Genesis Mission Design
The Genesis spacecraft will collect solar wind samples from a halo orbit about the Sun-Earth L1 point for two years, returning those samples to Earth in 2003 for on-Earth analysis and examination.
Genesis↗
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Results for “halo orbit”
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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Designing Low-Thrust Enabled Trajectories for a Heliophysics Smallsat Mission to Sun-Earth L5
A small satellite deployed to Sun-Earth L5 could serve as a low-cost platform to observe solar phenomena such as coronal mass ejections. However, the small satellite platform introduces significant challenges in the trajectory design process via limited thrusting capabilities, power and operational constraints, and fixed deployment conditions. To address these challenges, a strategy employing dynamical systems theory is used to design a low-thrust-enabled trajectory for a small satellite to reach the Sun-Earth L5 region. This procedure is demonstrated for a small satellite that launches as a secondary payload with a larger spacecraft destined for a Sun-Earth L2 halo orbit.
Stuart, Jeffrey↗
ExEP Yield Modeling Tool and Validation Test Results
EXOSIMS is an open-source simulation tool for parametric modeling of the detection yield and characterization of exoplanets. EXOSIMS has been adopted by the Exoplanet Exploration Programs Standards De nition and Evaluation Team (ExSDET) as a common mechanism for comparison of exoplanet mission concept studies. To ensure trustworthiness of the tool, we developed a validation test plan that leverages the Python-language unit-test framework, utilizes integration tests for selected module interactions, and performs end-to-end cross- validation with other yield tools. This paper presents the test methods and results, with the physics-based tests such as photometry and integration time calculation treated in detail and the functional tests treated summarily. The test case utilized a 4m unobscured telescope with an idealized coronagraph and an exoplanet population from the IPAC radial velocity (RV) exoplanet catalog. The known RV planets were set at quadrature to allow deterministic validation of the calculation of physical parameters, such as working angle, photon counts and integration time. The observing keepout region was tested by generating plots and movies of the targets and the keepout zone over a year. Although the keepout integration test required the interpretation of a user, the test revealed problems in the L2 halo orbit and the parameterization of keepout applied to some solar system bodies, which the development team was able to address. The validation testing of EXOSIMS was performed iteratively with the developers of EXOSIMS and resulted in a more robust, stable, and trustworthy tool that the exoplanet community can use to simulate exoplanet direct-detection missions from probe class, to WFIRST, up to large mission concepts such as HabEx and LUVOIR.
Nunez, Paul↗
Nancy Grace Roman Space Telescope Observatory Implementation and Challenges
NASA’s Nancy Grace Roman Space Telescope (Roman), previously referred to as Wide Field Infrared Survey Telescope (WFIRST), was named after Dr. Nancy Grace Roman, an astronomer and NASA pioneer of modern space-based astronomy who is known as the “mother of the Hubble Space Telescope”. Roman is a deep space infrared observatory with a Hubble-sized telescope and wide field of view instrument (greater than 100 times that of Hubble’s) that will conduct a high latitude time-domain survey, a high latitude imaging and spectroscopic survey, and a galactic bulge time-domain survey to characterize dark energy and expand the census of exoplanets in our galaxy while allowing a broad range of astrophysics research. Roman will also demonstrate exoplanet coronagraphy with active wave front control technology and provide general investigator programs for the science community. Roman is finishing the critical design phase and is planning for launch in 2026. It will operate in a quasi-halo orbit about Sun-Earth L2, 1.5 million kilometers from Earth, for a five-year primary mission life. The Observatory features an Optical Telescope Assembly with an existing, repurposed 2.4m primary mirror, a Wide Field Instrument with a focal plane array comprised of 18 HgCdTe near-infrared detectors and a grism, prism and filter elements for imaging and spectroscopy in support of the primary surveys, as well as a Coronagraph instrument technology demonstration with starlight suppression technology for direct imaging and spectroscopy of exoplanets. The telescope is mounted to the Instrument Carrier composite truss structure which also optically meters each instrument, includes a Launch Load and Vibration Isolation System to provide passive isolation of spacecraft jitter sources while also supporting the payload during launch and is attached to the Spacecraft Bus. The Spacecraft also includes a Solar Array Sunshield ,Deployable Aperture Cover, Lower Instrument Sunshade, High Gain Antenna System, and Outer Barrel Assembly. Figure 1 shows an overview of the Roman Observatory. When fully integrated, Roman will be the largest Observatory assembled and tested at NASA’s Goddard Space Flight Center. Figure 1. Roman Observatory Overview Development of scientific satellites is challenging by nature, as the pursuit to broaden scientific knowledge always pushes the boundary of what has come before. The implementation of the Roman mission is a prime example and expected challenges have been augmented by the foundational decision to use the existing telescope components, developed in the early 2000s by another Government agency for a different application. Other unique aspects of the Roman mission, such as its survey nature, the vast amount of data required to meet science objectives, and packaging of the Observatory elements around the existing telescope components, create constrained design spaces that drive competing requirements across Observatory subsystems. Given these challenges, systems engineering has been a critical discipline in balancing implementation decisions for the Roman mission and will continue to play a key role going forward. This paper will discuss details of the Roman Observatory configuration, as well as some of the systems engineering challenges and the decision-making process used to mature the Roman Space Telescope preliminary design to implementation.
Lisa Ml Bartusek↗
Mission Design Considerations for Robotic Lunar and Gateway Payload Return
A selection of lunar return trajectories is examined and assessed in terms of payload mass, vehicle mass, mission time, mission complexity, and total delta-V using a range of assumptions for the mission design based on historical precedence and near-future vehicle availability. Direct surface return trajectories using a single burn solution are compared with a range of Near-Rectilinear Halo Orbit return and Ballistic Return Trajectories. This study weighs anticipated mission needs, requirements and constraints with the aim of assessing the feasibility of commercial lunar cargo and/or sample return options utilizing NASA’s Gateway and Deep Space Logistics project.
lunar↗
Thermal Analysis of NRHO Entry Lunar Flyby Maneuver
NASA’s Gateway will serve as a platform for sustained human lunar exploration and an opportunity to gain experience with long duration deep-space architectures. The first two Gateway modules, the Power and Propulsion Element (PPE) and the Habitation and Logistics Outpost (HALO) are scheduled to launch together as a Co-Manifested Vehicle (CMV) in 2024. To reach the destination Near-Rectilinear Halo Orbit (NRHO), the CMV will perform a low-thrust electric propulsion spiral and a final transfer maneuver that will include a low altitude flyby of the lunar surface. This flyby thermal environment is more adverse than that expected during the NRHO itself and risks temperature exceedances on various spacecraft components. The highly transient nature of the flyby also required a different approach than the typical worst-case dissipation margins used to size the spacecraft thermal control systems. This paper presents analyses on the relative severity of different minimum perilune altitudes and results of an integrated CMV thermal model built in Thermal Desktop.
lunar flyby maneuver↗
Equuleus Launch Window Analysis and Mission Design
This paper presents the trajectory design process for EQUULEUS, a 6U CubeSat developed by JAXA and the University of Tokyo that is scheduled to launch as a piggyback of NASA’s Artemis 1. After separation from the upper stage of the Space Launch System, EQUULEUS will maneuver along a low-energy transfer to an Earth–Moon quasi-rectilinear halo orbit in 1-to-4 resonance with the lunar synodic period. As a secondary payload, the trajectory of EQUULEUS needs to be compatible with the requirements of the primary mission, but also robust against disturbances and potential changes in the deployment state. Realistic initial conditions spanning two years of potential launch windows are processed and the solution structure for optimal lunar transfers is analyzed. A host of candidate solutions is presented, compatibly with the fuel and power limitations of EQUULEUS. The global understanding of the solution space is shown to be insightful for the design of robust trajectories for limited control-authority spacecraft.
Kawakatsu, Yasuhiro↗
JWST REAL-TIME MID-COURSE CORRECTION MANEUVER MONITORING CONTINGENCY PREPARATION
The NASA James Webb Space Telescope (JWST) mission successfully launched on Dec 25, 2021, at 12:20 Coordinated Universal Time (UTC). During the 30-day transfer to the second Sun-Earth-Moon (SEM) libration point (L2), JWST executed three mid-course correction (MCC) maneuvers to insert into a quasi-halo orbit about L2. This paper covers the design and modeling for these three maneuvers with a focus on the timeline around the execution of each MCC maneuver. It will summarize the actual on-board events as well as the contingency preparation done for maneuver planning, monitoring, and final post-burn reconstruction of all three MCC maneuvers.
Wayne Yu↗
Flight Dynamics Planning and Operations Support for the JWST Mission
The James Webb Space Telescope (JWST) was launched from Kourou Spaceport on December 25, 2021, at 12:20 UTC on an Ariane 5 launch vehicle. The launch vehicle inserted JWST into a 30-day transfer trajectory to the Sun-Earth-Moon (SEM) Lagrange point L2 region. JWST executed three mid-course correction maneuvers (MCCs) to insert the spacecraft into a quasi-halo orbit about SEM L2; JWST will maintain its trajectory about L2 for at least 5.5 years, with a goal of at least 10.5 years. This paper summarizes the flight dynamics support for JWST, including the prelaunch nominal trajectory design, the launch window analysis, contingency planning for trajectory-related anomalies, mission operations support for the first 6 months, and a comparison of the planned and achieved actual JWST trajectory results. The orbit determination strategy, both planned and executed, will be summarized, and the method of addressing the anomalies as they occurred will be included.
Space Operations↗
Robust Trajectory Optimization and GN&C Performance Analysis for NRHO Rendezvous
This paper evaluates several candidate Near-Rectilinear Halo Orbits (NRHO) rendezvous trajectory designs using linear covariance (LinCov) analysis and determines the optimal locations for NRHO rendezvous translational maneuver locations. The performance of several candidate relative trajectory designs are determined as a function of relative navigation accuracy (angles only), inertial optical navigation (OpNav), range observability maneuvers, maneuver execution errors, relative maneuver targeting, and environment uncertainties. Further, the optimal locations of rendezvous maneuvers are determined for each of the candidate reference trajectories. The long-term goal of this research is to utilize LinCov and a genetic optimization algorithm (GA) to determine a complete end-to-end optimal NRHO trajectory design that is robust to navigation errors, maneuver execution errors, and environment uncertainties. This paper represents a first step toward this goal. Three candidate rendezvous trajectories with varying numbers of range-observability maneuvers are evaluated for their robustness to uncertainties, errors, and total trajectory correction delta-v performance. Some key elements of this analysis include relative navigation performance in an NRHO, relative trajectory dispersion performance, and total 3-sigma delta-v performance. This development provides the foundation to then determine an optimal and robust end-to-end NRHO rendezvous trajectory, including the determination of the optimal locations of range observability maneuvers, if needed.
Linear Covariance Analysis↗
JWST Real-Time Mid-Course Correction Maneuver Monitoring Contingency Preparation
The NASA James Webb Space Telescope (JWST) mission successfully launched on Dec 25, 2021, at 12:20 Coordinated Universal Time (UTC). During the 30-day transfer to the second Sun-Earth-Moon (SEM) libration point (L2), JWST executed three mid-course correction (MCC) maneuvers to insert into a quasi-halo orbit about L2. This paper covers the design and modeling for these three maneuvers with a focus on the timeline around the execution of each MCC maneuver. It will summarize the actual on-board events as well as the contingency preparation done for maneuver planning, monitoring, and final post-burn reconstruction of all three MCC maneuvers.
Toni Santacroce↗
Flight Dynamics Planning and Operations Support for the JWST Mission
The James Webb Space Telescope (JWST) was launched from Kourou Spaceport on December 25, 2021, at 12:20 UTC on an Ariane 5 launch vehicle. The launch vehicle inserted JWST into a 30-day transfer trajectory to the Sun-Earth-Moon (SEM) Lagrange point L2 region. JWST executed three mid-course correction maneuvers (MCCs) to insert the spacecraft into a quasi-halo orbit about SEM L2; JWST will maintain its trajectory about L2 for at least 5.5 years, with a goal of at least 10.5 years. This paper summarizes the flight dynamics support for JWST, including the prelaunch nominal trajectory design, the launch window analysis, contingency planning for trajectory-related anomalies, mission operations support for the first 6 months, and a comparison of the planned and achieved actual JWST trajectory results. The orbit determination strategy, both planned and executed, will be summarized, and the method of addressing the anomalies as they occurred will be included.
Karen Richon↗
Attitude Control Approach for Solar Cruiser, a Large, Deep Space Solar Sail
Solar Cruiser is a small satellite Technology Demonstration Mission (TDM) to mature solar sail propulsion technology using a solar sail larger than 1600 square meters, demonstrating performance as a propulsion system and a stable pointing platform for science observations in an artificial halo orbit sunward of the Sun-Earth Lagrange Point 1 (sub-L1). For the “sailcraft” to meet mission ob- jectives, there are several unique attitude control challenges that the Attitude De- termination and Control System (ADCS) must overcome. Large disturbance tor- ques, primarily due to sail deformations coupled with off-sun pointing angles, make it more difficult to maintain adequate controls performance and manage accumulated momentum on the control actuators. The large-amplitude, low- frequency flexible body modes of the sail, in concert with noisy sensors and ac- tuators, make it challenging to mitigate control-structure interactions and main- tain fine pointing capabilities. Stability and control performance during sail de- ployment is complicated by rapidly and widely varying inertias. The Solar Cruiser ADCS effectively addresses these challenges using a simple, traditional control system – including momentum management actuators that use bang- bang control to constrain internal accumulated momentum, a low-pass controls filter and an attitude determination Kalman Filter (KF) that blends multiple star tracker solutions, and a reaction wheel assembly (RWA) with controller gains tuned to specific configurations or inertias. Solar Cruiser’s attitude control ap- proach, and lessons learned from its development, establishes a state of the art of high value to future solar sail missions and other small spacecraft with large de- ployable structures operating in deep space.
solar sail↗
Attitude Control Approach for Solar Cruiser, a Large, Deep Space Solar Sail
Solar Cruiser is a small satellite Technology Demonstration Mission (TDM) to mature solar sail propulsion technology using a solar sail larger than 1600 square meters, demonstrating performance as a propulsion system and a stable pointing platform for science observations in an artificial halo orbit sunward of the Sun-Earth Lagrange Point 1 (sub-L1). For the “sailcraft” to meet mission objectives, there are several unique attitude control challenges that the Attitude Determination and Control System (ADCS) must overcome. Large disturbance torques, primarily due to sail deformations coupled with off-sun pointing angles, make it more difficult to maintain adequate controls performance and manage accumulated momentum on the control actuators. The large-amplitude, low-frequency flexible body modes of the sail, in concert with noisy sensors and actuators, make it challenging to mitigate control-structure interactions and maintain fine pointing capabilities. Stability and control performance during sail deployment is complicated by rapidly and widely varying inertias. The Solar Cruiser ADCS effectively addresses these challenges using a simple, traditional control system – including momentum management actuators that use bang-bang control to constrain internal accumulated momentum, a low-pass controls filter and an attitude determination Kalman Filter (KF) that blends multiple star tracker solutions, and a reaction wheel assembly (RWA) with controller gains tuned to specific configurations or inertias. Solar Cruiser’s attitude control approach, and lessons learned from its development, establishes a state of the art of high value to future solar sail missions and other small spacecraft with large deployable structures operating in deep space.
solar sail↗
Reflectivity Control Device (RCD) Roll Momentum Management for Solar Cruiser and Beyond
Reflectivity control devices (RCDs), thin film optical devices that are actuated (by application of a voltage across the device) to control the reflection of sunlight on their surfaces, can effectively enable roll momentum management (MM) of a solar sail operating in deep space, outside any magnetic field. This enables solar sails to maintain adequate attitude control in these challenging environments without sacrificing mission life by using expendables. They were planned and developed for use on Solar Cruiser, operating sunward of the Sun-Earth Lagrange Point 1 (sub-L1) in a propulsion-intensive artificial halo orbit. Located in pairs of panels near the distal end of each of four booms, each articulated at a slight angle out of the sail plane to produce a force tangent to the sail at a large moment arm, the RCDs provide roll torque capabilities by using simple switching logic to determine which panels should be activated (voltage applied) to produce torque in the desired direction. The performance capabilities and constraints of these devices across a range of operating conditions, namely sun-relative attitude, is well characterized, granting designers insight on how to properly size the system and optimize the system and component designs to maximize roll torque per unit area. With an informed design, RCDs can be very effective actuators for managing roll momentum with minimal detrimental system impacts (mass, power, complexity, etc.) on Solar Cruiser and other future deep space solar sail missions.
Andrew F Heaton↗
The NASA Solar Cruiser Solar Sail System – Ready for Heliophysics and Deep Space Missions
Mission Mature solar sail propulsion technology to enable near-term, compelling space missions for NASA and the nation. Mission Technology Goals Demonstrate solar sail propulsion technology to enable near- and midterm missions in cis-lunar space, sun-L1 halo orbits, non-Keplerian and other planetary orbits - Maneuver without regret capability in the Earth/Moon system - Space Domain Awareness - Increased Space Weather warning times - Sustained in-situ measurements within the Earth's magnetotail - High inclination solar orbital observations
Solar Sail↗
Momentum Management Strategies for Solar Cruiser and Beyond (ISSS 2023)
Solar Cruiser is a small (ESPA-class) satellite Technology Demonstration Mission (TDM) to mature solar sail propulsion technology using a solar sail larger than 1600 square meters, demonstrating performance both as a propulsion system and a stable pointing platform for science observations in an artificial halo orbit sunward of the Sun-Earth Lagrange Point 1 (sub-L1). To ensure attitude control throughout the mission, momentum accumulated on the reaction wheels (RWs) used for attitude control must be managed such that the sailcraft does not lose control due to RW momentum saturation. Momentum builds up on the wheels from environmental disturbance torques caused by solar radiation pressure combined with a center of mass (CM)/center of pressure (CP) offset, deformed sail shape, and an off-sun pointing angle, plus other factors. Solar Cruiser mitigates this momentum build up by utilizing an Active Mass Translator (AMT) that maintains pitch and yaw momentum by trimming the CM/CP offsets, and thrusters to maintain roll momentum. A survey was conducted by the Solar Cruiser team to assess the feasibility and tradeoffs of novel momentum management concepts such as Reflectivity Control Devices (RCD’s), different thruster configurations, and control vanes and other articulated control surfaces. In addition, techniques to reduce disturbance torque buildup, such as reducing boom tip deflections and clock angle control, were assessed. Similar sailcraft momentum management strategies can be used for future missions such as space weather monitoring and Earth magnetotail science missions
Solar Sail↗
Capability Gaps Assessment and Identification of Critical Technology Elements for Mars Transit Habitat
The Habitation Systems Development Office (HP40) at NASA Marshall Space Flight Center supports systems engineering, integration, and project management for next generation space habitats. For in space operations and eventual transport of humans to Mars, NASA will rely on a Mars Transit Habitat (TH). The TH will be designed for an up to 1,200-day Mars mission and will carry all food and supplies needed to support four crew for this duration. In the current concept of operations, Mars TH transfers to near rectilinear halo orbit (NRHO) following launch and docks at Gateway as a visiting vehicle. While there, the TH will complete system shakedown testing and a series of analog missions which will grow from 3 to 6+ months in duration TH also augments Gateway’s habitation capabilities beyond 60-days. Proposed Gateway-TH missions will far exceed the longest duration cislunar human missions to date. These shakedown missions will also be the first operational readiness tests of Mars TH’s long-duration deep space systems, and of the split crew (two crew on the surface, two crew in space) operations that are vital to the approach for the first human Mars mission. Once shakedown missions are complete, Mars TH departs Gateway to aggregate with the Mars propulsion system in NRHO before onboarding the crew and final supplies in Earth orbit via a co-manifested Orion-logistics module. Orion and the LM return to Earth prior to the now aggregated Deep Space Transport vehicle’s journey to Mars. Development of the Mars TH requires significant technology development and maturation. Each year the agency performs a capability gaps assessment, where gaps developed by subject matter experts (SMEs) in various engineering/science disciplines are linked to architectural elements in formulation and prioritized. A gap captures the difference between the current state-of-the-art and the maturity of the capability that is needed to enable or enhance a mission as it is currently envisioned in the government reference architecture. HP40 conducted a gap analysis for Mars TH which will be summarized in this poster. Gaps classified as enabling (which means the mission cannot achieve success without gap closure) were subsequently used to identify critical technology elements (CTEs) for Mars TH. This identification of CTEs was also informed by an examination of the product breakdown structure for Mars TH and focused conversations with SMEs in specific technology areas. CTEs identified for Mars TH to date include the following (note this is not a comprehensive list – CTEs listed represent those in MSFC’s capability areas): inflatable softgoods for habitation; enhanced CO2 recovery; life support systems with greater levels of reliability and maintainability; autonomous guidance, navigation, command and control; and radiators for the Mars TH application. The habitation systems development team is currently delving deeper into each CTE to assess technology approaches being pursued, their maturity, and the degree of difficulty in maturation to meet projected Mars TH timelines. This poster will summarize work to date on the identification of enabling capability gaps linked to Mars TH and provide insight into the associated CTEs and technology maturation efforts.
technology development↗