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NASA NTRS1998

A Status of the Advanced Space Transportation Program from Planning to Action

A Technology Plan for Enabling Commercial Space Business was presented at the 48th International Astronautical Congress in Turin, Italy. This paper presents a status of the program's accomplishments. Technology demonstrations have progressed in each of the four elements of the program; (1) Low Cost Technology, (2) Advanced Reusable Technology, (3) Space Transfer Technology and (4) Space Transportation Research. The Low Cost Technology program element is primarily focused at reducing development and acquisition costs of aerospace hardware using a "design to cost" philosophy with robust margins, adapting commercial manufacturing processes and commercial off-the-shelf hardware. The attributes of this philosophy for small payload launch are being demonstrated at the component, sub-system, and system level. The X-34 "Fastrac" engine has progressed through major component and subsystem demonstrations. A propulsion system test bed has been implemented for system-level demonstration of component and subsystem technologies; including propellant tankage and feedlines, controls, pressurization, and engine systems. Low cost turbopump designs, commercial valves and a controller are demonstrating the potential for a ten-fold reduction in engine and propulsion system costs. The Advanced Reusable Technology program element is focused on increasing life through high strength-to-weight structures and propulsion components, highly integrated propellant tanks, automated checkout and health management and increased propulsion system performance. The validation of rocket based combined cycle (RBCC) propulsion is pro,-,ressing through component and subsystem testing. RBCC propulsion has the potential to provide performance margin over an all rocket system that could result in lower gross liftoff weight, a lower propellant mass fraction or a higher payload mass fraction. The Space Transfer Technology element of the program is pursuing technology that can improve performance and dramatically reduce the propellant and structural mass of orbit transfer and deep space systems. Flight demonstration of ion propulsion is progressing towards launch. Ion propulsion is the primary propulsion for Deep Space 1; a flyby of comet West-kohoutek-lkemura and asteroid 3352 McAuliffe. Testing of critical solar-thermal propulsion subsystems have been accomplished and planning is continuing for the flight demonstration of an electrodynamic tether orbit transfer system. The forth and final element of the program, Space Transportation Research, has progressed in several areas of propulsion research. This element of the program is focused at long-term (25 years) breakthrough concepts that could bring launch costs to a factor of one hundred below today's cost or dramatically expand planetary travel and enable interstellar travel.

Lyles, Garry↗

NASA NTRS2015

Three Orbital Burns to Molniya Orbit Via Shuttle_Centaur G Upper Stage

An unclassified analytical trajectory design, performance, and mission study was done for the 1982 to 1986 joint National Aeronautics and Space Administration (NASA)-United States Air Force (USAF) Shuttle/Centaur G upper stage development program to send performance-demanding payloads to high orbits such as Molniya using an unconventional orbit transfer. This optimized three orbital burn transfer to Molniya orbit was compared to the then-baselined two burn transfer. The results of the three dimensional trajectory optimization performed include powered phase steering data and coast phase orbital element data. Time derivatives of the orbital elements as functions of thrust components were evaluated and used to explain the optimization's solution. Vehicle performance as a function of parking orbit inclination was given. Performance and orbital element data was provided for launch windows as functions of launch time. Ground track data was given for all burns and coasts including variation within the launch window. It was found that a Centaur with fully loaded propellant tanks could be flown from a 37 deg inclination low Earth parking orbit and achieve Molniya orbit with comparable performance to the baselined transfer which started from a 57 deg inclined orbit: 9,545 versus 9,552 lb of separated spacecraft weight, respectively. There was a significant reduction in the need for propellant launch time reserve for a 1 hr window: only 78 lb for the three burn transfer versus 320 lb for the two burn transfer. Conversely, this also meant that longer launch windows over more orbital revolutions could be done for the same amount of propellant reserve. There was no practical difference in ground tracking station or airborne assets needed to secure telemetric data, even though the geometric locations of the burns varied considerably. There was a significant adverse increase in total mission elapsed time for the three versus two burn transfer (12 vs. 1-1/4 hr), but could be accommodated by modest modifications to Centaur systems. Future applications were discussed. The three burn transfer was found to be a viable, arguably preferable, alternative to the two burn transfer.

Orbital trajectory design↗

NASA NTRSDate not supplied

Parametric Studies of Human Mars Entry, Descent, and Landing Systems

This paper will present a parametric analysis for entry, descent, and landing (EDL) concepts, enabling rapid systems assessment and tradespace exploration. The entry system uses a hypersonic inflatable aerodynamic decelerator (HIAD) technology. The baseline system includes elements for Mars aerocapture (AC) and EDL segments of the mission. The mission concept of operations (ConOps) begins at Mars arrival in a polar inclination. After performing an aerocapture (AC) maneuver into a pre-defined Mars parking orbit, the AC HIAD is jettisoned. The rest of thevehicle stays in the parking orbit for up to one year. The EDL sequence starts with a deorbit burn at the apoapsis of the parking orbit. After hypersonic entry with the EDL HIAD, the entry system uses a supersonic retropropulsion maneuver to slow the vehicle for the descent and landing segments of the mission. The vehicle will maintain a constant velocity of 2.5 m/s for 5 seconds prior to landing.The system includes a Mars Ascent Vehicle (MAV), Mars Descent Module (MDM), and two HIADs. The MDM includes a primary structure, tanks, engines, and radiators. The primary structure is an aluminum-lithium (Al-Li)cruciform design similar to the structural design of the Apollo Lunar Module. The cruciform planform layout results in four outer bays, with adequate volume in the corners between outer bays to package four landing gears. The central bay is reserved for packaging the MAV and the recessed MAV engines. Two of the outer bays accommodate main propellant tanks, with one LOX and one CH4 tank in each bay. The two remaining outer bays each house four rocket engine systems and associated support structure. Each HIAD comprises an inflatable structure, flexible thermal protection system, gas, and gas generators. The HIAD design used in this study is a stacked-toroid concept with pairing loop straps and radial/chevron straps. The baseline system lands a 20-t payload on the Mars surface. It is assumed the vehicle arrives at 6.2 km/s relative velocity at 90° inclination and is captured to a 1-Sol parking orbit.The baseline design includes many assumptions such as margins, arrival state, ConOps options, parking orbit, physical dimensions, propellant options, and technology concepts. The impact of these parameters are quantified through systems-level sensitivity analyses, which capture the global impact—not at a component level—but at the systems level. The systems-level sensitivities expose major design drivers and importance of each assumption for a design.Through tradespace exploration, a wide range of systems parameters are examined and compared for several feasible design options. Studies have been completed for the following input parameters: payload mass, propellant options, AC/EDL ballistic coefficient, lander thrust to weight ratio (T/W) (surrogate for the maximum EDL g’s),engine specific impulse (Isp), parking orbit, and inclination.The final paper will present and discuss the parametric approach used in the study. It will also include the results of recent systems analyses, sensitivity analyses, and tradespace exploration

Jamshid Samareh↗

NASA NTRS2014

An RF Sensor for Gauging Screen-Channel Liquid Acquisition Devices for Cryogenic Propellants

A key requirement of a low-gravity screen-channel liquid acquisition device (LAD) is the need to retain 100% liquid in the channel in response to propellant outflow and spacecraft maneuvers. The point at which a screen-channel LAD ingests vapor is known as breakdown, and can be measured several different ways such as: visual observation of bubbles in the LAD channel outflow; a sudden change in pressure drop between the propellant tank and LAD sump outlet; or, an indication by wet-dry sensors placed in the LAD channel or outflow stream. Here we describe a new type of sensor for gauging a screen-channel LAD, the Radio Frequency Mass Gauge (RFMG). The RFMG measures the natural electromagnetic modes of the screen-channel LAD, which is very similar to an RF waveguide, to determine the amount of propellant in the channel. By monitoring several of the RF modes, we show that the RFMG acts as a global sensor of the LAD channel propellant fill level, and enables detection of LAD breakdown even in the absence of outflow. This paper presents the theory behind the RFMG-LAD sensor, measurements and simulations of the RF modes of a LAD channel, and RFMG detection of LAD breakdown in a channel using a simulant fluid during inverted outflow and long-term stability tests.

microwave sensor↗

NASA NTRS2003

Propulsion Flight-Test Fixture

NASA Dryden Flight Research Center s new Propulsion Flight Test Fixture (PFTF), designed in house, is an airborne engine-testing facility that enables engineers to gather flight data on small experimental engines. Without the PFTF, it would be necessary to obtain such data from traditional wind tunnels, ground test stands, or laboratory test rigs. Traditionally, flight testing is reserved for the last phase of engine development. Generally, engines that embody new propulsion concepts are not put into flight environments until their designs are mature: in such cases, either vehicles are designed around the engines or else the engines are mounted in or on missiles. However, a captive carry capability of the PFTF makes it possible to test engines that feature air-breathing designs (for example, designs based on the rocket-based combined cycle) economically in subscale experiments. The discovery of unknowns made evident through flight tests provides valuable information to engine designers early in development, before key design decisions are made, thereby potentially affording large benefits in the long term. This is especially true in the transonic region of flight (from mach 0.9 to around 1.2), where it can be difficult to obtain data from wind tunnels and computational fluid dynamics. In January 2002, flight-envelope expansion to verify the design and capabilities of the PFTF was completed. The PFTF was flown on a specially equipped supersonic F-15B research testbed airplane, mounted on the airplane at a center-line attachment fixture, as shown in Figure 1. NASA s F-15B testbed has been used for several years as a flight-research platform. Equipped with extensive research air-data, video, and other instrumentation systems, the airplane carries externally mounted test articles. Traditionally, the majority of test articles flown have been mounted at the centerline tank-attachment fixture, which is a hard-point (essentially, a standardized weapon-mounting fixture). This hard-point has large weight margins, and, because it is located near the center of gravity of the airplane, the weight of equipment mounted there exerts a minimal effect on the stability and controllability of the airplane. The PFTF (see Figure 2) includes a one-piece aluminum structure that contains space for instrumentation, propellant tanks, and feed-system components. The PFTF also houses a force balance, on which is mounted the subscale engine or other experimental apparatus that is to be the subject of a flight test. The force balance measures a combination of inertial and aerodynamic forces and moments acting on the experimental apparatus.

Palumbo, Nate↗

NASA NTRS2022

Passive Cooling in Aerogel-Based Insulation Systems for Liquid Hydrogen Upper Stage Launch Vehicle Tanks

Spray-On Foam Insulation (SOFI) is typically used to protect the windward-facing side of flight tanks for space launch vehicle liquid hydrogen (LH2) stages. SOFI is an excellent insulator in ambient pressure environments, however, its performance pales in comparison to reflective-type systems such as Multi-Layer Insulation in the thermal radiation dominated vacuum environment of space. If a windward insulation system employing radiation shields could be devised to replace or supplement SOFI, the on-orbit heat load could be drastically reduced, and the residual propellants could then be used to facilitate secondary missions. Such insulation systems, utilizing aerogel blanket insulation, have been explored by the Cryogenics Test Laboratory at NASA Kennedy Space Center. Owing to its nano-porous structure, aerogel is an excellent adsorbent as well as an insulator at cryogenic temperatures and will readily uptake condensable background gasses such as air when protecting a surface near LH2 temperatures. When an adsorbed blanket is rapidly exposed to a vacuum, such as during the ascent of a rocket, it will release the background gas, producing a passive cooling effect that could potentially reduce or eliminate the heat load on the propellant tank for a time. This work discusses the setup and results of a test program carried out to determine the effectiveness and impact of this cooling effect as it pertains to LH2 upper stage tanks. A vertical-cylindrical liquid nitrogen (LN2) calorimeter test instrument acted as the upper stage analog. The cold-mass was wrapped with two layers of 10 mm thick aerogel blanket material, with and without aluminum heat shields depending on the test, and interspersed with numerous thermocouple temperature sensors. Gaseous argon was used as a stand-in for air as the condensable background gas. The test procedure mimicked vehicle cryo-loading/stabilization, and ascent phases, with the ascent phase accomplished using parallel vacuum pumps. Results show that the desorption cooling effect was enough to not only eliminate the heat load on the LN2 cold-mass entirely, but further refrigerated the liquid below its boiling point.

A.M. Swanger↗

NASA NTRS1999

Summary of LO2/Ethanol OMS/RCS Technology and Advanced Development 99-2744

NASA is pursuing non-toxic propellant technologies applicable to RLV and Space Shuttle orbital maneuvering system (OMS) and reaction control system (RCS). The primary objectives of making advancements in an OMS/RCS system are improved safety, reliability, and reduced operations and maintenance cost, while meeting basic operational and performance requirements. An OMS/RCS has a high degree of direct interaction with the vehicle and crew and requires subsystem and components that are compatible with integration into the vehicle with regard to external mold-line, power, and thermal control. In July 1997, a Phase I effort for the technology and advanced development of an upgrade of the space shuttle was conducted to define the system architecture, propellant tank, feed system, RCS thrusters, and OMS engine. Phase I of the project ran from July 1997 to October 1998. Phase II is currently being planned for the development and test of full-scale prototype of the system in 1999 and 2000. The choice of pressure-fed liquid oxygen (LO2) and ethanol is the result of numerous trade studies conducted from 1980 to 1996. Liquid oxygen and ethanol are clean burning, high-density propellants that provide a high degree of commonality with other spacecraft subsystems including life support, power, and thermal control, and with future human exploration and development of space missions. The key to this pressure-fed system is the use of subcooled liquid oxygen at 350 psia. In this approach, there is 80 degrees R of subcooling, which means that boil-off will not occur until the temperature has risen 80 R. The sub-cooling results naturally from loading propellants at 163 R, which is the saturation temperature at 14.7 psia, and then pressurizing to 350 psia on the launch pad. Thermal insulation and conditioning techniques are then used to limit the LO2 temperature to 185 R maximum, and maintain the sub-cooling. The other key is the wide temperature range of ethanol, -173 F to +300 F, which can provide heat to gasify liquid oxygen or provide a good coolant.

Curtis, Leslie A.↗

NASA NTRSDate not supplied

Validation of a Multiphase Computational Fluid Dynamics Model for Vapor Pull-Through in Normal and Low Gravity

On-orbit fluid transfer such as refueling of propellant tanks and life-support systems can enable long-duration space missions. For safe and efficient liquid transfer operations, prior knowledge of liquid positioning and liquid-vapor interface behavior while draining in a low-gravity environment is required. Numerical models capable of predicting vapor ingestion (or vapor pull-through) can be used to design liquid transfer operations while reducing liquid residuals, mission risk and settling thrust required to prevent vapor ingestion. An experimental program conducted in the 2.2 Second Drop Tower facility at NASA Glenn Research Center investigated the vapor ingestion phenomenon for a range of outflow rates and tank sizes providing a database for validation. This study presents a Computational Fluid Dynamics model capable of accurately predicting the vapor ingestion using the Volume-of-Fluid multiphase solver in commercial code STAR-CCM+. A description of the experimental setup and general trends from similar studies are presented. Comparisons of the numerical prediction and test data in normal and low gravity show good agreement and give confidence in pursuing design of full-scale propellant transfer systems.

Computational Fluid Dynamics↗

NASA NTRSDate not supplied

Validation of a Multiphase Computational Fluid Dynamics Model for Vapor Pull-Through in Normal and Low Gravity

On-orbit fluid transfer such as refueling of propellant tanks and life-support systems can enable long-duration space missions. For safe and efficient liquid transfer operations, prior knowledge of liquid positioning and liquid-vapor interface behavior while draining in a reduced-gravity environment is required. Numerical models capable of predicting vapor ingestion (or vapor pull-through) can be used to design liquid transfer operations while reducing liquid residuals, mission risk and settling thrust required to prevent vapor ingestion. An experimental program conducted in the 2.2 Second Drop Tower facility at NASA Lewis Research Center in 1969 investigated the vapor ingestion phenomenon for a range of outflow rates and tank sizes providing a database for validation. This study presents a Computational Fluid Dynamics model capable of accurately predicting the vapor ingestion using the Volume-of-Fluid multiphase solver in commercial code STAR-CCM+. A description of the experimental setup and general trends from similar studies are presented. Comparisons of the numerical prediction and test data in normal and low gravity show good agreement and give confidence in pursuing design of full-scale propellant transfer systems.

Computational Fluid Dynamics↗

NASA NTRSDate not supplied

Validation of A Multiphase Computational Fluid Dynamics Model for Vapor Pull-Through in Normal and Low Gravity

On-orbit fluid transfer such as refueling of propellant tanks and life-support systems can enable long-duration space missions. For safe and efficient liquid transfer operations, prior knowledge of liquid positioning and liquid-vapor interface behavior while draining in a reduced-gravity environment is required. Numerical models capable of predicting vapor ingestion (or vapor pull-through) can be used to design liquid transfer operations while reducing liquid residuals, mission risk and settling thrust required to prevent vapor ingestion. An experimental program conducted in the 2.2 Second Drop Tower facility at NASA Lewis Research Center in 1969 investigated the vapor ingestion phenomenon for a range of outflow rates and tank sizes providing a database for validation. This study presents a Computational Fluid Dynamics model capable of accurately predicting the vapor ingestion using the Volume-of-Fluid multiphase solver in commercial code STAR-CCM+. A description of the experimental setup and general trends from similar studies are presented. Comparisons of the numerical prediction and test data in normal and low gravity show good agreement and give confidence in pursuing design of full-scale propellant transfer systems.

Propellant Transfer↗

NASA NTRSDate not supplied

Exploring Cryogenic Propellant Behavior in Low-Gravity Environments, Insights from the Saturn AS-203 Vent Experiments and CFD Analysis

In the 1960s, NASA embarked on a series of groundbreaking flight tests on the Saturn AS-203, aiming to understand the complex dynamics of propellants in the distinctive low-gravity lunar environment. These tests centered on venting experiments, subjecting cryogenic liquid hydrogen to conditions beneath its saturation pressure while accelerating the vehicle to manage the propellant's positioning. During these experiments, NASA meticulously scrutinized the propellant tank using a suite of instruments, including temperature and pressure sensors, as well as a camera placed internal to the liquid hydrogen tank. The outcomes provided anecdotal evidence revealing the phenomenon of boiling along the tank's walls and the intriguing formation of liquid globules and droplets in the ullage during the venting process. Notably, the substantial drop in liquid temperature during these tests suggests adiabatic cooling as liquid hydrogen evaporates. This evaporation leads to a cooling of the remaining hydrogen due to the heat it releases. This paper presents the outcomes of our initial analysis, wherein CFD models were used to simulate the observed boiling phenomena and the bulk movement of the liquid hydrogen propellant, both qualitatively and quantitatively. The implications of these findings may extend to mission and vehicle designers, providing invaluable insights for crafting more efficient and effective in-space propulsion systems utilizing cryogenic propellant including impacts to vehicle control systems. Understanding propellant behavior under these conditions may better inform GNC teams, ensuring more stable vehicle operations when utilizing cryogenic propellants. This includes essential considerations for cryogenic propellant transfer and storage systems, integral to NASA's forthcoming Artemis missions. While we recognize the challenges tied to CFD models, this study represents a step forward, highlighting current progress and signaling the potential for refining our predictive understanding in the future.

Computational Fluid Dynamics↗

NASA NTRSDate not supplied

Exploring Cryogenic Propellant Behavior in Low-Gravity Environments, Insights from the Saturn AS-203 Vent Experiments and CFD Analysis

Computational Fluid Dynamics↗

NASA NTRS2018

Robust Exploration and Commercial Missions to the Moon Using Nuclear Thermal Rocket Propulsion and In Situ Propellants Derived from Lunar Polar Ice Deposits

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable “access through cislunar space” necessary for commercial development and sustained human presence on the Moon. It is a demonstrated technology capable of generating both high thrust and high specific impulse (I(sub sp) ~900 s)— twice that of today’s best chemical rockets. Nuclear lunar transfer vehicles—consisting of a propulsion stage using three ~16.5-klb(sub f) small nuclear rocket engines (SNREs), an in-line propellant tank, plus the payload—can enable a variety of reusable lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong “tourism” missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing a robust in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in situ propellant production used to supply strategically located propellant depots and transportation nodes. The processing of lunar polar ice (LPI) deposits (estimated to be ~2 billion metric tons) for propellant production—specifically liquid oxygen (LO(sub 2)) and hydrogen (LH(sub 2))—can significantly reduce the launch mass requirements from Earth and can enable reusable, surface-based lunar landing vehicles (LLVs) using LO(sub 2)/LH(sub 2) chemical rocket engines. Afterwards, LO(sub 2)/LH(sub 2) propellant depots can be established in lunar polar and equatorial orbits to supply the LTS. At this point a modified version of the conventional NTR called the LO(sub 2)-augmented NTR, or LANTR, would be introduced into the LTS, allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants (LDPs) for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an “afterburner” into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engine’s choked sonic throat—essentially “scramjet propulsion in reverse.” By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and I(sub sp) values while the reactor core power level remains relatively constant. A LANTR-based LTS offers unique mission capabilities including short transit time crewed cargo transports. Even a “commuter” shuttle service may be possible, allowing “one-way” trip times to and from the Moon on the order of 36 hr or less. If only 1% of the postulated trapped water ice were available for use in lunar orbit, such a supply could support routine commuter flights to the Moon for many thousands of years. This report outlines an evolving LTS architecture that uses propellants derived from LPI and examines a variety of mission types and transfer vehicle designs along with their operating characteristics and increasing demands on LDP production as mission complexity and velocity change ΔV requirements increase. A comparison of the LDP production and mining requirements using LPI and volcanic glass to produce lunar-derived liquid oxygen (LUNOX) via the hydrogen reduction process is included, and the synergy with an evolving helium-3 mining industry is also discussed.

Borowski, Stanley K.↗

NASA NTRS2017

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

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

Spacecraft design↗

NASA NTRS2017

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

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

Nuclear rocket engines↗

NASA NTRS2017

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

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

Nuclear Thermal Propulsion↗

NASA NTRS2018

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

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

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NASA NTRS1974

Reusable Agena study. Volume 1: Executive summary

The shuttle Agena upper stage interim tug concept is based on a building block approach. These building block concepts are extensions of existing ascent Agena configurations. Several current improvements, have been used in developing the shuttle/Agena upper stage concepts. High-density acid is used as the Agena upper stage oxidizer. The baffled injector is used in the main engine. The DF-224 is a fourth generation computer currently in development and will be flight proven in the near future. The Agena upper stage building block concept uses the current Agena as a baseline, adds an 8.5-inch (21.6 cm) extension to the fuel tank for optimum mixture ratio, uses monomethyl hydrazine as fuel, exchanges a 150:1 nozzle extension for the existing 45:1, exchanges an Autonetics DF-224 for the existing Honeywell computer, and adds a star sensor for guidance update. These modifications to the current Agena provide a 5-foot (1.52m) diameter shuttle/Agena upper stage that will fly all Vandenberg Air Force Base missions in the reusable mode without resorting to a kick motor. The delta V velocity of the Agena is increased by use of a strap-on propellant tank option. This option provides a shuttle/Agena upper stage with the capability to place almost 3900 pounds (1769 kg) into geosynchronous orbit (24 hour period) without the aid of kick motors.

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