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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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NASA NTRSDate not supplied

Trajectory Design Considerations for Low Lunar Orbit to Near Rectilinear Halo Orbit Transfers

NASA’s Artemis program plans to use a Near-Rectilinear Halo Orbit (NRHO) to stage assets for a lunar exploration campaign. While the proximity operations in the NRHO are simple and very similar to flat-space dynamics, the far-field ren-dezvous has considerations that are non-intuitive, especially to those that may be steeped in the dynamics of rendezvous in Keplerian orbits. This paper will discuss some key drivers to the delta-V performance of far-field rendezvous from Low Lunar Orbit to the Gateway space station in its 9:2 resonant south-ward NRHO.

NRHO↗
NASA NTRS2019

HelioSwarm: Swarm Mission Design in High Altitude Orbit for Heliophysics

Resolving the complex three-dimensional turbulent structures that characterize the solar wind requires contemporaneous spatially and temporally distributed measurements. HelioSwarm is a mission concept that will deploy multiple, co-orbiting satellites to use the solar wind as a natural laboratory for understanding the fundamental, universal process of plasma turbulence. The HelioSwarm transfer trajectory and science orbit use a lunar gravity assist to deliver the ESPA-class nodes attached to a large data transfer hub to a P/2 lunar resonant orbit. Once deployed in the science orbit, the free-flying, propulsive nodes use simple Cartesian relative motion patterns to establish baseline separations both along and across the solar wind flow direction.

lunar resonant orbit↗
NASA NTRS2019

HelioSwarm: Swarm Mission Design in High Altitude Orbit for Heliophysics

Resolving the complex three-dimensional turbulent structures that characterize the solar wind requires contemporaneous spatially and temporally distributed measurements. HelioSwarm is a mission concept that will deploy multiple, co-orbiting satellites to use the solar wind as a natural laboratory for understanding the fundamental, universal process of plasma turbulence. The HelioSwarm transfer trajectory and science orbit use a lunar gravity assist to deliver the ESPA-class nodes attached to a large data transfer hub to a P/2 lunar resonant orbit. Once deployed in the science orbit, the free-flying, propulsive nodes use simple Cartesian relative motion patterns to establish baseline separations both along and across the solar wind flow direction.

lunar resonant orbit↗
NASA NTRS2011

Targeting Low-Energy Transfers to Low Lunar Orbit

A targeting scheme is presented to build trajectories from a specified Earth parking orbit to a specified low lunar orbit via a low-energy transfer and up to two maneuvers. The total transfer delta V (velocity) is characterized as a function of the Earth parking orbit inclination and the departure date for transfers to each given low lunar orbit. The transfer delta V (velocity) cost is characterized for transfers constructed to low lunar polar orbits with any longitude of ascending node and for transfers that arrive at the Moon at any given time during a month.

mission designs↗
NASA NTRSDate not supplied

Overview of the Lunar Transit Trajectory Performed by the Power and Propulsion Element of NASA’s Gateway

NASA has committed to returning to the moon, landing the first woman and the next man on its surface. To support a sustained lunar presence, NASA is designing an orbital platform to be assembled in an orbit near the moon called the Near Rectilinear Halo Orbit (NRHO). This platform is known as the Gateway and its purpose it to support missions primarily to the lunar south pole. As NASA continues to study ways to reduce the cost of lunar exploration, a simplification implemented in 2020combinedthe first two elements of the Gateway, the Power and Propulsion Element (PPE) and NASA’s Habitation and Logistics Outpost (HALO), onto a single commercial launch vehicle (CLV). When launched together, the PPE and HALO make up the first two elements of NASA’s Gateway and exceed the performance capacity of commercially available launch vehicles to deliver directly to the moon. The delivery of the Gateway is enabled by and takes advantage of the high efficiency of the PPEs high-power Solar Electric Propulsion (SEP) system to transfer a significant starting mass from an initial Earth orbit to final insertion into the NRHO. The PPE is a 50-kW class high power solar electric propulsion stage comprised of two different types of electric thruster strings. The SEP system is operated in two different modes, a high thrust and high Isp (specific impulse), to both maximize the final delivered mass and attempt to minimize the time spent in the Van Allen Belts early in the transit trajectory. Additionally, this SEP system brings the capability to the assembled Gateway for transfer between orbits in cislunar space. This paper captures the preliminary low thrust lunar transfer reference trajectory to be flown by the PPE it delivers itself and HALO via a spiral trajectory from launch vehicle insertion to insertion into the NRHO as well as a preliminary reference round trip transfer of the Gateway from the NRHO to a Distant Retrograde Orbit (DRO) using the PPE. Once in its final NRHO, the Gateway is being designed to enable long duration human and robotic exploration of the lunar south pole as a precursor to Mars. The Gateway’s low thrust lunar transit trajectory is envisioned to prove out the application of and flight of a high-power SEP system as a demonstration of technologies applicable to the enabling of future human space exploration.

human exploration↗
NASA NTRSDate not supplied

Rocket Lab Venus - Enabling Low-Cost Interplanetary Missions

Rocket Lab’s mission to Venus [1], launching in January 2025, aims to demonstrate that small launch vehicles, such as Electron, and high-energy small spacecraft, such as Photon, can enable a new paradigm of regular, low-cost interplanetary missions. The Electron launch vehicle will deliver the Photon spacecraft carrying the entry probe into low earth orbit, at which point Photon will take over and put both itself and the attached probe on an Earth-Venus transfer trajectory, with assistance from a Lunar flyby. The primary science objective of the mission is to sample the Venusian cloud layers between atitudes of 45 and 60 km above the surface with an autofluorescing nephelometer. The ~1kg scientific instrument will be carried by a small ~17 kg direct entry probe which will separate from the Photon spacecraft approximately 30 minutes prior to atmospheric entry. The probe outer mould line is a scaled down version of the Deep Space 2 probes [2]. NASA’s newly developed 3D Woven Carbon Phenolic, a derivative of the Heatshield for Extreme Entry Environment Technology (HEEET) material [3], is an enabling technology for this mission, and it will comprise the probe forebody heat shield.The probe will communicate directly back to the Earth during its descent via a small on-board radio frequency (RF) antenna. The hyperbolic Venus encounter will see the probe enter the Venusia atmosphere at approximately 11 km/s with an entry flight path angle (EFPA) of about 10°. This relatively low entry angle was selected to increase the duration of the descent phase, reduce the peak deceleration, and reduce the sensitivity of the nominal entry environments to navigational errors. For reference, the Pioneer Venus probes entered at EFPAs varying from about 25° to 70°, resulting in peak decelerations between around 220 and 450 g’s and peak stagnation point heat fluxes ranging from about 40 to 75 MW/m2 [4]. In comparison, the Rocket Lab Venus probe will experience a lower preak deceleration of around 60 g’s and a lower peak stagnation point heat flux of about 15 MW/m2, at the cost of a greater total heat load and correspondingly higher thermal protection system (TPS) mass fraction. This paper recounts the development history of the TPS for the Rocket Lab Venus probe, with a focus on the unique requirements of the mission. In particular, the effect on the TPS design of the mission’s tight budget, short development timeline, slim mass and RF link margins, and mild but long duration entry heating. Rocket Lab’s initial investigations into a ‘hybrid’ TPS [5] based on a scaled version of the Pioneer Venus full-density carbon phenolic forebody shield, and the abandonment of this approach in favour of NASA’s 3D Woven Carbon Phenolic is presented. Additionally, the leveraging of previous testing under the HEEET project at higher conditions and the associated TPS margining strategy that was adopted to avoid the need for extensive mission-specific arc jet testing is discussed. Finally, this paper concludes with a brief consideration of the feasibility of a future low-cost aerocapture technology demonstrator mission based on the 2025 Rocket Lab Venus mission and entry probe.

Lyle Campbell↗
NASA NTRS1997

Lunar Prospector Mission Design

The National Aeronautics and Space Administration (NASA) has selected Lunar Prospector as one of the discovery missions to conduct solar system exploration science investigations. The mission is NASA's first lunar voyage to investigate key science objectives since Apollo and will be launched in September 1997. In keeping with discovery program requirements to reduce total mission cost and utilize new technology, Lunar Prospector's mission design and control will focus on the use of innovative and proven trajectory analysis programs. As part of this effort, the Ames Research Center and the Goddard Space Flight Center have become partners in the Lunar Prospector trajectory team to provide the trajectory analysis and orbit determination support. The mission design of Lunar Prospector is characterized by a direct minimum energy transfer trajectory to the moon with an injection into a circular 100 km altitude polar mapping orbit. The mission design will incorporate the optimization of the reference trajectory for a new launch vehicle. An extended mission includes a periselene altitude approaching 10 km. As part of the navigation, a new lunar potential model will be used that incorporates data from the recent Clementine mission. This paper provides understanding of the design of the Lunar Prospector mission and the optimization process of the nominal trajectory, orbit determination, launch windows, and lunar orbit targets, and discusses the feasibility of meeting the mission goals.

Folta, David↗
NASA NTRS1998

Lunar prospector mission design and trajectory support

The Lunar Prospector mission is the first dedicated NASA lunar mapping mission since the Apollo Orbiter program which was flown over 25 years ago. Competitively selected under the NASA Discovery Program, Lunar Prospector was launched on January 7, 1998 on the new Lockheed Martin Athena 2 launch vehicle. The mission design of Lunar Prospector is characterized by a direct minimum energy transfer trajectory to the moon with three scheduled orbit correction maneuvers to remove launch and cislunar injection errors prior to lunar insertion. At lunar encounter, a series of three lunar orbit insertion maneuvers and a small circularization burn were executed to achieve a 100 km altitude polar mapping orbit. This paper will present the design of the Lunar Prospector transfer, lunar insertion and mapping orbits, including maneuver and orbit determination strategies in the context of mission goals and constraints. Contingency plans for handling transfer orbit injection and lunar orbit insertion anomalies are also summarized. Actual flight operations results are discussed and compared to pre-launch support analysis.

Lozier, David↗
NASA NTRS1994

High Earth orbit design for lunar assisted small Explorer class missions

Small Expendable launch vehicles are capable of injecting modest payloads into high Earth orbits having apogee near the lunar distance. However, lunar and solar perturbations can quickly lower perigee and cause premature reentry. Costly perigee raising maneuvers by the spacecraft are required to maintain the orbit. In addition, the range of inclinations achievable is limited to those of launch sites unless costly spacecraft maneuvers are performed. This study investigates the use of a lunar swingby in a near-Hohmann transfer trajectory to raise perigee into the 8 to 25 solar radius range and reach a wide variety of inclinations without spacecraft maneuvers. It is found that extremely stable orbits can be obtained if the postencounter spacecraft orbital period is one-half of a lunar sidereal revolution and the Earth-vehicle-Moon geometry is within a specified range. Criteria for achieving stable orbits with various perigee heights and ecliptic inclinations are developed, and the sensitivity of the resulting mission orbits to transfer trajectory injection (TTI) errors is examined. It is shown that carefully designed orbits yield lifetimes of several years, with excellent ground station coverage characteristics and minimal eclipses. A phasing loop error correction strategy is considered with the spacecraft propulsion system delta V demand for TTI error correction and a postlunar encounter apogee trim maneuver typically in the 30 to 120 meters per second range.

Mathews, M.↗
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 NTRS1992

Transfer trajectory design for the SOHO libration-point mission

It is shown that transfer trajectories to the halo orbit exist throughout the year, with a one week closure of the launch window each month because of unfavorable lunar perturbations that cannot be corrected in the case of transfer trajectory insertion errors. The launch window for transfers to large-amplitude Lissajous orbits is virtually the same as that for transfers to the baseline halo orbit. Details of these trajectories are described and questions about groundstation coverage are discussed.

Dunham, D. W.↗
NASA NTRS1999

Optimization of Insertion Cost for Transfer Trajectories to Libration Point Orbits

The objective of this work is the development of efficient techniques to optimize the cost associated with transfer trajectories to libration point orbits in the Sun-Earth-Moon four body problem, that may include lunar gravity assists. Initially, dynamical systems theory is used to determine invariant manifolds associated with the desired libration point orbit. These manifolds are employed to produce an initial approximation to the transfer trajectory. Specific trajectory requirements such as, transfer injection constraints, inclusion of phasing loops, and targeting of a specified state on the manifold are then incorporated into the design of the transfer trajectory. A two level differential corrections process is used to produce a fully continuous trajectory that satisfies the design constraints, and includes appropriate lunar and solar gravitational models. Based on this methodology, and using the manifold structure from dynamical systems theory, a technique is presented to optimize the cost associated with insertion onto a specified libration point orbit.

Howell, K. C.↗
NASA NTRS2001

Conceptual Design and Analysis of an MXER Tether Boost Station

Momentum-exchange / electrodynamic reboost (MXER) tether systems show great promise for use in propellantless orbital transfer. Work conducted at MSFC (Marshall Space Flight Center) and TUI (Tethers Unlimited, Inc.) over the last two years has updated and refined a tether conceptual design that could be used to boost payloads from LEO (Low Earth Orbit) to GTO (Geostationary Transfer Orbit) or on a trans-lunar trajectory. Efficient configurations have been created that can be packaged, assembled, deployed, and maintained, improving their operational potential and flexibility.

Sorensen, Kirk↗
NASA NTRS2007

Optimization of Return Trajectories for Orbital Transfer Vehicle between Earth and Moon

In this paper, optimum trajectories in Earth Transfer Orbit (ETO) for a lunar transportation system are proposed. This paper aims at improving the payload ratio of the reusable orbital transfer vehicle (OTV), which transports the payload from Low Earth Orbit (LEO) to Lunar Low Orbit (LLO) and returns to LEO. In ETO, we discuss ballistic flight using chemical propulsion, multi-impulse flight using electrical propulsion, and aero-assisted flight using aero-brake. The feasibility of the OTV is considered.

Funase, Ryu↗
NASA NTRS1993

An interactive tool for design and support of lunar, gravity assist, and libration point trajectories

The 'Swingby' interactive tool for design, analysis, and support of maneuver planning for missions that involve transfers to the moon, the lunar orbits, multiple-gravity assists, and the libration-point orbits, runs on a PC and uses pull-down and pop-up menus to allow users to change physical constants and select perturbations for inclusion. Swingby can also target by ascertain parameters, or allow the computer to generate goals on the basis of mission requirements. Rapid mission analyses, operational maneuvers, and contingency plans are thereby obtainable.

Carrico, J.↗
NASA NTRS2021

Multi-revolution Extension of Solar-perturbed Moon-to-moon Transfer Families

Lunar flybys and solar perturbations present a golden opportunity to naturally modify trajectories of spacecraft launched in Moon-bound orbits. Leveraging the energy gained, or lost, from lunar flybys and multi-body effects allows for spacecraft to reach destinations not previously within their designed ?v budgets. Previous research studies have demonstrated the viability of collecting zero revolution Moon-to-Moon transfers into a referential database. This paper extends this capability to multi-revolution transfers and describes their main characteristics and increased versatility. Finally, a trajectory example is given for the escape phase of the NEA Scout mission.

Lantoine, Gregory↗
NASA NTRS1993

Trajectory design for the Deep Space Program Science Experiment (DSPSE) mission

In 1994, the Deep Space Program Science Experiment (DSPSE) spacecraft will become the first spacecraft to perform, in succession, both a lunar orbiting mission and a deep-space asteroid encounter mission. The primary mission objective is to perform a long-duration flight-test of various new-technology lightweight components, such as sensors, in a deep-space environment. The mission has two secondary science objectives: to provide high-resolution imaging of the entire lunar surface for mapping purposes and flyby imaging of the asteroid 1620 Geographos. The DSPSE mission is sponsored by the Strategic Defense Initiative Organization (SDIO). As prime contractor, the Naval Research Laboratory (NRL) is building the spacecraft and will conduct mission operations. The Goddard Space Flight Center's (GSFC) Flight Dynamics Division is supporting NRL in the areas of The Deep Space Network (DSN) will provide tracking support. The DSPSE mission will begin with a launch from the Western Test Range in late January 1994. Following a minimum 1.5-day stay in a low-Earth parking orbit, a solid kick motor burn will boost DSPSE into an 18-day, 2.5-revolution phasing orbit transfer trajectory to the Moon. Two burns to insert DSPSE into a lunar polar orbit suitable for the mapping mission will be followed by mapping orbit maintenance and adjustment operations over a period of 2 sidereal months. In May 1994, a lunar orbit departure maneuver, in conjunction with a lunar swingby 26 days later, will propel DSPSE onto a heliocentric transfer that will intercept Geographos on September 1, 1994. This paper presents the characteristics, deterministic delta-Vs, and design details of each trajectory phase of this unique mission, together with the requirements, constraints, and design considerations to which each phase is subject. Numerous trajectory plots and tables of significant trajectory events are included. Following a discussion of the results of a preliminary launch window analysis, a summary of the deterministic impulsive delta-V budget required to establish the baseline mission trajectory design is presented.

Carrington, D.↗