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NASA NTRS2011

Orion Entry Handling Qualities Assessments

The Orion Command Module (CM) is a capsule designed to bring crew back from the International Space Station (ISS), the moon and beyond. The atmospheric entry portion of the flight is deigned to be flown in autopilot mode for nominal situations. However, there exists the possibility for the crew to take over manual control in off-nominal situations. In these instances, the spacecraft must meet specific handling qualities criteria. To address these criteria two separate assessments of the Orion CM s entry Handling Qualities (HQ) were conducted at NASA s Johnson Space Center (JSC) using the Cooper-Harper scale (Cooper & Harper, 1969). These assessments were conducted in the summers of 2008 and 2010 using the Advanced NASA Technology Architecture for Exploration Studies (ANTARES) six degree of freedom, high fidelity Guidance, Navigation, and Control (GN&C) simulation. This paper will address the specifics of the handling qualities criteria, the vehicle configuration, the scenarios flown, the simulation background and setup, crew interfaces and displays, piloting techniques, ratings and crew comments, pre- and post-fight briefings, lessons learned and changes made to improve the overall system performance. The data collection tools, methods, data reduction and output reports will also be discussed. The objective of the 2008 entry HQ assessment was to evaluate the handling qualities of the CM during a lunar skip return. A lunar skip entry case was selected because it was considered the most demanding of all bank control scenarios. Even though skip entry is not planned to be flown manually, it was hypothesized that if a pilot could fly the harder skip entry case, then they could also fly a simpler loads managed or ballistic (constant bank rate command) entry scenario. In addition, with the evaluation set-up of multiple tasks within the entry case, handling qualities ratings collected in the evaluation could be used to assess other scenarios such as the constant bank angle maintenance case. The 2008 entry assessment was divided into two sections (see Figure 1). Entry I was the first, high speed portion of a lunar return and Entry II was the second, lower speed portion of a lunar return, which is similar (but not identical) to a typical ISS return.

Bihari, B.↗

NASA NTRS1986

Force/torque display for space teleoperation control experiments and evaluation

Experiments were performed at the Johnson Space Center (JSC), Manipulator Development Facility using the full scale Shuttle Remote Manipulator System (SRMS) to evaluate the effect of visual presentation through perspective display of the orthogonal forces and torques sensed at the manipulator end effector. The experiments investigated the effect of the display information on the management of forces and torques generated during payload berthing and deployment, as well as simulated satellite module change-out operations. The evaluation also addressed (1) issues of display format, including: force/torque scaling, point of resolution, and display mixing with video generated imagery, and (20) task related variables of payload size, alternative sources of guidance information, and control mode. The results are presented of a first-pass informal anlaysis of the analog, strip chart-recorded data from these evaluation tests. The results provide a relative measure of improvement in force management through the use of such a display, as well as information regarding the impact of display variables and task demands on operator performance.

Corker, K.↗

NASA NTRS2019

Design of Experiment for the Measurement of Aerosol Droplet Size Distribution of Temperature-Controlled Thermally Atomized Printed Electronic Inks

In-Space Manufacturing (ISM) centers around NASA’s growing need and ability to produce space technologies on demand in space. As the future of long term presence in space and deep space exploration approach, fundamental questions of our dependence on earth resupply to Low Earth Orbit (LEO) remain unanswered. ISM is leading various effort to evaluate the feasibility of producing essential spares and redundant parts on demand to enable a sustainable space-based supply chain model for part resupply. Among the parts and systems being considered, Avionics form the neural network of modern day aircraft and space vehicles providing a wealth of information ranging from Guidance, Navigation, and Control (GN&C) systems to on board Environmental Control and Life Support (ECLS) systems. Recent advances in the use of Aerosol Jet Technology to print Avionics components ranging from electrical traces on a circuit board to complex transistors and sensors raise the possibility of using such technology to reproduce or recreate electronic parts on demand with the help of custom electronics 3D printers. The challenge herein lies within the ability of such printers to generate and deposit an aerosol of electronic material utilizing processes independent of or enhanced by gravity to ensure controllably identical or improved behavior of the aerosol in an International Space Station (ISS) laboratory and on the ground. The behavior as well as the hazards and properties associated with such aerosols in a microgravity environment must be understood well in order to merit a feasible approach to utilizing them for manufacturing in space. In this report, we outline the experimental setup of a modified conventional vaping device to be used as the ideal gravity independent thermal atomization mechanism to generate aerosol. Our objective is to identify the ideal mass, density, and volume of our aerosolized droplets of ink to conclude that there exist a threshold of aerosolized ink droplet sizes that are indeed independent of the effects of gravity and remain stable after atomization. We use a Malvern Spraytec® Spray Particle Size Analyzer to perform real-time laser diffraction measurement of our ink droplets during atomization. The droplet size between conductive ink, dielectric ink and vegetable glycerin have been measured and contrasted. Furthermore, the mechanism of thermal atomization versus traditional pneumatic and ultrasonic atomization for operation in microgravity have been explored.

Wu, Chih-Hao↗

NASA NTRS2001

Achieving the Earth Science Enterprise Vision for the 21st Century: Platform Challenges

The ESE observational architecture of the future vision is dramatically different from that of today. The vision suggests observations from multiple orbits, collaborating space assets, and even seamless integration of space and other assets. Observations from GEO or from Libration points rather than from LEO suggest spacecraft carrying instruments with large deployable apertures. Minimization of launch costs suggests that these large apertures have long life, be extremely mass and volume efficient, and have low life cycle cost. Another significant challenge associated with high latitude orbits is high precision pointing and control. Finally, networks of spacecraft flying in predetermined constellation will be required either to apply complementary assets to an observation or to extend the virtual aperture beyond that attainable with a single spacecraft. These changes dictate development of new technology on several fronts, which are outlined in this paper. A section on high speed communications will outline requirements and approaches now envisioned. Sensorwebs will be developed from the viewpoint of work already begun for both space and for terrestrial networks. Precision guidance, navigation and control will be addressed from the perspective of precision flying for repeat pass interferometry and extreme pointing stability for advanced altimetry. A separate section will address requirements for distributed systems. Large lightweight deployables will be discussed with an emphasis on inflatable technology and its predicted benefits for large aperture instruments. For each technology area listed, current state-of-the-art, technological approaches for future development, and projected levels of performance are outlined.

Lemmerman, Loren↗

NASA NTRS2017

Spacecraft Alignment Determination and Control for Dual Spacecraft Precision Formation Flying

Many proposed formation flying missions seek to advance the state of the art in spacecraft science imaging by utilizing precision dual spacecraft formation flying to enable a virtual space telescope. Using precision dual spacecraft alignment, very long focal lengths can be achieved by locating the optics on one spacecraft and the detector on the other. Proposed science missions include astrophysics concepts with spacecraft separations from 1000 km to 25,000 km, such as the Milli-Arc-Second Structure Imager (MASSIM) and the New Worlds Observer, and Heliophysics concepts for solar coronagraphs and X-ray imaging with smaller separations (50m 500m). All of these proposed missions require advances in guidance, navigation, and control (GNC) for precision formation flying. In particular, very precise astrometric alignment control and estimation is required for precise inertial pointing of the virtual space telescope to enable science imaging orders of magnitude better than can be achieved with conventional single spacecraft instruments. This work develops design architectures, algorithms, and performance analysis of proposed GNC systems for precision dual spacecraft astrometric alignment. These systems employ a variety of GNC sensors and actuators, including laser-based alignment and ranging systems, optical imaging sensors (e.g. guide star telescope), inertial measurement units (IMU), as well as micro-thruster and precision stabilized platforms. A comprehensive GNC performance analysis is given for Heliophysics dual spacecraft PFF imaging mission concept.

Formation flying↗

NASA NTRS2017

Spacecraft Alignment Determination and Control for Dual Spacecraft Precision Formation Flying

Many proposed formation flying missions seek to advance the state of the art in spacecraft science imaging by utilizing precision dual spacecraft formation flying to enable a virtual space telescope. Using precision dual spacecraft alignment, very long focal lengths can be achieved by locating the optics on one spacecraft and the detector on the other. Proposed science missions include astrophysics concepts with spacecraft separations from 1000 km to 25,000 km, such as the Milli-Arc-Second Structure Imager (MASSIM) and the New Worlds Observer, and Heliophysics concepts for solar coronagraphs and X-ray imaging with smaller separations (50m-500m). All of these proposed missions require advances in guidance, navigation, and control (GNC) for precision formation flying. In particular, very precise astrometric alignment control and estimation is required for precise inertial pointing of the virtual space telescope to enable science imaging orders of magnitude better than can be achieved with conventional single spacecraft instruments. This work develops design architectures, algorithms, and performance analysis of proposed GNC systems for precision dual spacecraft astrometric alignment. These systems employ a variety of GNC sensors and actuators, including laser-based alignment and ranging systems, optical imaging sensors (e.g. guide star telescope), inertial measurement units (IMU), as well as microthruster and precision stabilized platforms. A comprehensive GNC performance analysis is given for Heliophysics dual spacecraft PFF imaging mission concept.

Virtual Telescope↗

NASA NTRS2018

Spacecraft Alignment Determination and Control for Dual Spacecraft Precision Formation Flying

Many proposed formation flying missions seek to advance the state of the art in spacecraft science imaging by utilizing dual-spacecraft precision formation flying (PFF) to enable a 'virtual' telescope (VT). Using precision dual-spacecraft alignment, very long focal lengths can be achieved by locating the optics on one spacecraft and the detector on the other. Proposed science missions include astrophysics concepts for X-ray imaging and exo-planet observation with large spacecraft separations (1000 km to 80,000 km), and heliophysics concepts for X-ray or extreme ultra-violet (EUV) imaging or solar coronagraphs with smaller separations (50m - 500m). These proposed missions require advances in guidance, navigation, and control (GN&C) for PFF to enable high resolution science imaging. For many applications, the dual-spacecraft dynamics are coupled through the GN&C system when the relative ranging and position alignment sensor components are not co-located with their respective spacecraft mass centers. We develop a model-based PFF system design approach for the VT application, considering the coupling inherent in precision dual-spacecraft inertial alignment. These systems employ a variety of GN&C sensors and actuators, including laser-based alignment and ranging systems, camera-based imaging sensors, inertial measurement units (IMU), as well as microthruster systems and image motion compensation platforms. Results of a GN&C performance assessment reveal how data from relative position sensors can be employed in a Kalman filter framework to significantly improve alignment estimation performance. The assessment provides a comparison of two different GN&C formation flying architectures, illustrating the performance trades inherent in the choice of PFF system architecture in the VT application.

Formation Flying↗

NASA NTRSDate not supplied

Integrated Design Results for the MSR DAC-0.0 Mars Ascent Vehicle

The NASA Mars Sample Return (MSR) Campaign endeavors to return Martian regolith, rock, and atmospheric samples to Earth for scientific study. One of many significant challenges to overcome in the return of these samples lies in transporting them from the Martian surface to space. In order to surmount this challenge, the Campaign has conceptualized the need for a Mars Ascent Vehicle (MAV) to perform this function and deliver Martian samples to orbit. There, the samples will be ejected and captured by a separate spacecraft for return to Earth. Many concepts for a MAV have existed in the past, but it has not been until now that an integrated, detailed design solution has been developed and analyzed. Preliminary assessments of the initial architecture examined multiple methods of propulsion. The team ultimately determined that a Two Stage to Orbit (TSTO) solid propulsion vehicle would provide the most effective performance and be the most technologically ready to support this mission. Following the decision to adopt a TSTO solid propelled vehicle, the first official Design Analysis Cycle, DAC-0.0, was performed in Spring 2020 to formally advance the fidelity of the vehicle to a maturity level acceptable for NASA Key Decision Point A (KDP-A). This paper describes the resultant MAV design concept developed as part of the DAC-0.0 study by the NASA Marshall Space Flight Center (MSFC), in association with the NASA Jet Propulsion Laboratory (JPL). The TSTO vehicle features two solid rocket motors, one powering each stage. Their thrust vectors are controlled with Thrust Vector Control (TVC) systems consisting of independent electromechanical actuators acting on gimballed nozzles. The vehicle is designed to deliver up to 0.47kg of Martian samples to a Mars circular orbit of 343km at 27° inclination. Due to the unique environmental conditions that this vehicle is required to operate in, the subsystem design teams were compelled to develop creative and unorthodox designs to ensure a successful mission. The detailed design and analysis of these subsystems are discussed in this paper and include topics on the MAV Guidance, Navigation, and Control (GNC); structures and mechanisms; integrated vehicle thermal; avionics and flight software; a hydrazine-based Reaction Control System (RCS); aerosciences; and vehicle assembly, integration, and test (AI&T) considerations, among others. Following the conclusion of the MAV DAC-0.0, additional alternative architecture concepts were also studied to further reduce the mass of the overall system. The results of these studies will also be examined in this paper.

Darius Yaghoubi↗

NASA NTRSDate not supplied

Integrated Design Results for the MSR DAC-0.0 Mars Ascent Vehicle

The NASA Mars Sample Return (MSR) Campaign endeavors to return Martian regolith, rock, and atmospheric samples to Earth for scientific study. One of many significant challenges to overcome in the return of thesesamples lies in transporting them from the Martian surface to space. In order to surmount this challenge, the Campaign has conceptualized the need for a Mars Ascent Vehicle (MAV) to perform this function and deliver Martian samples to orbit. There, the samples will be ejected and captured by a separate spacecraft for return to Earth. Many concepts for a MAV have existed in the past, but it has not been until now that an integrated, detailed design solution has been developed and analyzed. Preliminary assessments of the initial architecture examined multiple methods of propulsion. The team ultimately determined that a Two Stage to Orbit (TSTO) solid propulsion vehicle would provide the most effective performance and be the most technologically ready to support this mission. Following the decision to adopt a TSTO solid propelled vehicle, the first official Design Analysis Cycle, DAC-0.0, was performed in Spring 2020 to formally advance the fidelity of the vehicle to a maturity level acceptable for NASA Key Decision Point A (KDP-A). This paper describes the resultant MAV design concept developed as part of the DAC-0.0 study by the NASA Marshall Space Flight Center (MSFC), in association with the NASA Jet Propulsion Laboratory (JPL). The TSTO vehicle features two solid rocket motors, one powering each stage. Their thrust vectors are controlled with Thrust Vector Control (TVC) systems consisting of independent electromechanical actuators acting on gimballed nozzles. The vehicle is designed to deliver up to 0.47kg of Martian samples to a Mars circular orbit of 343km at 27° inclination. Due to the unique environmental conditions that this vehicle is required to operate in, the subsystem design teams were compelled to develop creative and unorthodox designs to ensure a successful mission. The detailed design and analysis of these subsystems are discussed in this paper and include topics on the MAV Guidance, Navigation, and Control (GNC); structures and mechanisms; integrated vehicle thermal; avionics and flight software; a hydrazine-based Reaction Control System (RCS); aerosciences; and vehicle assembly, integration, and test (AI&T) considerations, among others. Following the conclusion of the MAV DAC-0.0, additional alternative architecture concepts were also studied to further reduce the mass of the overall system.

Mars↗

NASA NTRS1992

Dynamic interactions between hypersonic vehicle aerodynamics and propulsion system performance

Described here is the development of a flexible simulation model for scramjet hypersonic propulsion systems. The primary goal is determination of sensitivity of the thrust vector and other system parameters to angle of attack changes of the vehicle. Such information is crucial in design and analysis of control system performance for hypersonic vehicles. The code is also intended to be a key element in carrying out dynamic interaction studies involving the influence of vehicle vibrations on propulsion system/control system coupling and flight stability. Simple models are employed to represent the various processes comprising the propulsion system. A method of characteristics (MOC) approach is used to solve the forebody and external nozzle flow fields. This results in a very fast computational algorithm capable of carrying out the vast number of simulation computations needed in guidance, stability, and control studies. The three-dimensional fore- and aft body (nozzle) geometry is characterized by the centerline profiles as represented by a series of coordinate points and body cross-section curvature. The engine module geometry is represented by an adjustable vertical grid to accommodate variations of the field parameters throughout the inlet and combustor. The scramjet inlet is modeled as a two-dimensional supersonic flow containing adjustable sidewall wedges and multiple fuel injection struts. The inlet geometry including the sidewall wedge angles, the number of injection struts, their sweepback relative to the vehicle reference line, and strut cross-section are user selectable. Combustion is currently represented by a Rayleigh line calculation including corrections for variable gas properties; improved models are being developed for this important element of the propulsion flow field. The program generates (1) variation of thrust magnitude and direction with angle of attack, (2) pitching moment and line of action of the thrust vector, (3) pressure and temperature distributions throughout the system, and (4) performance parameters such as thrust coefficient, specific impulse, mass flow rates, and equivalence ratio. Preliminary results are in good agreement with available performance data for systems resembling the NASP vehicle configuration.

Flandro, G. A.↗

NASA NTRS1991

Automated rendezvous and capture system

This paper describes an ARC system that is an attempt to simplify opration, reduce energy requirements, reduce weight, and provide longterm use and reliability. The ARC system is a laser/optical/holographic (LOH) control system for guidance, rendezvous, and docking (RVD). The LOH/RVD utilizes a hologram, residing at the target platform. Excited by a laser diode, the hologram projects an image at a given distance from the platform. A vision system in the automated chase vehicle sees the projected image and, by optical comparisons, guides the chase vehicle to that image, reaching a proximity conductive to soft docking. The vision system then shifts to a second hologram image holding at close proximity (2mm) to the target platform and guides to it for controlled, precise docking at the rendezvous point. The holographic image projections from the target platform, are composed of color hues and may be circular, triangular or of any other shape and texture that may enhance the ability of the chase vehicle's vision system to analyze information pertinent to velocity, attitude, and roll of the target platform. Any movement of the image, whether planned or errant, will be translated by the vision system into synchronous adjustments throughout the vehicle approach path.

Kader, Jack B.↗

NASA NTRS1993

Orbit determination and orbit control for the Earth Observing System (EOS) AM spacecraft

Future NASA Earth Observing System (EOS) Spacecraft will make measurements of the earth's clouds, oceans, atmosphere, land and radiation balance. These EOS Spacecraft will be part of the NASA Mission to Planet Earth. This paper specifically addresses the EOS AM Spacecraft, referred to as 'AM' because it has a sun-synchronous orbit with a 10:30 AM descending node. This paper describes the EOS AM Spacecraft mission orbit requirements, orbit determination, orbit control, and navigation system impact on earth based pointing. The EOS AM Spacecraft will be the first spacecraft to use the TDRSS Onboard Navigation System (TONS) as the primary means of navigation. TONS flight software will process one-way forward Doppler measurements taken during scheduled TDRSS contacts. An extended Kalman filter will estimate spacecraft position, velocity, drag coefficient correction, and ultrastable master oscillator frequency bias and drift. The TONS baseline algorithms, software, and hardware implementation are described in this paper. TONS integration into the EOS AM Spacecraft Guidance, Navigation, and Control (GN&C) System; TONS assisted onboard time maintenance; and the TONS Ground Support System (TGSS) are also addressed.

Herberg, Joseph R.↗

NASA NTRSDate not supplied

A CFD Validation Study using Artemis 1 Orbital Slosh Test Data

Advancements in liquid propellant management science and technologies are key to increasing safety, decreasing cost, and increasing payload mass of NASA space missions. Liquid propellant usually comprises a large portion of the total mass of launch vehicles and spacecraft, so predicting and controlling the motion of it are important. Computational fluid dynamics (CFD) programs are critical to predicting slosh dynamics, but CFD programs require experimental validation with physically relevant test data before the results can be trusted. Low-gravity slosh test data is lacking, and most of what exists is inadequate for CFD validation. A potential slosh risk for Orion, and the lack of validated low-gravity slosh models, led to performing on-orbit slosh tests during the Artemis 1 mission to assess the impacts of slosh on the guidance, navigation, and control of Orion. This work details the validation of a coupled slosh-motion CFD model using the Orion orbital slosh test data set.

CFD↗

NASA NTRSDate not supplied

A CFD Validation Study Using Artemis 1 Orbital Slosh Test Data

CFD↗

NASA NTRS2005

Expected EDL navigation performance with spacecraft to spacecraft data

Pinpoint landing (defined for the purpose of this discussion as landing within 1km of a preselected target) is a key Advanced Entry, Descent and Landing (EDL) technology for future Mars landers. Key scientific goals for Mars exploration, such as the search for water and characterization of aqueous process on Mars, the study of mineralogy and weathering of the Martian surface and the search for preserved biosignatures in Martian rocks, requires placing landers at pre-defined locations of greatest scientific interest. The capability to land within 1 km of a pre-defined landing site will improve safety and enable landing within roving range of sites of scientific interest while avoiding hazardous areas. A critical component of the closed-loop guidance, navigation and control (GN&C) system required for pinpoint landing is position and velocity estimation in real time. Spacecraft-to-spacecraft navigation will take advantage of the UHF link between two spacecraft (i.e. to an orbiter from an approaching lander for EDL telemetry relay) to build radiometric data, specifically the velocity between the two spacecraft along the radio beam, that are processed to determine position and velocity in real time. The improved onboard state knowledge provided by spacecraft-to-spacecraft navigation will reduce the landed position error and improve the performance of entry guidance. Results from the first of two years planned for this effort are documented here, including selection and documentation of prototype algorithms that will go forward into flight code along with analysis results used to define the algorithm set.

descent and landing↗

NASA NTRS2011

Docking Offset Between the Space Shuttle and the International Space Station and Resulting Impacts to the Transfer of Attitude Reference and Control

The Space Shuttle does not dock at an exact 90 degrees to the International Space Station (ISS) x-body axis. This offset from 90 degrees, along with error sources within their respective attitude knowledge, causes the two vehicles to never completely agree on their attitude, even though they operate as a single, mated stack while docked. The docking offset can be measured in flight when both vehicles have good attitude reference and is a critical component in calculations to transfer attitude reference from one vehicle to another. This paper will describe how the docking offset and attitude reference errors between both vehicles are measured and how this information would be used to recover Shuttle attitude reference from ISS in the event of multiple failures. During STS-117, ISS on-board Guidance, Navigation and Control (GNC) computers began having problems and after several continuous restarts, the systems failed. The failure took the ability for ISS to maintain attitude knowledge. This paper will also demonstrate how with knowledge of the docking offset, the contingency procedure to recover Shuttle attitude reference from ISS was reversed in order to provide ISS an attitude reference from Shuttle. Finally, this paper will show how knowledge of the docking offset can be used to speed up attitude control handovers from Shuttle to ISS momentum management. By taking into account the docking offset, Shuttle can be commanded to hold a more precise attitude which better agrees with the ISS commanded attitude such that start up transients with the ISS momentum management controllers are reduced. By reducing start-up transients, attitude control can be transferred from Shuttle to ISS without the use of ISS thrusters saving precious on-board propellant, crew time and minimizing loads placed upon the mated stack.

Helms, W. Jason↗

NASA NTRS1994

SCL: An off-the-shelf system for spacecraft control

In this age of shrinking military, civil, and commercial space budgets, an off-the-shelf solution is needed to provide a multimission approach to spacecraft control. A standard operational interface which can be applied to multiple spacecraft allows a common approach to ground and space operations. A trend for many space programs has been to reduce operational staff by applying autonomy to the spacecraft and to the ground stations. The Spacecraft Command Language (SCL) system developed by Interface and Control Systems, Inc. (ICS) provides an off-the-shelf solution for spacecraft operations. The SCL system is designed to provide a hyper-scripting interface which remains standard from program to program. The spacecraft and ground station hardware specifics are isolated to provide the maximum amount of portability from system to system. Uplink and downlink interfaces are also isolated to allow the system to perform independent of the communications protocols chosen. The SCL system can be used for both the ground stations and the spacecraft, or as a value added package for existing ground station environments. The SCL system provides an expanded stored commanding capability as well as a rule-based expert system on-board. The expert system allows reactive control on-board the spacecraft for functions such as electrical power systems (EPS), thermal control, etc. which have traditionally been performed on the ground. The SCL rule and scripting capability share a common syntax allowing control of scripts from rules and rules from scripts. Rather than telemeter over sampled data to the ground, the SCL system maintains a database on-board which is available for interrogation by the scripts and rules. The SCL knowledge base is constructed on the ground and uploaded to the spacecraft. The SCL system follows an open-systems approach allowing other tasks to communicate with SCL on the ground and in space. The SCL system was used on the Clementine program (launched January 25, 1994) and is required to have bidirectional communications with the guidance, navigation, and control (GNC) algorithms which were written as another task. Sequencing of the spacecraft maneuvers are handled by SCL, but the low-level thruster pulse commands are handled by the GNC software. Attitude information is reported back as telemetry, allowing the SCL expert system to inference on the changing data. The Clementine SCL flight software was largely reused from another Naval Center for Space Technology (NCST) satellite program. This paper details the SCL architecture and how an off-the-shelf solution makes sense for multimission spacecraft programs. The Clementine mission will be used as a case study in the application of the SCL to a 'fast track' program. The benefits of such a system in a 'better, cheaper, faster' climate will be discussed.

Buckley, Brian↗

NASA NTRS2021

Spacecraft Line-Of-Sight Jitter Management and Mitigation Lessons Learned And Engineering Best Practices

Predicting, managing, controlling, and testing spacecraft line-of-sight (LoS) jitter caused by micro-vibrations due to on-board internal disturbance sources is a formidable multidisciplinary engineering task. It is especially challenging for those missions hosting high-performance (e.g., nano-radian/milli-arcsecond class), vibration-sensitive optical sensor payloads with stringent pointing stability requirements. The Nation Aeronautics and Space Administration (NASA) and the European Space Agency (ESA) are planning technically aggressive spaceflight missions that include ultra-high-performance optical payloads with delicate, highly vibration-sensitive scientific and observational instruments. The guidance, navigation, and control community of practice will need to leverage collective experiences and document their best practices and lessons learned to address future micro-vibration challenges. To identify lessons learned and best practices the NASA Engineering & Safety Center sponsored a 2-day Spacecraft LoS Jitter Workshop in late 2019. The workshop’s goal was to provide a multidisciplinary forum to elicit deeper understanding of the issues related to addressing the spacecraft LoS jitter/micro-vibration problem. The primary objective was to identify, document, and share lessons learned, best practices, and preferred options for jitter-related analysis and test activities. Representatives from NASA, ESA, along with NASA’s industrial partners, independent consultant subject matter experts, and members of academia participated in the workshop. This paper will describe the motivation for the workshop and summarize the identified findings and recommendations.

Swanson, Davin K.↗