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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

Lunar Surface Access Module Descent Engine Turbopump Technology: Detailed Design

The need for a high specific impulse LOX/LH2 pump-fed lunar lander engine has been established by NASA for the new lunar exploration architecture. Studies indicate that a 4-engine cluster in the thrust range of 9,000-lbf each is a candidate configuration for the main propulsion of the manned lunar lander vehicle. The lander descent engine will be required to perform minor mid-course corrections, a Lunar Orbit Insertion (LOI) burn, a de-orbit burn, and the powered descent onto the lunar surface. In order to achieve the wide range of thrust required, the engines must be capable of throttling approximately 10:1. Working under internal research and development funding, NASA Marshall Space Flight Center (MSFC) has been conducting the development of a 9,000-lbf LOX/LH2 lunar lander descent engine testbed. This paper highlights the detailed design and analysis efforts to develop the lander engine Fuel Turbopump (FTP) whose operating speeds range from 30,000-rpm to 100,000-rpm. The capability of the FTP to operate across this wide range of speeds imposes several structural and dynamic challenges, and the small size of the FTP creates scaling and manufacturing challenges that are also addressed in this paper.

Alarez, Erika↗
NASA NTRS2010

Lunar Surface Access Module Descent Engine Turbopump Technology: Detailed Design

The need for a high specific impulse LOX/LH2 pump-fed lunar lander engine has been established by NASA for the new lunar exploration architecture. Studies indicate that a 4-engine cluster in the thrust range of 9,000-lbf each is a candidate configuration for the main propulsion of the manned lunar lander vehicle. The lander descent engine will be required to perform multiple burns including the powered descent onto the lunar surface. In order to achieve the wide range of thrust required, the engines must be capable of throttling approximately 10:1. Working under internal research and development funding, NASA Marshall Space Flight Center (MSFC) has been conducting the development of a 9,000-lbf LOX/LH2 lunar lander descent engine technology testbed. This paper highlights the detailed design and analysis efforts to develop the lander engine Fuel Turbopump (FTP) whose operating speeds range from 30,000-rpm to 100,000-rpm. The capability of the FTP to operate across this wide range of speeds imposes several structural and dynamic challenges, and the small size of the FTP creates scaling and manufacturing challenges that are also addressed in this paper.

Alvarez, Erika↗
NASA NTRS2009

Nuclear Thermal Rocket/Vehicle Characteristics And Sensitivity Trades For NASA's Mars Design Reference Architecture (DRA) 5.0 Study

This paper summarizes Phase I and II analysis results from NASA's recent Mars DRA 5.0 study which re-examined mission, payload and transportation system requirements for a human Mars landing mission in the post-2030 timeframe. Nuclear thermal rocket (NTR) propulsion was again identified as the preferred in-space transportation system over chemical/aerobrake because of its higher specific impulse (I(sub sp)) capability, increased tolerance to payload mass growth and architecture changes, and lower total initial mass in low Earth orbit (IMLEO) which is important for reducing the number of Ares-V heavy lift launches and overall mission cost. DRA 5.0 features a long surface stay (approximately 500 days) split mission using separate cargo and crewed Mars transfer vehicles (MTVs). All vehicles utilize a common core propulsion stage with three 25 klbf composite fuel NERVA-derived NTR engines (T(sub ex) approximately 2650 - 2700 K, p(sub ch) approximately 1000 psia, epsilon approximately 300:1, I(sub sp) approximately 900 - 910 s, engine thrust-toweight ratio approximately 3.43) to perform all primary mission maneuvers. Two cargo flights, utilizing 1-way minimum energy trajectories, pre-deploy a cargo lander to the surface and a habitat lander into a 24-hour elliptical Mars parking orbit where it remains until the arrival of the crewed MTV during the next mission opportunity (approximately 26 months later). The cargo payload elements aerocapture (AC) into Mars orbit and are enclosed within a large triconicshaped aeroshell which functions as payload shroud during launch, then as an aerobrake and thermal protection system during Mars orbit capture and subsequent entry, descent and landing (EDL) on Mars. The all propulsive crewed MTV is a 0-gE vehicle design that utilizes a fast conjunction trajectory that allows approximately 6-7 month 1-way transit times to and from Mars. Four 12.5 kW(sub e) per 125 square meter rectangular photovoltaic arrays provide the crewed MTV with approximately 50 kW(sub e) of electrical power in Mars orbit for crew life support and spacecraft subsystem needs. Vehicle assembly involves autonomous Earth orbit rendezvous and docking between the propulsion stages, in-line propellant tanks and payload elements. Nine Ares-V launches -- five for the two cargo MTVs and four for the crewed MTV -- deliver the key components for the three MTVs. Details on mission, payload, engine and vehicle characteristics and requirements are presented and the results of key trade studies are discussed.

Borowski, Stanley K.↗
NASA NTRS2008

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

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

Klem, Mark D.↗
NASA NTRS2020

A Flexible Lunar Architecture for Exploration (FLARE) Supporting NASA’s Artemis Program

The Flexible Lunar Architecture for Exploration (FLARE) is a concept to deliver four crew to the lunar surface for a minimum of seven days and then return them safely to Earth. FLARE can be implemented whenever the component vehicles are operational. FLARE was developed as an alternative to NASA’s Human Landing System (HLS) reference architecture from the Design Analysis Cycle (DAC) #2 created in 2019. The DAC2 guidelines required utilization of the Gateway vehicle in a Near- Rectilinear Halo Orbit (NRHO). Instead, FLARE chooses a Low Lunar Frozen Polar Orbit (LLFPO) for lunar rendezvous of components, and an optional Gateway vehicle. The LLFPO provides a stable orbit that overflies the south pole every 2 h, ensuring easy access to the lunar surface for surface aborts with a much lower propellant requirement than NRHO. The minimum FLARE concept uses one Space Launch System (SLS) launch, one Orion, one European Service Module (ESM), and one human lander (launched on commercial vehicle(s)). FLARE adds the SpaceTug, based upon the mature and successful ULA “Common” Centaur Upper Stage vehicle, with modifications to create an Earth-Moon transfer vehicle. In the FLARE baseline mission, the SpaceTug provides propulsion needed to return the Orion + ESM from LLFPO to Earth. The SpaceTug also provides propulsion to deliver the separate human lander components – the Descent Element (DE) and the Ascent Element (AE) - from Low Earth Orbit (LEO) to LLFPO. The SLS Block 1 then launches the Orion + ESM and completes a rendezvous with the mated DE + AE components in LLFPO. FLARE offers optional phases beyond the baseline mission. The SpaceTug can deliver components of the planned Gateway, including the Power and Propulsion Element (PPE) and the Habitable and Logistics Outpost (HALO), to LLFPO. FLARE provides an option to deliver precursor equipment to the lunar surface to enhance and extend the human mission. With these components, including an inflatable habitation module and airlock, individual crew mobility vehicle(s), an In-Situ Resource Utilization (ISRU) demonstration, and science and technology experiments, the crew can explore and conduct science on the lunar surface for up to 14 days.

Moon 2024↗
NASA NTRSDate not supplied

Computational Investigation of Retropropulsion Operating Environments with a GPU-Enabled Detached Eddy Simulation Approach

Human exploration of the surface of Mars will require an extended powered descent phase of flight, during which aerodynamic-propulsive interference effects can be significant. Characterization of these environments to enable implementation of this technology into a flight vehicle will rely heavily on computational simulation. This work advances the understanding of retropropulsion aerodynamics through application of a massively parallel detached eddy simulation approach on a GPU-accelerated computational framework, yielding data that are largely unachievable with conventional high-performance computing resources. This work includes time-dependent and time-averaged forces and moments on a conceptual, full-scale vehicle in environments and operating conditions relevant to human Mars exploration. Conditions are examined where the engine exhaust flow transitions between over-expanded and under-expanded flow structures, and flight operation will require the ability to maintain control of the vehicle during such a transition. These transitions occur as the vehicle decelerates, and as such, this investigation includes supersonic, transonic, and subsonic flight conditions. Options for vehicle control during powered flight include differential throttling of the engines. This paper provides an overview of the computational campaign, approach, and discussion of results in characterizing the resulting aerodynamics for differential throttling with retropropulsion in atmospheric environments.

EDL↗
NASA NTRS1987

Rekindled vision of hypersonic travel

NASA has joined with the DOD to conduct the National Aerospace Plane (NASP) program, whose experimental test vehicle will be designated the X-30. NASP will study the X-30's takeoff from a runway under its own power, acceleration to high Mach number on the basis of airbreathing propulsion, emergence into LEO, reentry into the earth atmosphere, and descent to a powered horizontal landing. NASP will thereby generate technology base data for three distinct types of aircraft: upper-atmosphere hypersonic-cruise aircraft, LEO space transports, and military transatmospheric vehicles. The current concept-validation phase of NASP focuses on airbreathing propulsion, lightweight/high-strength heat-resistant materials, and computational fluid dynamics.

Colladay, Raymond S.↗
NASA NTRS2015

Resource Prospector Propulsion System Cold Flow Testing

Resource Prospector (RP) is a NASA mission being led by NASA Ames Research Center with current plans to deliver a scientific payload package aboard a rover to the lunar surface. As part of an early risk reduction activity, Marshall Space Flight Center (MSFC) and Johnson Space Flight Center (JSC) have jointly developed a government-version concept of a lunar lander for the mission. The spacecraft consists of two parts, the lander and the rover which carries the scientific instruments. The lander holds the rover during launch, cruise, and landing on the surface. Following terminal descent and landing the lander portion of the spacecraft become dormant after the rover embarks on the science mission. The lander will be equipped with a propulsion system for lunar descent and landing, as well as trajectory correction and attitude control maneuvers during transit to the moon. Hypergolic propellants monomethyl hydrazine and nitrogen tetroxide will be used to fuel sixteen 70-lbf descent thrusters and twelve 5-lbf attitude control thrusters. A total of four metal-diaphragm tanks, two per propellant, will be used along with a high-pressure composite-overwrapped pressure vessel for the helium pressurant gas. Many of the major propulsion system components are heritage missile hardware obtained by NASA from the Air Force. In parallel with the flight system design activities, a simulated propulsion system based on flight drawings was built for conducting a series of water flow tests 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. The primary objective of the cold flow testing was to simulate the RP propulsion system fluid flow operation through water flow testing 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. All design and build efforts, including the analytical modeling, have been performed. The cold flow testing of the propulsion system was set up and conducted at a NASA MSFC test facility. All testing was completed in the summer of 2014, and this paper documents the results of that testing and the associated fluid system modeling efforts.

Williams, Hunter↗
NASA NTRS2010

COMPASS Final Report: Low Cost Robotic Lunar Lander

The COllaborative Modeling for the Parametric Assessment of Space Systems (COMPASS) team designed a robotic lunar Lander to deliver an unspecified payload (greater than zero) to the lunar surface for the lowest cost in this 2006 design study. The purpose of the low cost lunar lander design was to investigate how much payload can an inexpensive chemical or Electric Propulsion (EP) system deliver to the Moon s surface. The spacecraft designed as the baseline out of this study was a solar powered robotic lander, launched on a Minotaur V launch vehicle on a direct injection trajectory to the lunar surface. A Star 27 solid rocket motor does lunar capture and performs 88 percent of the descent burn. The Robotic Lunar Lander soft-lands using a hydrazine propulsion system to perform the last 10% of the landing maneuver, leaving the descent at a near zero, but not exactly zero, terminal velocity. This low-cost robotic lander delivers 10 kg of science payload instruments to the lunar surface.

McGuire, Melissa L.↗
NASA NTRS2018

Propulsion System Options for a Potential Sample Return Lander (SRL) for Mars

A potential Mars Sample Return (MSR) campaign would be very unique challenge and has been the focus of technology development and planning efforts at JPL for several decades. The latest mission concept study focuses on a potential Sample Return Lander (SRL). The potential SRL examined several propulsion system approaches using monopropellant and bi-propellants for the terminal descent phase. A ranking was developed based on a mathematical optimization program. This paper focuses on the evaluations conducted for these system options. Historically, monopropellant and bi-propellant propulsion systems are usually considered for this application; this study took a fresh look at both conventional State of the Art (SOA) pressure fed monopropellant and bi-propellant propulsion system and augmenting the system using small Electrically Driven Pumps (EDP) for both these propulsion systems.

Preudhomme, Michael↗
NASA NTRS2021

Mars Reconnaissance Orbiter Maneuver Plan Following Mars 2020 Landing

The Mars Reconnaissance Orbiter spacecraft continues to perform valuable science observations at Mars, provide telecommunication relay for surface assets, and characterize landing sites for future missions. The spacecraft provided the primary relay support for the Mars 2020 mission during entry, descent, and landing on February 18, 2021. This paper discusses the propulsive maneuver plan following the Mars 2020 landing to return the spacecraft to its primary science orbit by November 2021 and maintain it through 2029. Alternate maneuver plans considered for supporting Mars 2020 surface operations and the entry, descent, and landing of the ExoMars 2022 mission are also described.

Menon, Premkumar R.↗
NASA NTRS2013

Descent and Landing Triggers for the Orion Multi-Purpose Crew Vehicle Exploration Flight Test-1

The Orion Multi-Purpose Crew Vehicle (MPCV) will perform a flight test known as Exploration Flight Test-1 (EFT-1) currently scheduled for 2014. One of the primary functions of this test is to exercise all of the important Guidance, Navigation, Control (GN&C), and Propulsion systems, along with the flight software for future flights. The Descent and Landing segment of the flight is governed by the requirements levied on the GN&C system by the Landing and Recovery System (LRS). The LRS is a complex system of parachutes and flight control modes that ensure that the Orion MPCV safely lands at its designated target in the Pacific Ocean. The Descent and Landing segment begins with the jettisoning of the Forward Bay Cover and concludes with sensing touchdown. This paper discusses the requirements, design, testing, analysis and performance of the current EFT-1 Descent and Landing Triggers flight software.

Bihari, Brian D.↗
NASA NTRS2015

NASA Propulsion Sub-System Concept Studies and Risk Reduction Activities for Resource Prospector Lander

NASA's exploration roadmap is focused on developing technologies and performing precursor missions to advance the state of the art for eventual human missions to Mars. One of the key components of this roadmap is various robotic missions to Near-Earth Objects, the Moon, and Mars to fill in some of the strategic knowledge gaps. The Resource Prospector (RP) project is one of these robotic precursor activities in the roadmap. RP is a multi-center and multi-institution project to investigate the polar regions of the Moon in search of volatiles. The mission is rated Class D and is approximately 10 days, assuming a five day direct Earth to Moon transfer. Because of the mission cost constraint, a trade study of the propulsion concepts was conducted with a focus on available low-cost hardware for reducing cost in development, while technical risk, system mass, and technology advancement requirements were also taken into consideration. The propulsion system for the lander is composed of a braking stage providing a high thrust to match the lander's velocity with the lunar surface and a lander stage performing the final lunar descent. For the braking stage, liquid oxygen (LOX) and liquid methane (LCH4) propulsion systems, derived from the Morpheus experimental lander, and storable bi-propellant systems, including the 4th stage Peacekeeper (PK) propulsion components and Space Shuttle orbital maneuvering engine (OME), and a solid motor were considered for the study. For the lander stage, the trade study included miniaturized Divert Attitude Control System (DACS) thrusters (Missile Defense Agency (MDA) heritage), their enhanced thruster versions, ISE-100 and ISE-5, and commercial-off-the-shelf (COTS) hardware. The lowest cost configuration of using the solid motor and the PK components while meeting the requirements was selected. The reference concept of the lander is shown in Figure 1. In the current reference configuration, the solid stage is the primary provider of delta-V. It will generate 15,000-lbf of thrust with a single burn of ~ 80's seconds. The lander stage is a bi-propellant, pressure-regulated, pulsing liquid propulsion system to perform all other functions.

Trinh, Huu P.↗
NASA NTRS2013

Mars Science Laboratory Cruise Propulsion Maneuvering Operations

Mars Science Laboratory "Curiosity" is NASA's most recent mission to Mars, launched in November 2011, and landed in August 2012. It is a subcompact car-sized nuclear powered rover designed for a long duration mission, with an extensive suite of science instruments. Entry, descent and landing used a unique "skycrane" concept. This report describes the propulsive maneuvering operations during cruise from Earth to Mars, to control attitudes and to target the vehicle for entry. The propulsion subsystem, mission operations, and flight performance are discussed. All trajectory control maneuvers were well within accuracy requirements, and all turns and spin corrections were nominal.

Baker, Raymond S.↗
NASA NTRS2020

Mid Lift-to-Drag Ratio Rigid Vehicle 6-DoF Performance for Human Mars Entry, Descent, and Landing: A Fractional Polynomial Powered Descent Guidance Approach

Defining a feasible vehicle design and mission architecture capable of reliably delivering apayload of 20 metric tons (mt) or more is a great challenge for landing humans on Mars. TheMid Lift-to-Drag Rigid Vehicle (MRV), a rigid decelerator studied in NASA’s Entry, Descent,and Landing Architecture Study (EDLAS), has shown to be a viable vehicle candidate forfuture human Mars missions. As the vehicle concept matures, models of increasing fidelity areadded to the six-degree-of-freedom (6DoF) EDL simulation. This paper presents 6DoFsimulation results using model updates for vehicle mass properties, fineness ratio, andaerodynamic-propulsive interactions. Additionally, an assessment of the Fractional-Polynomial Powered Descent Guidance (FP2DG) performance is presented, and the vehicleperformance is compared with the Tunable Apollo Powered Descent Guidance (TAPDG).Finally, Monte Carlo results of the vehicle design trades are presented.

Johnson, Breanna J.↗
NASA NTRS2010

Cryogenic Propellant Management Device: Conceptual Design Study

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.

Wollen, Mark↗
NASA NTRS2020

Computational Investigation of Powered Descent for Human-Scale Mars Landers

NASA human-scale Mars entry, descent, and landing vehicles currently under consideration have system requirements beyond the current state of the art. Specifically, heritage approaches for deceleration using parachutes and initiating engines at subsonic speeds are not adequate to land these large vehicles (50-60 metric ton entry mass) precisely (within 50 meters of a surface target). Therefore, the present approach for landing humans on Mars utilizes entry vehicles with large aerodynamic surface areas that transition to the descent and landing phase by initiating retropropulsion at supersonic speeds. Vehicle performance, including stability and control, depends on accurate characterization of the aerodynamic-propulsive interference effects, especially at engine initiation, when the vehicles are transitioning from an entry attitude to that suitable for descent and landing. This paper summarizes the human Mars entry vehicle configurations, retropropulsion integration, performance assumptions, and expected flight environment at engine initiation. Cases spanning supersonic, transonic, and subsonic conditions during powered descent are analyzed using four computational fluid dynamics codes, all of which have been previously applied to supersonic retropropulsion at varying scales and conditions. The results of the analysis are used to develop an aerodynamic-propulsive interference model that has since been applied in trajectory simulations to augment the fidelity of human-scale Mars entry vehicle performance.

Ashley M Korzun↗
NASA NTRS1999

Human Mars Mission: Weights and Mass Properties

This paper presents a final report on The Human Mars Mission Weights and Measures. The topics included in this report are: 1) Trans-Earth Injection Storage Human Mars Mission (HMM) Chemical Design Reference Mission (DRM) v4.0a Weight Breakout; 2) Ascent Stage HMM Chemical DRM v4.0a Weight Breakout; 3) Descent Stages HMM Chemical DRM v4.0a Weight Breakout; 4) Trans-Mars Injection Stages HMM Chemical DRM v4.0a Weight Breakout; 5) Trans-Earth Injection Stage HMM Solar Electric Propulsion (SEP) DRM v4.0a Weight Breakout; 6) Ascent Stage HMM SEP DRM v4.0a Weight Breakout; 7) Descent Stages HMM SEP DRM v4.0a Weight Breakout; 8) Trans-Mars Injection Stages HMM SEP DRM v4.0a Weight Breakout; 9) Crew Taxi Stage HMM SEP DRM v4.0 Weight Breakout; 10)Trans-Earth Injection Stage HMM Nuclear DRM v4.0a Weight Breakout; 11) Ascent Stage HMM Nuclear DRM v4.0a Weight Breakout; 12) Descent Stages HMM Nuclear DRM v4.0a Weight Breakout; 13) Trans-Mars Injection Stages HMM Nuclear DRM v4.0a Weight Breakout; and 14) HMM Mass Properties Coordinate System.

Brothers, Bobby↗