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All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

Constellation Architecture Team-Lunar: Lunar Habitat Concepts

This paper will describe lunar habitat concepts that were defined as part of the Constellation Architecture Team-Lunar (CxAT-Lunar) in support of the Vision for Space Exploration. There are many challenges to designing lunar habitats such as mission objectives, launch packaging, lander capability, and risks. Surface habitats are required in support of sustaining human life to meet the mission objectives of lunar exploration, operations, and sustainability. Lunar surface operations consist of crew operations, mission operations, EVA operations, science operations, and logistics operations. Habitats are crewed pressurized vessels that include surface mission operations, science laboratories, living support capabilities, EVA support, logistics, and maintenance facilities. The challenge is to deliver, unload, and deploy self-contained habitats and laboratories to the lunar surface. The CxAT-Lunar surface campaign analysis focused on three primary trade sets of analysis. Trade set one (TS1) investigated sustaining a crew of four for six months with full outpost capability and the ability to perform long surface mission excursions using large mobility systems. Two basic habitat concepts of a hard metallic horizontal cylinder and a larger inflatable torus concept were investigated as options in response to the surface exploration architecture campaign analysis. Figure 1 and 2 depicts the notional outpost configurations for this trade set. Trade set two (TS2) investigated a mobile architecture approach with the campaign focused on early exploration using two small pressurized rovers and a mobile logistics support capability. This exploration concept will not be described in this paper. Trade set three (TS3) investigated delivery of a "core' habitation capability in support of an early outpost that would mature into the TS1 full outpost capability. Three core habitat concepts were defined for this campaign analysis. One with a four port core habitat, another with a 2 port core habitat, and the third investigated leveraging commonality of the lander ascent module and airlock pressure vessel hard shell. The paper will describe an overview of the various habitat concepts and their functionality. The Crew Operations area includes basic crew accommodations such as sleeping, eating, hygiene and stowage. The EVA Operations area includes additional EVA capability beyond the suit-port airlock function such as redundant airlock(s), suit maintenance, spares stowage, and suit stowage. The Logistics Operations area includes the enhanced accommodations for 180 days such as closed loop life support systems hardware, consumable stowage, spares stowage, interconnection to the other Hab units, and a common interface mechanism for future growth and mating to a pressurized rover. The Mission & Science Operations area includes enhanced outpost autonomy such as an IVA glove box, life support, and medical operations.

Toups, Larry↗
NASA NTRSDate not supplied

Rapid Spacecraft Payload Development: In-Orbit Demonstration of Flight Software Reuse, Scalability, and Dependability

As space mission design trends towards shared, multi-mission platforms and high-performance onboard computing architectures, the number of spacecraft launched into operation is also steadily rising. Through ridesharing, spacecraft miniaturization, and other cost-reduction measures, the barriers to space are lowering, resulting in compounded growth in the amount of flight software being deployed. To meet the needs of both the growing quantity and evolving nature of spacecraft, flight software design must accordingly adapt to support more efficient development, solutions to computational resource-sharing, and software reusability. This paper focuses on a software payload demonstrating several core technologies that improve the state-of-the-art in these identified areas. Launched into low-earth orbit in January 2022, our software payload was conceived, designed, and delivered in a span of merely two months. It was developed on top of the NASA core Flight System (cFS) framework and the Distributed Spacecraft Autonomy (DSA) Comm cFS application, which translates cFS software bus messages across a Data Distribution Service (DDS) network. The flight software, packaged in Linux container images, was deployed as one of 18 flight applications managed through the Unibap SpaceCloud Framework. The applications were run on a Unibap iX5-102 radiation-tolerant payload computer, hosted on the D-Orbit SCV-004 spacecraft as part of an ESA-sponsored in-orbit technology test. Our payload, referred to as the DSA D-Orbit software, demonstrates the reusability of the DSA Comm app in a substantially different context and purpose as its original mission. Comm’s original design goal was to reliably distribute messages between spacecraft swarms of arbitrary size and dynamic network topology. However, we leverage this same functionality to introduce redundancy and opportunistic parallel data processing in the context of a representative onboard image processing workload. This adaptive mission architecture was enabled in part by the SpaceCloud Framework’s use of container virtualization as the payload integration interface. By using a base container image with common high-level language runtimes and libraries, we were able to rapidly design, develop, and validate our image processing application without many of the technological barriers common to flight software development. We present details the goals, approach, results, and lessons learned through this technology demonstration experiment and contextualize those observations against present and future challenges in spacecraft software development.

computer programming↗
NASA NTRS2010

Reconfigurable Fault Tolerance for FPGAs

The invention allows a field-programmable gate array (FPGA) or similar device to be efficiently reconfigured in whole or in part to provide higher capacity, non-redundant operation. The redundant device consists of functional units such as adders or multipliers, configuration memory for the functional units, a programmable routing method, configuration memory for the routing method, and various other features such as block RAM, I/O (random access memory, input/output) capability, dedicated carry logic, etc. The redundant device has three identical sets of functional units and routing resources and majority voters that correct errors. The configuration memory may or may not be redundant, depending on need. For example, SRAM-based FPGAs will need some type of radiation-tolerant configuration memory, or they will need triple-redundant configuration memory. Flash or anti-fuse devices will generally not need redundant configuration memory. Some means of loading and verifying the configuration memory is also required. These are all components of the pre-existing redundant FPGA. This innovation modifies the voter to accept a MODE input, which specifies whether ordinary voting is to occur, or if redundancy is to be split. Generally, additional routing resources will also be required to pass data between sections of the device created by splitting the redundancy. In redundancy mode, the voters produce an output corresponding to the two inputs that agree, in the usual fashion. In the split mode, the voters select just one input and convey this to the output, ignoring the other inputs. In a dual-redundant system (as opposed to triple-redundant), instead of a voter, there is some means to latch or gate a state update only when both inputs agree. In this case, the invention would require modification of the latch or gate so that it would operate normally in redundant mode, and would separately latch or gate the inputs in non-redundant mode.

Shuler, Robert, Jr.↗
NASA NTRS2010

Methods and circuitry for reconfigurable SEU/SET tolerance

A device is disclosed in one embodiment that has multiple identical sets of programmable functional elements, programmable routing resources, and majority voters that correct errors. The voters accept a mode input for a redundancy mode and a split mode. In the redundancy mode, the programmable functional elements are identical and are programmed identically so the voters produce an output corresponding to the majority of inputs that agree. In a split mode, each voter selects a particular programmable functional element output as the output of the voter. Therefore, in the split mode, the programmable functional elements can perform different functions, operate independently, and/or be connected together to process different parts of the same problem.

Shuler, Jr., Robert L.↗
NASA NTRS1975

Space Tug avionics definition study. Volume 1: Executive summary

A top down approach was used to identify, compile, and develop avionics functional requirements for all flight and ground operational phases. Such requirements as safety mission critical functions and criteria, minimum redundancy levels, software memory sizing, power for tug and payload, data transfer between payload, tug, shuttle, and ground were established. Those functional requirements that related to avionics support of a particular function were compiled together under that support function heading. This unique approach provided both organizational efficiency and traceability back to the applicable operational phase and event. Each functional requirement was then allocated to the appropriate subsystems and its particular characteristics were quantified.

Source record↗
NASA NTRS1978

Reliable dual-redundant sensor failure detection and identification for the NASA F-8 DFBW aircraft

A technique was developed which provides reliable failure detection and identification (FDI) for a dual redundant subset of the flight control sensors onboard the NASA F-8 digital fly by wire (DFBW) aircraft. The technique was successfully applied to simulated sensor failures on the real time F-8 digital simulator and to sensor failures injected on telemetry data from a test flight of the F-8 DFBW aircraft. For failure identification the technique utilized the analytic redundancy which exists as functional and kinematic relationships among the various quantities being measured by the different control sensor types. The technique can be used not only in a dual redundant sensor system, but also in a more highly redundant system after FDI by conventional voting techniques reduced to two the number of unfailed sensors of a particular type. In addition the technique can be easily extended to the case in which only one sensor of a particular type is available.

Deckert, J. C.↗
NASA NTRS1976

Redundant Strapdown Laser Gyro Navigation System

For the last several years, NASA has pursued the development of low-cost high-reliability inertial navigation systems that would satisfy a broad spectrum of future space and avionics missions. Two specific programs have culminated in the construction of a Redundant Strapdown Laser Gyro Navigation System. These two programs were for development of a space ultrareliable modular computer (SUMC) and a redundant laser gyro inertial measurement unit (IMU). The SUMC is a digital computer that employs state-of-the-art large-scale integrated circuits configured in a functional modular breakdown. The redundant laser gyro IMU is a six-pack strapdown sensor package in a dodecahedron configuration which uses six laser gyros to provide incremental angular positions and six accelerometers for linear velocity outputs. The sensor arrangement allows automatic accommodation of two failures; a third failure can be tolerated provided it can be determined. The navigation system also includes redundant power supplies, built-in test-equipment (BITE) circuits for failure detection, and software which provides for navigation, redundancy management, and automatic calibration and alignment.

Mcpherson, B. W.↗
NASA NTRS2011

Comparison Modeling of System Reliability for Future NASA Projects

A National Aeronautics and Space Administration (NASA) supported Reliability, Maintainability, and Availability (RMA) analysis team developed a unique RMA analysis methodology using cut set and importance measure analysis in order to comparison model proposed avionics computing architectures. In this paper we will present this efficient application of the RMA analysis methodology for importance measures that includes Reliability Block Diagram (RED) Analysis, Comparison modeling, Cut Set Analysis, and Importance Measure Analysis. We will also demonstrate that integrating RMA early in the system design process as a key to success by providing a fundamental decision metric supporting design selection. The RMA analysis methodology presented in this paper and applied to the avionics architectures enhances the usual way of predicting the need for redundancy based on failure rates or subject matter expert opinion. Using the REDs and the minimal cut sets, along with the Fussell-Vesely (FV) factors, importance measures are calculated for each functional element in the architectures. This paper presents an application of the FV importance measures and presents an improved methodology for using importance measures in success space (instead of failure space) to compare architectures. These importance measures are used to determine which functional element would be most likely to cause a system failure, thus, quickly identifying the path to increase the overall system reliability by either procuring more reliable functional elements or adding redundancy. This application of the RMA analysis methodology, using RBD analysis, cut set analysis, and the importance measure analysis, allows the avionics design team to better understand and compare the vulnerabilities in each of the architectures, enabling them to address the deficiencies in the design architectures more efficiently, while balancing the need to design for optimum weight and space allocations.

Gillespie, Amanda M.↗
NASA NTRS1981

Automated Power Systems Management (APSM)

A breadboard power system incorporating autonomous functions of monitoring, fault detection and recovery, command and control was developed, tested and evaluated to demonstrate technology feasibility. Autonomous functions including switching of redundant power processing elements, individual load fault removal, and battery charge/discharge control were implemented by means of a distributed microcomputer system within the power subsystem. Three local microcomputers provide the monitoring, control and command function interfaces between the central power subsystem microcomputer and the power sources, power processing and power distribution elements. The central microcomputer is the interface between the local microcomputers and the spacecraft central computer or ground test equipment.

Bridgeforth, A. O.↗
NASA NTRS1978

On reliable control system designs

A mathematical model for use in the design of reliable multivariable control systems is discussed with special emphasis on actuator failures and necessary actuator redundancy levels. The model consists of a linear time invariant discrete time dynamical system. Configuration changes in the system dynamics are governed by a Markov chain that includes transition probabilities from one configuration state to another. The performance index is a standard quadratic cost functional, over an infinite time interval. The actual system configuration can be deduced with a one step delay. The calculation of the optimal control law requires the solution of a set of highly coupled Riccati-like matrix difference equations. Results can be used for off-line studies relating the open loop dynamics, required performance, actuator mean time to failure, and functional or identical actuator redundancy, with and without feedback gain reconfiguration strategies.

Birdwell, J. D.↗
NASA NTRS1991

On an adaptive truss manipulator space crane concept

The paper describes an adaptive truss manipulator (ATM) space crane concept for in-space assembly and construction. The underlying mechanism of an ATM is that the batten members of the constituent octahedral modules are controllable in length. Through geometric transformations of the constituent modules, the basic manipulator functions such as articulation can be performed by the ATM. The mechanism can also provide deployment/retraction, high dexterity motion, and bracing operation. The advantages of an ATM over the conventional multijoint-multilink anthropomorphic manipulator are its compact stowage volume for in-space storage and mobility, deployment as needed, high dexterity in complex workspace, and high redundancy of the actuator function. The kinematic description of an ATM is formulated in the global cartesian coordinate system.

Chen, Gun-Shing↗
NASA NTRS1973

Atmosphere Explorer control system software (version 2.0)

The Atmosphere Explorer Control System (AECS) was developed to provide automatic computer control of the Atmosphere Explorer spacecraft and experiments. The software performs several vital functions, such as issuing commands to the spacecraft and experiments, receiving and processing telemetry data, and allowing for extensive data processing by experiment analysis programs. The AECS was written for a 48K XEROX Data System Sigma 5 computer, and coexists in core with the XDS Real-time Batch Monitor (RBM) executive system. RBM is a flexible operating system designed for a real-time foreground/background environment, and hence is ideally suited for this application. Existing capabilities of RBM have been used as much as possible by AECS to minimize programming redundancy. The most important functions of the AECS are to send commands to the spacecraft and experiments, and to receive, process, and display telemetry data.

Mocarsky, W.↗
NASA NTRS1979

On reliable control system designs with and without feedback reconfigurations

This paper contains an overview of a theoretical framework for the design of reliable multivariable control systems, with special emphasis on actuator failures and necessary actuator redundancy levels. Using a linear model of the system, with Markovian failure probabilities and quadratic performance index, an optimal stochastic control problem is posed and solved. The solution requires the iteration of a set of highly coupled Riccati-like matrix difference equations; if these converge one has a reliable design; if they diverge, the design is unreliable, and the system design cannot be stabilized. In addition, it is shown that the existence of a stabilizing constant feedback gain and the reliability of its implementation is equivalent to the convergence properties of a set of coupled Riccati-like matrix difference equations. In summary, these results can be used for offline studies relating the open loop dynamics, required performance, actuator mean time to failure, and functional or identical actuator redundancy, with and without feedback gain reconfiguration strategies.

Birdwell, J. D.↗
NASA NTRS1991

Balancing reliability and cost to choose the best power subsystem

A mathematical model is presented for computing total (spacecraft) subsystem cost including both the basic subsystem cost and the expected cost due to the failure of the subsystem. This model is then used to determine power subsystem cost as a function of reliability and redundancy. Minimum cost and maximum reliability and/or redundancy are not generally equivalent. Two example cases are presented. One is a small satellite, and the other is an interplanetary spacecraft.

Suich, Ronald C.↗
NASA NTRS1990

Characteristics of the GOES I-M Imager and Sounder

The design and the parameters of the improved thermal-imaging and sounding instruments (the Imager and the Sounder) that will be part of the instrument complements of the next-generation Geostationary Operational Environmental Satellite (GOES I-M) are discussed. The new design incorporates many features that enhance instrumental reliability over the previous GOES radiometric instruments, such as independently functioning Sounder and Imager, redundancy, and more reliable position sensors and lubrication methods. Tables are presented which list the instrument parameters of the GOES I-M Imager and Sounder and the performance characteristics of the two instruments.

Ernst, Thomas J.↗
NASA NTRS1991

Balancing reliability and cost to choose the best power subsystem

The justification of the increased cost of a more reliable power subsystem is discussed. A mathematical model is presented for computing total spacecraft subsystem cost including both the basic subsystem cost and the expected cost due to the failure of the subsystem. This is then used to determine the power subsystem cost as a function of reliability and redundancy. Minimum cost and maximum reliability and/or redundancy are not generally equivalent. Two example cases are presented. One is a small satellite, and the other is an interplanetary spacecraft.

Suich, Ronald C.↗
NASA NTRS2002

Fine Sun Sensor Field of View Calibration

The fine Sun sensor (FSS) used on many spacecraft consists of two independent single-axis sensors, nominally mounted perpendicularly, that detect Sun angle across a typical field of view of +/- 32 degrees. The nonlinear function that maps the measured counts into an observed angle is called the transfer function. The FSS transfer function provided by the manufacturer consists of nine parameters for each of the two sensitive axes. An improved transfer function has been previously reported that achieves a significant accuracy improvement across the entire field of view. This new function expands the parameter set to 12 coefficients per axis and includes cross terms combining counts from both axes. To make best use of the FSS for spacecraft attitude determination, it must be calibrated after launch. We are interested in simplifying the postlaunch calibration procedure for estimating improvements to the 24 parameters in the transfer function. This paper discusses how to recombine the terms of the transfer function to reduce their redundancy without decreasing its accuracy and then presents an attitude dependent procedure for estimating the parameters. The end result is a calibration algorithm that is easier to use and does not sacrifice accuracy. Results of calibration using on-orbit data are presented.

Sedlak, Joseph E.↗
NASA NTRS2014

Design, Integration, Certification and Testing of the Orion Crew Module Propulsion System

The Orion Multipurpose Crew Vehicle (MPCV) is NASA's next generation spacecraft for human exploration of deep space. Lockheed Martin is the prime contractor for the design, development, qualification and integration of the vehicle. A key component of the Orion Crew Module (CM) is the Propulsion Reaction Control System, a high‐flow hydrazine system used during re‐entry to orient the vehicle for landing. The system consists of a completely redundant helium (GHe) pressurization system and hydrazine fuel system with monopropellant thrusters. The propulsion system has been designed, integrated, and qualification tested in support of the Orion program's first orbital flight test, Exploration Flight Test One (EFT‐1), scheduled for 2014. A subset of the development challenges and lessons learned from this first flight test campaign will be discussed in this paper for consideration when designing future spacecraft propulsion systems. The CONOPS and human rating requirements of the CM propulsion system are unique when compared with a typical satellite propulsion reaction control system. The system requires a high maximum fuel flow rate. It must operate at both vacuum and sea level atmospheric pressure conditions. In order to meet Orion's human rating requirements, multiple parts of the system must be redundant, and capable of functioning after spacecraft system fault events.

McKay, Heather↗