants into a gaseous form. For commercial applications, where the reactants are stored as gases, the system is simplified. The resulting gas-phase streams of fuel and oxidizer are then further divided for the individual components of the igniter. One stream each of the fuel and oxidizer is introduced to a mixing bottle/apparatus where they are mixed to a fuel-rich composition with an O/F mass-based mixture ratio of under 1.0. This premixed flow then feeds into the catalytic microtube device. The total flow is on the order of 0.01 g/s. The microtube device is composed of a pair of sub-millimeter diameter platinum tubes connected only at the outlet so that the two outlet flows are parallel to each other. The tubes are each approximately 10 cm long and are heated via direct electric resistive heating. This heating brings the gasses to their minimum required ignition temperature, which is lower than the auto-thermal ignition temperature, and causes the onset of both surface and gas phase ignition producing hot temperatures and a highly reacting flame. The combustion products from the catalytic tubes, which are below the melting point of platinum, are injected into the center of another combustion stage, called the primary augmenter. The reactants for this combustion stage come from the same source but the flows of non-premixed methane and oxygen gas are split off to a secondary mixing apparatus and can be mixed in a near-stoichiometric to highly lean mixture ratio. The primary augmenter is a component that has channels venting this mixed gas to impinge on each other in the center of the augmenter, perpendicular to the flow from the catalyst. The total crosssectional area of these channels is on a similar order as that of the catalyst. The augmenter has internal channels that act as a manifold to distribute equally the gas to the inward-venting channels. This stage creates a stable flame kernel as its flows, which are on the order of 0.01 g/s, are ignited by the combustion products of the catalyst. This stage is designed to produce combustion products in the flame kernel that exceed the autothermal ignition temperature of oxygen and methane.

Schneider, Steven J.↗

NASA NTRS2012

Roadmap for In-Space Propulsion Technology

NASA has created a roadmap for the development of advanced in-space propulsion technologies for the NASA Office of the Chief Technologist (OCT). This roadmap was drafted by a team of subject matter experts from within the Agency and then independently evaluated, integrated and prioritized by a National Research Council (NRC) panel. The roadmap describes a portfolio of in-space propulsion technologies that could meet future space science and exploration needs, and shows their traceability to potential future missions. Mission applications range from small satellites and robotic deep space exploration to space stations and human missions to Mars. Development of technologies within the area of in-space propulsion will result in technical solutions with improvements in thrust, specific impulse (Isp), power, specific mass (or specific power), volume, system mass, system complexity, operational complexity, commonality with other spacecraft systems, manufacturability, durability, and of course, cost. These types of improvements will yield decreased transit times, increased payload mass, safer spacecraft, and decreased costs. In some instances, development of technologies within this area will result in mission-enabling breakthroughs that will revolutionize space exploration. There is no single propulsion technology that will benefit all missions or mission types. The requirements for in-space propulsion vary widely according to their intended application. This paper provides an updated summary of the In-Space Propulsion Systems technology area roadmap incorporating the recommendations of the NRC.

Meyer, Michael↗

NASA NTRS2012

Mars Sample Return Using Solar Sail Propulsion

Many Mars Sample Return (MSR) architecture studies have been conducted over the years. A key element of them is the Earth Return Stage (ERS) whose objective is to obtain the sample from the Mars Ascent Vehicle (MAV) and return it safely to the surface of the Earth. ERS designs predominantly use chemical propulsion [1], incurring a significant launch mass penalty due to the low specific impulse of such systems coupled with the launch mass sensitivity to returned mass. It is proposed to use solar sail propulsion for the ERS, providing a high (effective) specific impulse propulsion system in the final stage of the multi-stage system. By doing so to the launch mass of the orbiter mission can be significantly reduced and hence potentially decreasing mission cost. Further, solar sailing offers a unique set of non-Keplerian low thrust trajectories that may enable modifications to the current approach to designing the Earth Entry Vehicle by potentially reducing the Earth arrival velocity. This modification will further decrease the mass of the orbiter system. Solar sail propulsion uses sunlight to propel vehicles through space by reflecting solar photons from a large, mirror-like surface made of a lightweight, reflective material. The continuous photonic pressure provides propellantless thrust to conduct orbital maneuvering and plane changes more efficiently than conventional chemical propulsion. Because the Sun supplies the necessary propulsive energy, solar sails require no onboard propellant, thus reducing system mass. This technology is currently at TRL 7/8 as demonstrated by the 2010 flight of the Japanese Aerospace Exploration Agency, JAXA, IKAROS mission. [2]

Johnson, Les↗

NASA NTRS2013

Assessment of tbe Performance of Ablative Insulators Under Realistic Solid Rocket Motor Operating Conditions (a Doctoral Dissertation)

Ablative insulators are used in the interior surfaces of solid rocket motors to prevent the mechanical structure of the rocket from failing due to intense heating by the high-temperature solid-propellant combustion products. The complexity of the ablation process underscores the need for ablative material response data procured from a realistic solid rocket motor environment, where all of the potential contributions to material degradation are present and in their appropriate proportions. For this purpose, the present study examines ablative material behavior in a laboratory-scale solid rocket motor. The test apparatus includes a planar, two-dimensional flow channel in which flat ablative material samples are installed downstream of an aluminized solid propellant grain and imaged via real-time X-ray radiography. In this way, the in-situ transient thermal response of an ablator to all of the thermal, chemical, and mechanical erosion mechanisms present in a solid rocket environment can be observed and recorded. The ablative material is instrumented with multiple micro-thermocouples, so that in-depth temperature histories are known. Both total heat flux and thermal radiation flux gauges have been designed, fabricated, and tested to characterize the thermal environment to which the ablative material samples are exposed. These tests not only allow different ablative materials to be compared in a realistic solid rocket motor environment but also improve the understanding of the mechanisms that influence the erosion behavior of a given ablative material.

Martin, Heath Thomas↗

NASA NTRS2013

Benefits of Hybrid-Electric Propulsion to Achieve 4x Increase in Cruise Efficiency for a VTOL Aircraft

Electric propulsion enables radical new vehicle concepts, particularly for Vertical Takeoff and Landing (VTOL) aircraft because of their significant mismatch between takeoff and cruise power conditions. However, electric propulsion does not merely provide the ability to normalize the power required across the phases of flight, in the way that automobiles also use hybrid electric technologies. The ability to distribute the thrust across the airframe, without mechanical complexity and with a scale-free propulsion system, is a new degree of freedom for aircraft designers. Electric propulsion is scale-free in terms of being able to achieve highly similar levels of motor power to weight and efficiency across a dramatic scaling range. Applying these combined principles of electric propulsion across a VTOL aircraft permits an improvement in aerodynamic efficiency that is approximately four times the state of the art of conventional helicopter configurations. Helicopters typically achieve a lift to drag ratio (L/D) of between 4 and 5, while the VTOL aircraft designed and developed in this research were designed to achieve an L/D of approximately 20. Fundamentally, the ability to eliminate the problem of advancing and retreating rotor blades is shown, without resorting to unacceptable prior solutions such as tail-sitters. This combination of concept and technology also enables a four times increase in range and endurance while maintaining the full VTOL and hover capability provided by a helicopter. Also important is the ability to achieve low disc-loading for low ground impingement velocities, low noise and hover power minimization (thus reducing energy consumption in VTOL phases). This combination of low noise and electric propulsion (i.e. zero emissions) will produce a much more community-friendly class of vehicles. This research provides a review of the concept brainstorming, configuration aerodynamic and mission analysis, as well as subscale prototype construction and flight testing that verifies transition flight control. A final down-selected vehicle is also presented.

Fredericks, William J.↗

NASA NTRS2013

Revolutionary Aeropropulsion Concept for Sustainable Aviation: Turboelectric Distributed Propulsion

In response to growing aviation demands and concerns about the environment and energy usage, a team at NASA proposed and examined a revolutionary aeropropulsion concept, a turboelectric distributed propulsion system, which employs multiple electric motor-driven propulsors that are distributed on a large transport vehicle. The power to drive these electric propulsors is generated by separately located gas-turbine-driven electric generators on the airframe. This arrangement enables the use of many small-distributed propulsors, allowing a very high effective bypass ratio, while retaining the superior efficiency of large core engines, which are physically separated but connected to the propulsors through electric power lines. Because of the physical separation of propulsors from power generating devices, a new class of vehicles with unprecedented performance employing such revolutionary propulsion system is possible in vehicle design. One such vehicle currently being investigated by NASA is called the "N3-X" that uses a hybrid-wing-body for an airframe and superconducting generators, motors, and transmission lines for its propulsion system. On the N3-X these new degrees of design freedom are used (1) to place two large turboshaft engines driving generators in freestream conditions to minimize total pressure losses and (2) to embed a broad continuous array of 14 motor-driven fans on the upper surface of the aircraft near the trailing edge of the hybrid-wing-body airframe to maximize propulsive efficiency by ingesting thick airframe boundary layer flow. Through a system analysis in engine cycle and weight estimation, it was determined that the N3-X would be able to achieve a reduction of 70% or 72% (depending on the cooling system) in energy usage relative to the reference aircraft, a Boeing 777-200LR. Since the high-power electric system is used in its propulsion system, a study of the electric power distribution system was performed to identify critical dynamic and safety issues. This paper presents some of the features and issues associated with the turboelectric distributed propulsion system and summarizes the recent study results, including the high electric power distribution, in the analysis of the N3-X vehicle.

Kim, Hyun Dae↗

NASA NTRS2014

RS-34 Phoenix In-Space Propulsion System Applied to Active Debris Removal Mission

In-space propulsion is a high percentage of the cost when considering Active Debris Removal mission. For this reason it is desired to research if existing designs with slight modification would meet mission requirements to aid in reducing cost of the overall mission. Such a system capable of rendezvous, close proximity operations, and de-orbit of Envisat class resident space objects has been identified in the existing RS-34 Phoenix. RS-34 propulsion system is a remaining asset from the de-commissioned United States Air Force Peacekeeper program; specifically the pressure-fed storable bi-propellant Stage IV Post Boost Propulsion System. The National Aeronautics and Space Administration (NASA) Marshall Space Flight Center (MSFC) gained experience with the RS-34 propulsion system on the successful Ares I-X flight test program flown in the Ares I-X Roll control system (RoCS). The heritage hardware proved extremely robust and reliable and sparked interest for further utilization on other potential in-space applications. Subsequently, MSFC has obtained permission from the USAF to obtain all the remaining RS-34 stages for re-use opportunities. The MSFC Advanced Concepts Office (ACO) was commissioned to lead a study for evaluation of the Rocketdyne produced RS-34 propulsion system as it applies to an active debris removal design reference mission for resident space object targets including Envisat. Originally designed, the RS-34 Phoenix provided in-space six-degrees-of freedom operational maneuvering to deploy payloads at multiple orbital locations. The RS-34 Concept Study lead by sought to further understand application for a similar orbital debris design reference mission to provide propulsive capability for rendezvous, close proximity operations to support the capture phase of the mission, and deorbit of single or multiple large class resident space objects. Multiple configurations varying the degree of modification were identified to trade for dry mass optimization and propellant load. The results of the RS-34 Phoenix Concept Study show that the system is technically sufficient to successfully support all of the missions to rendezvous, capture, and de-orbit targets including Envisat and Hubble Space Telescope. The results and benefits of the RS-34 Orbital Debris Application Concept Study are presented in this paper.

Esther, Elizabeth A.↗