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NASA NTRS1998

Development and Testing of Automatically Generated ACS Flight Software for the MAP Spacecraft

By integrating the attitude determination and control system (ACS) analysis and design, flight software development, and flight software testing processes, it is possible to improve the overall spacecraft development cycle, as well as allow for more thorough software testing. One of the ways to achieve this integration is to use code-generation tools to automatically generate components of the ACS flight software directly from a high-fidelity (HiFi) simulation. In the development of the Microwave Anisotropy Probe (MAP) spacecraft, currently underway at the NASA Goddard Space Flight Center, approximately 1/3 of the ACS flight software was automatically generated. In this paper, we will examine each phase of the ACS subsystem and flight software design life cycle: analysis, design, and testing. In the analysis phase, we scoped how much software we would automatically generate and created the initial interface. The design phase included parallel development of the HiFi simulation and the hand-coded flight software components. Everything came together in the test phase, in which the flight software was tested, using results from the HiFi simulation as one of the bases of comparison for testing. Because parts of the spacecraft HiFi simulation were converted into flight software, more care needed to be put into its development and configuration control to support both the HiFi simulation and flight software. The components of the HiFi simulation from which code was generated needed to be designed based on the fact that they would become flight software. This process involved such considerations as protecting against mathematical exceptions, using acceptable module and parameter naming conventions, and using an input/output interface compatible with the rest of the flight software. Maintaining good configuration control was an issue for the HiFi simulation and the flight software, and a way to track the two systems was devised. Finally, an integrated test approach was devised to support flight software testing at both the unit- and build-test levels using the HiFi simulation to generate data for performance verification. Another benefit of the simulation and code-generation application used on the MAP project is that it supported bringing flight software and test data into the HiFi simulation environment. It was possible to integrate parts of the hand-coded flight software into the HiFi simulation, and also possible to import flight software test data for comparison and performance verification. This capability was used to incorporate the flight software Kalman filter into the HiFi simulation. This enabled us to greatly increase the amount of testing that could be done on the filter, because we could exert a greater degree of control over the software-only simulation than over the flight software test environment. Also, since the simulation could be used to run the Kalman filter faster than real time, our testing efficiency was greatly increased. We will conclude our discussion with a summary of the lessons learned thus far using automatically- generated code for the MAP project, and the spacecraft status as we work towards our scheduled launch in the year 2000.

ODonnell, James R., Jr.↗

NASA NTRS2011

Early Oscillation Detection for DC/DC Converter Fault Diagnosis

The electrical power system of a spacecraft plays a very critical role for space mission success. Such a modern power system may contain numerous hybrid DC/DC converters both inside the power system electronics (PSE) units and onboard most of the flight electronics modules. One of the faulty conditions for DC/DC converter that poses serious threats to mission safety is the random occurrence of oscillation related to inherent instability characteristics of the DC/DC converters and design deficiency of the power systems. To ensure the highest reliability of the power system, oscillations in any form shall be promptly detected during part level testing, system integration tests, flight health monitoring, and on-board fault diagnosis. The popular gain/phase margin analysis method is capable of predicting stability levels of DC/DC converters, but it is limited only to verification of designs and to part-level testing on some of the models. This method has to inject noise signals into the control loop circuitry as required, thus, interrupts the DC/DC converter's normal operation and increases risks of degrading and damaging the flight unit. A novel technique to detect oscillations at early stage for flight hybrid DC/DC converters was developed.

Wang, Bright L.↗

NASA NTRS2014

Independent Verification of Mars-GRAM 2010 with Mars Climate Sounder Data

The Mars Global Reference Atmospheric Model (Mars-GRAM) is an engineering-level atmospheric model widely used for diverse mission and engineering applications. Applications of Mars-GRAM include systems design, performance analysis, and operations planning for aerobraking, entry, descent and landing, and aerocapture. Atmospheric influences on landing site selection and long-term mission conceptualization and development can also be addressed utilizing Mars-GRAM. Mars-GRAM's perturbation modeling capability is commonly used, in a Monte Carlo mode, to perform high-fidelity engineering end-to-end simulations for entry, descent, and landing. Mars-GRAM is an evolving software package resulting in improved accuracy and additional features. Mars-GRAM 2005 has been validated against Radio Science data, and both nadir and limb data from the Thermal Emission Spectrometer (TES). From the surface to 80 km altitude, Mars-GRAM is based on the NASA Ames Mars General Circulation Model (MGCM). Above 80 km, Mars-GRAM is based on the University of Michigan Mars Thermospheric General Circulation Model (MTGCM). The most recent release of Mars-GRAM 2010 includes an update to Fortran 90/95 and the addition of adjustment factors. These adjustment factors are applied to the input data from the MGCM and the MTGCM for the mapping year 0 user-controlled dust case. The adjustment factors are expressed as a function of height (z), latitude and areocentric solar longitude (Ls).

Justh, Hilary L.↗

NASA NTRS2017

Designing to Sample the Unknown: Lessons from OSIRIS-REx Project Systems Engineering

On September 8, 2016, the third NASA New Frontiers mission launched on an Atlas V 411. The Origins, Spectral Interpretation, Resource Identification, Security-Regolith Explorer (OSIRIS-REx) will rendezvous with asteroid Bennu in 2018, collect a sample in 2020, and return that sample to Earth in September 2023. The development team has overcome a number of challenges in order to design and build a system that will make contact with an unexplored, airless, low-gravity body. This paper will provide an overview of the mission, then focus in on the system-level challenges and some of the key system-level processes. Some of the lessons here are unique to the type of mission, like discussion of operating at a largely-unknown, low-gravity object. Other lessons, particularly from the build phase, have broad implications. The OSIRIS-REx risk management process was particularly effective in achieving an on-time and under-budget development effort. The systematic requirements management and verification and the system validation also helped identify numerous potential problems. The final assessment of the OSIRIS-REx performance will need to wait until the sample is returned in 2023, but this post-launch assessment will capture some of the key systems-engineering lessons from the development team.

Asteroid↗

NASA NTRS2024

xEMU Thermal Vacuum Testing Overview

The Exploration Extravehicular Mobility Unit (xEMU) project was the culimation of over a decade of spacesuit development that was performed in-house at the NASA Johnson Space Center. This project reached a level where almost fully completed development fidelity spacesuits had been designed, assembled and tested in an integrated configuration. The xEMU Development-Verification-Test (DVT) hardware was assembled into two different xEMU test articles and underwent a series of thermal-vacuum tests at the Johnson Space Center’s Chamber B. These tests not only gathered data on thermal performance, but also exercised the life support system, communication system, suit information systems, and suit avionics. This complex test has and will continue to produce many meaningful reports ranging from component level test results (for example on spacesuit boots), test design of heater cages to simulate thermal environments for a spacesuit test article, and higher level thermal performance of subsystems (such as the Portable Life Support System) or the entire assembly. This paper provides an overview of the test configuration and also top level results from this highly successful integrated test of the xEMU.

PLSS↗

NASA NTRSDate not supplied

xEMU Thermal Vacuum Testing Overview

PLSS↗

NASA NTRS2008

Integrated System Test Approaches for the NASA Ares I Crew Launch Vehicle

NASA is maturing test and evaluation plans leading to flight readiness of the Ares I crew launch vehicle. Key development, qualification, and verification tests are planned . Upper stage engine sea-level and altitude testing. First stage development and qualification motors. Upper stage structural and thermal development and qualification test articles. Main Propulsion Test Article (MPTA). Upper stage green run testing. Integrated Vehicle Ground Vibration Testing (IVGVT). Aerodynamic characterization testing. Test and evaluation supports initial validation flights (Ares I-Y and Orion 1) and design certification.

Cockrell, Charles↗

NASA NTRS1990

A noise assessment and prediction system

A system has been designed to provide an assessment of noise levels that result from testing activities at Aberdeen Proving Ground, Md. The system receives meteorological data from surface stations and an upper air sounding system. The data from these systems are sent to a meteorological model, which provides forecasting conditions for up to three hours from the test time. The meteorological data are then used as input into an acoustic ray trace model which projects sound level contours onto a two-dimensional display of the surrounding area. This information is sent to the meteorological office for verification, as well as the range control office, and the environmental office. To evaluate the noise level predictions, a series of microphones are located off the reservation to receive the sound and transmit this information back to the central display unit. The computer models are modular allowing for a variety of models to be utilized and tested to achieve the best agreement with data. This technique of prediction and model validation will be used to improve the noise assessment system.

Olsen, Robert O.↗

NASA NTRS2008

Using Optically Stimulated Electron Emission as an Inspection Method to Monitor Surface Contamination

During redesign of the Space Shuttle reusable solid rocket motor (RSRM), NASA amended the contract with ATK Launch Systems (then Morton Thiokol Inc.) with Change Order 966 to implement a contamination control and cleanliness verification method. The change order required: (1) A quantitative inspection method (2) A written record of actual contamination levels versus a known reject level (3) A method that is more sensitive than existing methods of visual and black light inspection. Black light inspection is only useful for inspection of contaminants that fluoresce near the 365 nm spectral line and is not useful for inspection of most silicones that will not produce strong fluorescence. Black light inspection conducted by a qualified inspector under controlled light is capable of detecting Conoco HD-2 grease in gross amounts and is very subjective due to operator sensitivity. Optically stimulated electron emission (OSEE), developed at the Materials and Process Laboratory at Marshall Space Flight Center (MSFC), was selected to satisfy Change Order 966. OSEE offers several important advantages over existing laboratory methods with similar sensitivity, e.g., spectroscopy and nonvolatile residue sampling, which provide turn around time, real time capability, and full coverage inspection capability. Laboratory methods require sample gathering and in-lab analysis, which sometimes takes several days to get results. This is not practical in a production environment. In addition, these methods do not offer full coverage inspection of the large components

Lingbloom, Mike S.↗

NASA NTRS2016

Functional Fault Model Development Process to Support Design Analysis and Operational Assessment

A functional fault model (FFM) is an abstract representation of the failure space of a givensystem. As such, it simulates the propagation of failure effects along paths between the origin ofthe system failure modes and points within the system capable of observing the failure effects. Asa result, FFMs may be used to diagnose the presence of failures in the modeled system. FFMsnecessarily contain a significant amount of information about the design, operations, and failuremodes and effects. One of the important benefits of FFMs is that they may be qualitative, ratherthan quantitative and, as a result, may be implemented early in the design process when there ismore potential to positively impact the system design. FFMs may therefore be developed andmatured throughout the monitored system's design process and may subsequently be used toprovide real-time diagnostic assessments that support system operations. This paper provides anoverview of a generalized NASA process that is being used to develop and apply FFMs. FFMtechnology has been evolving for more than 25 years. The FFM development process presented inthis paper was refined during NASA's Ares I, Space Launch System, and Ground SystemsDevelopment and Operations programs (i.e., from about 2007 to the present). Process refinementtook place as new modeling, analysis, and verification tools were created to enhance FFMcapabilities. In this paper, standard elements of a model development process (i.e., knowledgeacquisition, conceptual design, implementation & verification, and application) are describedwithin the context of FFMs. Further, newer tools and analytical capabilities that may benefit thebroader systems engineering process are identified and briefly described. The discussion isintended as a high-level guide for future FFM modelers.

Melcher, Kevin J.↗

NASA NTRS2024

Runtime Verification of Hard Realtime Systems With Copilot: A Tutorial

This presentation is a tutorial on RV using Copilot, a runtime verification framework for real-time embedded systems. Copilot monitors are written in a compositional, stream-based language with support for a variety of Temporal Logics (TL), which results in robust, high-level specifications that are easier to understand than their traditional counterparts. The framework translates monitor specifications into C code with static memory requirements, which can be compiled to run on embedded hardware.

runtime monitoring↗

NASA NTRS2001

Using Dynamic Interface Modeling and Simulation to Develop a Launch and Recovery Flight Simulation for a UH-60A Blackhawk

Joint Shipboard Helicopter Integration Process (JSHIP) is a Joint Test and Evaluation (JT&E) program sponsored by the Office of the Secretary of Defense (OSD). Under the JSHDP program is a simulation effort referred to as the Dynamic Interface Modeling and Simulation System (DIMSS). The purpose of DIMSS is to develop and test the processes and mechanisms that facilitate ship-helicopter interface testing via man-in-the-loop ground-based flight simulators. Specifically, the DIMSS charter is to develop an accredited process for using a flight simulator to determine the wind-over-the-deck (WOD) launch and recovery flight envelope for the UH-60A ship/helicopter combination. DIMSS is a collaborative effort between the NASA Ames Research Center and OSD. OSD determines the T&E and warfighter training requirements, provides the programmatics and dynamic interface T&E experience, and conducts ship/aircraft interface tests for validating the simulation. NASA provides the research and development element, simulation facility, and simulation technical experience. This paper will highlight the benefits of the NASA/JSHIP collaboration and detail achievements of the project in terms of modeling and simulation. The Vertical Motion Simulator (VMS) at NASA Ames Research Center offers the capability to simulate a wide range of simulation cueing configurations, which include visual, aural, and body-force cueing devices. The system flexibility enables switching configurations io allow back-to-back evaluation and comparison of different levels of cueing fidelity in determining minimum training requirements. The investigation required development and integration of several major simulation system at the VMS. A new UH-60A BlackHawk interchangeable cab that provides an out-the-window (OTW) field-of-view (FOV) of 220 degrees in azimuth and 70 degrees in elevation was built. Modeling efforts involved integrating Computational Fluid Dynamics (CFD) generated data of an LHA ship airwake and integrating a real-time ship motion model developed based on a batch model from Naval Surface Warfare Center. Engineering development and integration of a three degrees-of-freedom (DOF) dynamic seat to simulate high frequency rotor-dynamics dependent motion cues for use in conjunction with the large motion system was accomplished. The development of an LHA visual model in several different levels of resolution and an aural cueing system in which three separate fidelity levels could be selected were developed. VMS also integrated a PC-based E&S simFUSION system to investigate cost effective IG alternatives. The DIMSS project consists of three phases that follow an approved Validation, Verification and accreditation (VV&A) process. The first phase will support the accreditation of the individual subsystems and models. The second will follow the verification and validation of the integrated subsystems and models, and will address fidelity requirements of the integrated models and subsystems. The third and final phase will allow the verification and validation of the full system integration. This VV&A process will address the utility of the simulated WOD launch and recovery envelope. Simulations supporting the first two stages have been completed and the data is currently being reviewed and analyzed.

Sweeney, Christopher↗

NASA NTRS1978

Advanced rotorcraft technology: Task force report

The technological needs and opportunities related to future civil and military rotorcraft were determined and a program plan for NASA research which was responsive to the needs and opportunities was prepared. In general, the program plan places the primary emphasis on design methodology where the development and verification of analytical methods is built upon a sound data base. The four advanced rotorcraft technology elements identified are aerodynamics and structures, flight control and avionic systems, propulsion, and vehicle configurations. Estimates of the total funding levels that would be required to support the proposed program plan are included.

Source record↗

NASA NTRS2015

Verification of Functional Fault Models and the Use of Resource Efficient Verification Tools

Functional fault models (FFMs) are a directed graph representation of the failure effect propagation paths within a system's physical architecture and are used to support development and real-time diagnostics of complex systems. Verification of these models is required to confirm that the FFMs are correctly built and accurately represent the underlying physical system. However, a manual, comprehensive verification process applied to the FFMs was found to be error prone due to the intensive and customized process necessary to verify each individual component model and to require a burdensome level of resources. To address this problem, automated verification tools have been developed and utilized to mitigate these key pitfalls. This paper discusses the verification of the FFMs and presents the tools that were developed to make the verification process more efficient and effective.

reliability↗

NASA NTRS2003

Low-Power SOI CMOS Transceiver

The work aims at developing a low-power Silicon on Insulator Complementary Metal Oxide Semiconductor (SOI CMOS) Transceiver for deep-space communications. RF Receiver must accomplish the following tasks: (a) Select the desired radio channel and reject other radio signals, (b) Amplify the desired radio signal and translate them back to baseband, and (c) Detect and decode the information with Low BER. In order to minimize cost and achieve high level of integration, receiver architecture should use least number of external filters and passive components. It should also consume least amount of power to minimize battery cost, size, and weight. One of the most stringent requirements for deep-space communication is the low-power operation. Our study identified that two candidate architectures listed in the following meet these requirements: (1) Low-IF receiver, (2) Sub-sampling receiver. The low-IF receiver uses minimum number of external components. Compared to Zero-IF (Direct conversion) architecture, it has less severe offset and flicker noise problems. The Sub-sampling receiver amplifies the RF signal and samples it using track-and-hold Subsampling mixer. These architectures provide low-power solution for the short- range communications missions on Mars. Accomplishments to date include: (1) System-level design and simulation of a Double-Differential PSK receiver, (2) Implementation of Honeywell SOI CMOS process design kit (PDK) in Cadence design tools, (3) Design of test circuits to investigate relationships between layout techniques, geometry, and low-frequency noise in SOI CMOS, (4) Model development and verification of on-chip spiral inductors in SOI CMOS process, (5) Design/implementation of low-power low-noise amplifier (LNA) and mixer for low-IF receiver, and (6) Design/implementation of high-gain LNA for sub-sampling receiver. Our initial results show that substantial improvement in power consumption is achieved using SOI CMOS as compared to standard CMOS process. Potential advantages of SOI CMOS for deep-space communication electronics include: (1) Radiation hardness, (2) Low-power operation, and (3) System-on-Chip (SOC) solutions.

Fujikawa, Gene↗

NASA NTRS2005

Multi-Mission System Architecture Platform: Design and Verification of the Remote Engineering Unit

The Multi-Mission System Architecture Platform (MSAP) represents an effort to bolster efficiency in the spacecraft design process. By incorporating essential spacecraft functionality into a modular, expandable system, the MSAP provides a foundation on which future spacecraft missions can be developed. Once completed, the MSAP will provide support for missions with varying objectives, while maintaining a level of standardization that will minimize redesign of general system components. One subsystem of the MSAP, the Remote Engineering Unit (REU), functions by gathering engineering telemetry from strategic points on the spacecraft and providing these measurements to the spacecraft's Command and Data Handling (C&DH) subsystem. Before the MSAP Project reaches completion, all hardware, including the REU, must be verified. However, the speed and complexity of the REU circuitry rules out the possibility of physical prototyping. Instead, the MSAP hardware is designed and verified using the Verilog Hardware Definition Language (HDL). An increasingly popular means of digital design, HDL programming provides a level of abstraction, which allows the designer to focus on functionality while logic synthesis tools take care of gate-level design and optimization. As verification of the REU proceeds, errors are quickly remedied, preventing costly changes during hardware validation. After undergoing the careful, iterative processes of verification and validation, the REU and MSAP will prove their readiness for use in a multitude of spacecraft missions.

Sartori, John↗

NASA NTRS2006

Certification Testing Approach for Propulsion System Design

The Certification of Propulsion Systems is costly and complex, involving development and qualification testing. The desire of the certification process is to assure all requirements can be demonstrated to be compliant. The purpose of this paper is to address the technical design concerns of certifying a propulsion system for flight. Presented are Pressurization, Tankage, Feed System and Combustion Instability concerns. Propulsion System Engineers are challenged with the dilemma for testing new systems to specific levels to reduce risk yet maintain budgetary targets. A methodical approach is presented to define the types of test suitable to address the technical issues for qualifying systems for retiring the risk levels. Experience of the lessons learned from supporting the Shuttle Program for Main Propulsion and On Orbit Propulsions Systems as well as previous collaborations on design concerns for certifying propulsion systems are utilized to address design concerns and verification approaches.

Rodriguez, Henry↗

NASA NTRS2010

Verification and Implementation of Operations Safety Controls for Flight Missions

Approximately eleven years ago, the International Space Station launched the first module from Russia, the Functional Cargo Block (FGB). Safety and Mission Assurance (S&MA) Operations (Ops) Engineers played an integral part in that endeavor by executing strict flight product verification as well as continued staffing of S&MA's console in the Mission Evaluation Room (MER) for that flight mission. How were these engineers able to conduct such a complicated task? They conducted it based on product verification that consisted of ensuring that safety requirements were adequately contained in all flight products that affected crew safety. S&MA Ops engineers apply both systems engineering and project management principles in order to gain a appropriate level of technical knowledge necessary to perform thorough reviews which cover the subsystem(s) affected. They also ensured that mission priorities were carried out with a great detail and success.

Jones, Cheryl L.↗