Search NASA⌕ Search

SEARCH · Search NASA

Results for “lunar transfer”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Lunar transfer vehicle studies

Lunar transportation architectures exist for several different mission scenarios. Direct flights from Earth are possible, as the Apollo program clearly demonstrated. Alternatively, a space transfer vehicle could be constructed in space by using the Space Station as a base of operations, or multiple vehicles could be launched from Earth and dock in LEO without using a space station for support. Similarly, returning personnel could proceed directly to Earth or rendezvous at the Space Station for a ride back home on the Space Shuttle. Multiple design concepts exist which are compatible with these scenarios and which can support requirements of cargo, personnel, and mission objectives. Regardless of the ultimate mission selected, some technologies will certainly play a key role in the design and operation of advanced lunar transfer vehicles. Current technologies are capable of delivering astronauts to the lunar surface, but improvements are needed to affordably transfer the material and equipment that will be needed for establishing a lunar base. Materials and structures advances, in particular, will enable the development of more capable cryogenic fluid management and propulsion systems, improved structures, and more efficient vehicle assembly, servicing and processing.

Keeley, Joseph T.↗
NASA NTRS1989

Lunar transfer vehicle design issues with electric propulsion systems

This paper describes parametric design studies of electric propulsion lunar transfer vehicles. In designing a lunar transfer vehicle, selecting the 'best' operating points for the design parameters allows significant reductions in the mass in low earth orbit (LEO) for the mission. These parameters include the specific impulse, the power level, and the propulsion technology. Many of the decisions regarding the operating points are controlled by the propulsion and power system technologies that are available for the spacecraft. The relationship between these technologies is discussed and analyzed here. It is found that both ion and MPD propulsion offer significant LEO mass reductions over O2/H2 for lunar transfer vehicle missions. The recommended operating points for the lunar transfer vehicle are an I(sp) of 5000 lb(f)-s/lb(m) and a 1 MW power level. For large lunar missions, krypton may be the best choice for ion propulsion.

Palaszewski, Bryan↗
NASA NTRS2010

Targeting Low-Energy Ballistic Lunar Transfers

Numerous low-energy ballistic transfers exist between the Earth and Moon that require less fuel than conventional transfers, but require three or more months of transfer time. An entirely ballistic lunar transfer departs the Earth from a particular declination at some time in order to arrive at the Moon at a given time along a desirable approach. Maneuvers may be added to the trajectory in order to adjust the Earth departure to meet mission requirements. In this paper, we characterize the (Delta)V cost required to adjust a low-energy ballistic lunar transfer such that a spacecraft may depart the Earth at a desirable declination, e.g., 28.5(white bullet), on a designated date. This study identifies the optimal locations to place one or two maneuvers along a transfer to minimize the (Delta)V cost of the transfer. One practical application of this study is to characterize the launch period for a mission that aims to launch from a particular launch site, such as Cape Canaveral, Florida, and arrive at a particular orbit at the Moon on a given date using a three-month low-energy transfer.

Low-energy transfers↗
NASA NTRS1992

Preliminary structural design of a lunar transfer vehicle aerobrake

An aerobrake for a lunar transfer vehicle was designed through the use of the Taguchi design method, structural finite element analyses, and structural sizing routines. A minimum weight aerobrake structural configuration was the objective of the study. Six design parameters were chosen to represent the aerobrake structural configuration. The design parameters included honeycomb core thickness, diameter-to-depth ratio, shape, material, number of concentric ring frames and number of radial frames. Each parameter had 3 levels. The optimum aerobrake configuration resulting from the study was approximately half the weight of the average of all the other experimental configurations. The parameters having the most significant impact on the aerobrake structural weight were identified.

Bush, Lance B.↗
NASA NTRS1992

Preliminary structural design of a lunar transfer vehicle aerobrake

An aerobrake concept for a Lunar transfer vehicle was weight optimized through the use of the Taguchi design method, structural finite element analyses and structural sizing routines. Six design parameters were chosen to represent the aerobrake structural configuration. The design parameters included honeycomb core thickness, diameter to depth ratio, shape, material, number of concentric ring frames, and number of radial frames. Each parameter was assigned three levels. The minimum weight aerobrake configuration resulting from the study was approx. half the weight of the average of all twenty seven experimental configurations. The parameters having the most significant impact on the aerobrake structural weight were identified.

Bush, Lance B.↗
NASA NTRS1992

Weight optimization of an aerobrake structural concept for a lunar transfer vehicle

An aerobrake structural concept for a lunar transfer vehicle was weight optimized through the use of the Taguchi design method, finite element analyses, and element sizing routines. Six design parameters were chosen to represent the aerobrake structural configuration. The design parameters included honeycomb core thickness, diameter-depth ratio, shape, material, number of concentric ring frames, and number of radial frames. Each parameter was assigned three levels. The aerobrake structural configuration with the minimum weight was 44 percent less than the average weight of all the remaining satisfactory experimental configurations. In addition, the results of this study have served to bolster the advocacy of the Taguchi method for aerospace vehicle design. Both reduced analysis time and an optimized design demonstrated the applicability of the Taguchi method to aerospace vehicle design.

Bush, Lance B.↗
NASA NTRSDate not supplied

Ballistic Lunar Transfers to Near Rectilinear Halo Orbit: Operational Considerations

This paper presents a study of ballistic lunar transfer (BLT) trajectories from Earth launch to insertion into a near rectilinear halo orbit (NRHO). BLTs have favorable properties for uncrewed launches to orbits in the vicinity of the Moon, such as dramatically reduced spacecraft ΔV requirements. Results are described from a detailed set of related mission design studies: BLTs with and without an outbound lunar flyby, insertion and rendezvous of a single spacecraft with a target in an NRHO, and insertion and rendezvous of multiple spacecraft with a target NRHO in quick succession. Monte Carlo analyses are presented of simulated rendezvous operations with realistic errors. These analyses are presented to inform future missions to NRHOs.

Nathan L Parrish↗
NASA NTRS1993

Heating analysis for a Lunar Transfer Vehicle at near-equilibrium flow conditions

A heating analysis for a 15.2 m diameter Lunar Transfer Vehicle (LTV) at 0 and 10.6 deg angle of attack for a nominal trajectory through the earth's atmosphere is described. The analysis utilizes the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) with thin-layer, Navier-Stokes, thermochemical nonequilibrium options. Radiative heating levels are calculated using the Langley Optimized RAdiative Nonequilibrium (LORAN) and the Non-EQuilibrium AIr Radiation (NEQAIR) codes. At peak heating, the shock layer is substantially in equilibrium. Comprehensive spatial and spectral grid convergence studies have been implemented to quantify grid effects on the convective and radiative heating levels. Axisymmetric tests including the coupled effects of radiative energy transfer show negligible change to the convective heating and a 20 percent reduction in the radiative heating.

Gnoffo, Peter A.↗
NASA NTRS1991

A method for the construction of a lunar transfer trajectory using ballistic capture

The weak stability boundary (WSB) method for the design and optimization of lunar transfer trajectories is described. Considerable savings of propellant are shown over classical methods of orbit transfer such as the Hohmann transfer method. The savings in Delta V required of the spacecraft ranges from 100 to 200 m/s, which translates into a 5-10 percent reduction in spacecraft propellant for a science payload. Another advantage of the WSB method involves the utilization of low thrust propellant systems. Since the transfer trajectory is nearly completely ballistic, thrusting propulsive maneuvers may be performed over a long time duration.

Miller, James K.↗
NASA NTRS1993

Minimum accommodation for aerobrake assembly. Phase 2: Structural concepts for a lunar transfer vehicle aerobrake which can be assembled on orbit

A multidisciplinary conceptual study was conducted to define a reusable lunar transfer vehicle (LTV) aerobrake which could be launched on a Space Shuttle of Titan 4 and assembled on orbit at Space Station Freedom. A major objective was to design an aerobrake, with integrated structure and thermal protection systems, which has a mass less than 20 percent (9040 lb) of the LTV lunar return mass. The aerobrake segmentation concepts, the structural concepts, a joint concept for assembly, and a structural design with analysis of the aerobrake are described. Results show that a 50-foot diameter LTV aerobrake can be designed for on-orbit assembly which will achieve the 20 percent mass budget.

Dorsey, John T.↗
NASA NTRS1993

Structural concepts for a lunar transfer vehicle aerobrake which can be assembled on orbit

A multidisciplinary conceptual study was conducted to define a reusable Lunar transfer vehicle (LTV) aerobrake which could be launched on a Space Shuttle or Titan IV and assembled on orbit at Space Station Freedom. A major objective of this study was to design an aerobrake, with integrated structure and thermal protection systems, which has a mass less than 20 percent (9040 lb) of the LTV lunar return mass. This paper describes the aerobrake segmentation concepts, the structural concepts, a joint concept for assembly, and a structural design with analysis of the aerobrake. Results show that a 50-foot diameter LTV aerobrake can be designed for on-orbit assembly which will achieve the 20 percent mass budget.

Dorsey, John T.↗
NASA NTRS2017

Low Thrust Cis-Lunar Transfers Using a 40 kW-Class Solar Electric Propulsion Spacecraft

This paper captures trajectory analysis of a representative low thrust, high power Solar Electric Propulsion (SEP) vehicle to move a mass around cis-lunar space in the range of 20 to 40 kW power to the Electric Propulsion (EP) system. These cis-lunar transfers depart from a selected Near Rectilinear Halo Orbit (NRHO) and target other cis-lunar orbits. The NRHO cannot be characterized in the classical two-body dynamics more familiar in the human spaceflight community, and the use of low thrust orbit transfers provides unique analysis challenges. Among the target orbit destinations documented in this paper are transfers between a Southern and Northern NRHO, transfers between the NRHO and a Distant Retrograde Orbit (DRO) and a transfer between the NRHO and two different Earth Moon Lagrange Point 2 (EML2) Halo orbits. Because many different NRHOs and EML2 halo orbits exist, simplifying assumptions rely on previous analysis of orbits that meet current abort and communication requirements for human mission planning. Investigation is done into the sensitivities of these low thrust transfers to EP system power. Additionally, the impact of the Thrust to Weight ratio of these low thrust SEP systems and the ability to transit between these unique orbits are investigated.

cislunar↗
NASA NTRS2003

Lunar transfer trajectory design and the four-body problem

The existence of a ballistic trajectory from the Earth to orbit about the Moon was long considered to be impossible based on analysis of the three-body problem. In 1990 a ballistic trajectory from the Earth to lunar orbit was discovered while analyzing a plan to salvage the Muses A (Hiten) spacecraft. This trajectory utilized the Sun's gravity in conjunction with the Earth and Moon's gravity and was thus the first example of a practical four-body trajectory design. This paper presents a review of lunar transfer trajectories that go beyond three-body theory and the Jacobi integral. These include Hiten, Lunar A and the Genesis return trajectory from the vicinity of the Moon to Earth.It is shown that these trajectories may be analyzed by piecing together segments where three-body motion dominates.

Muses↗
NASA NTRS1992

An investigation into critical aspects of a new form of low energy lunar transfer, the Belbruno-Miller trajectories

This paper documents an integrated Belbruno-Miller (B-M) trajectory and its corresponding injection period. The B-M trajectories use Weak Stability Boundaries (WSB) resulting from four-body perturbative dynamics between the earth, moon, sun, and spacecraft to significantly reduce maneuver requirements for lunar transfer. It is determined that the presented nominal trajectory has a viable 4-day injection period considering maneuver constraints. The nominal B-M trajectory produces saving varying from 150 m/s to 222 m/s over traditional Hohmann means thus pointing to the practical utility of the WSB method. Energy savings are extracted at the price of extended flight time of approximately 6-months. Usefulness of this procedure was recently dramatized by the Japanese spacecraft Hiten when it arrived a the moon on October 2, 1991 following its entry into B-M trajectory on April 25, 1991.

Krish, V.↗
NASA NTRS2010

Concept for a Lunar Transfer Vehicle for Small Satellite Delivery to the Moon from the International Space Station

The International Space Station (ISS) has developed as a very capable center for scientific research in Lower Earth Orbit. An additional potential of the ISS that has not thus far been exploited, is the use of this orbiting plat-form for the assembly and launching of vehicles that could be sent to more distant destinations. This paper reports the results of a recent study that looked at an architecture and conceptual flight system design for a lunar transfer vehicle (LTV) that could be delivered to the ISS in segments, assembled, loaded with payload and launched from the ISS with the objective of delivering multiple small and micro satellites to lunar orbit. The design of the LTV was optimized for low cost and high payload capability, as well as ease of assembly. The resulting design would use solar electric propulsion (SEP) to carry a total payload mass of ~250 kg from the ISS to a 100 km lunar orbit. A preliminary concept of operations was developed considering currently available delivery options and ISS capabili-ties that should prove flexible enough to accommodate a variety of payloads and missions. This paper will present an overview of the study, including key trades, mission and flight system design, and notional operational concept.

Elliott, John↗
NASA NTRSDate not supplied

Overview of the Lunar Transfer Trajectory of the Co-Manifested First Elements of NASA’s Gateway

This paper documents the current design reference mission planned for the first two elements of NASA’s Gateway. When launched together, the Power and Pro-pulsion Element and Habitation and Logistics Outpost comprise the Co-Manifested Vehicle (CMV).1The low-thrust transfer between the initial parking orbit and the final insertion into the operational Near Rectilinear Halo Orbit is described. While each specific trajectory depends on launch date, trends are identified in the dynamics and orientation of the CMV as it traverses its spiral orbit. This paper describes the interplay between various assumptions and constraints on the development of the low thrust lunar transfer.

low thrust↗
NASA NTRSDate not supplied

Overview of the Lunar Transfer Trajectory of the Co-Manifested First Elements of NASA’s Gateway

This paper documents the current design reference mission planned for the first two elements of NASA’s Gateway. When launched together, the Power and Propulsion Element and Habitation and Logistics Outpost comprise the Co-Manifested Vehicle (CMV). The low-thrust transfer between the initial parking orbit and the final insertion into the operational Near Rectilinear Halo Orbit is described. While each specific trajectory depends on launch date, trends are iden-tified in the dynamics and orientation of the CMV as it traverses its spiral orbit. This paper describes the interplay between various assumptions and constraints on the development of the low thrust lunar transfer.

electric propulsion↗