SEARCH · Engineering Papers

Results for “Engineering Model”

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.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 793 records · Page 44

All sources NASA NTRS DOE OSTI

NASA NTRS2016

Integration Tests of the 4 kW-class High Voltage Hall Accelerator Power Processing Unit with the HiVHAc and the SPT-140 Hall Effect Thrusters

NASAs Science Mission Directorate is sponsoring the development of a 4 kW-class Hall propulsion system for implementation in NASA science and exploration missions. The main components of the system include the High Voltage Hall Accelerator (HiVHAc), an engineering model power processing unit (PPU) developed by Colorado Power Electronics, and a xenon flow control module (XFCM) developed by VACCO Industries. NASA Glenn Research Center is performing integrated tests of the Hall thruster propulsion system. This presentation presents results from integrated tests of the PPU and XFCM with the HiVHAc engineering development thruster and a SPT-140 thruster provided by Space System Loral. The results presented in this paper demonstrate thruster discharge initiation, open-loop and closed-loop control of the discharge current with anode flow for both the HiVHAc and the SPT-140 thrusters. Integrated tests with the SPT-140 thruster indicated that the PPU was able to repeatedly initiate the thrusters discharge, achieve steady state operation, and successfully throttle the thruster between 1.5 and 4.5 kW. The measured SPT-140 performance was identical to levels reported by Space Systems Loral.

Hall Thrusters↗

NASA NTRSDate not supplied

Controls for Electrified Aircraft Propulsion

Advanced electrified aircraft propulsion (EAP) concepts with integrated power, propulsion, and thermal systems require the development of equally sophisticated controllers to fly safe and efficient missions. Hybrid electric aircraft utilize electric machines mechanically coupled to the engine shafts to extract and insert power for a variety of purposes. Additionally, electric machines may drive propulsive fans pulling from a combination of on-board energy storage devices and engine extracted power. NASA’s investment in hybrid electric aircraft hardware-in-the-loop testing enables controls research on representative engine models using a novel emulation and scaling methodology. High-voltage, high-power electrical powertrain presents several technical risks at altitude. High-power requires high efficiency to minimize losses. Superconducting electric machines and power distribution research at NASA has identified key challenges with potential controls solutions. These risks necessitate the development of controllers robust to model uncertainty, disturbances, and fault conditions. System health management and fault detection schemes play an important role in a multi-layer controls approach that utilizes a supervisor which oversees inner loops for the highly coupled power, propulsion, and thermal systems. Such a supervisory controller enables coordinates energy transfer between engines, energy storage devices, electric machines, propulsors, and heat exchangers. Existing methods have shown an improvement in engine operability using the hybrid electric powertrain.

Halle E Buescher↗

NASA NTRS2010

A Comparison of Flow-Through Versus Non-Flow-Through Proton Exchange Membrane Fuel Cell Systems for NASA's Exploration Missions

As part of the Exploration Technology Development Program (ETDP) under the auspices of the Exploration Systems Mission Directorate (ESMD), NASA is developing both primary fuel cell power systems and regenerative fuel cell (RFC) energy storage systems within the fuel cell portion of the Energy Storage Project. This effort is being led by the NASA Glenn Research Center (GRC) in partnership with the NASA Johnson Space Center (JSC), Jet Propulsion Laboratory (JPL), NASA Kennedy Space Center (KSC), and industrial partners. The development goals are to improve fuel cell and electrolysis stack electrical performance, reduce system mass, volume, and parasitic power requirements, and increase system life and reliability. A major focus of this effort has been the parallel development of both flow-through and non-flow-through proton exchange membrane (PEM) primary fuel cell power systems. The plan has been, at the appropriate time, to select a single primary fuel cell technology for eventual flight hardware development. Ideally, that appropriate time would occur after both technologies have achieved a technology readiness level (TRL) of six, which represents an engineering model fidelity PEM fuel cell system being successfully tested in a relevant environment. Budget constraints in fiscal year 2009 and beyond have prevented NASA from continuing to pursue the parallel development of both primary fuel cell options. Because very limited data exists for either system, a toplevel, qualitative assessment based on engineering judgement was performed expeditiously to provide guidance for a selection. At that time, the non-flow-through technology was selected for continued development because of potentially major advantages in terms of weight, volume, parasitic power, reliability, and life. This author believes that the advantages are significant enough, and the potential benefits great enough, to offset the higher state of technology readiness of flow-through technology. This paper summarizes the technical considerations which helped form the engineering judgement that led to the final decision.

Hoberecht, Mark A.↗

NASA NTRSDate not supplied

GASP Advancements With Symbolic Computations, Optimization, and Decoupled Numerical Methods

NASA is exploring advancements in hybrid-electric propulsion concepts for aircraft design. A new analysis capability has been deployed at NASA’s Ames Research Center using a new mathematical modeling framework in Python that leverages symbolic representation of mathematical expressions, optimization-based problem formulations, and standalone numerical methods. This capability, called Gascon, has been verified against NASA’s General Aviation Synthesis Program (GASP) results for several vehicles that exercise the analysis capability for a variety of aircraft classes and engine models, including both turbofans and turboprops. We believe this capability will set the stage for more rapid development of novel aircraft models with more flexible design assessments using optimization, and more traceable analysis for reproducibility. Additionally, a novel Pareto optimization-based design space exploration that shows the trade between range flown and fuel burn by manipulating the thrust split during cruise for a true parallel hybrid aircraft was developed by leveraging the unique features of Gascon.

aircraft design↗

NASA NTRS1976

Summary report on effects at temperature, humidity, and fuel-air ratio on two air-cooled light aircraft engines

Five different engine models were tested to experimentally characterize emissions and to determine the effects of variation in fuel-air ratio and spark timing on emissions levels and other operating characteristics such as cooling, misfiring, roughness, power acceleration, etc. The results are given of two NASA reports covering the Avco Lycoming 0-320-D engine testing and the recently obtained results on the Teledyne Continental TSIO-360-C engine.

Kempke, E. E., Jr.↗

NASA NTRS1991

Supersonic propulsion simulation by incorporating component models in the large perturbation inlet (LAPIN) computer code

An approach to simulating the internal flows of supersonic propulsion systems is presented. The approach is based on a fairly simple modification of the Large Perturbation Inlet (LAPIN) computer code. LAPIN uses a quasi-one dimensional, inviscid, unsteady formulation of the continuity, momentum, and energy equations. The equations are solved using a shock capturing, finite difference algorithm. The original code, developed for simulating supersonic inlets, includes engineering models of unstart/restart, bleed, bypass, and variable duct geometry, by means of source terms in the equations. The source terms also provide a mechanism for incorporating, with the inlet, propulsion system components such as compressor stages, combustors, and turbine stages. This requires each component to be distributed axially over a number of grid points. Because of the distributed nature of such components, this representation should be more accurate than a lumped parameter model. Components can be modeled by performance map(s), which in turn are used to compute the source terms. The general approach is described. Then, simulation of a compressor/fan stage is discussed to show the approach in detail.

Cole, Gary L.↗

NASA NTRS2021

Systems Engineering and Assurance Modeling (SEAM): A Web-Based Solution for Integrated Mission Assurance

We present an overview of the Systems Engineering and Assurance Modeling (SEAM) platform, a web-browser-based tool which is designed to help engineers evaluate the radiation vulnerabilities and develop an assurance approach for electronic parts in space systems. The SEAM framework consists of three interconnected modeling tools, a SysML compatible system description tool, a Goal Structuring Notation (GSN) visual argument tool, and Bayesian Net and Fault Tree extraction and export tools. The SysML and GSN sections also have a coverage check application that ensures that every radiation fault identified on the SysML side is also addressed in the assurance case in GSN. The SEAM platform works on space systems of any degree of radiation hardness but is especially helpful for assessing radiation performance in systems with commercial-off-the-shelf (COTS) electronic components.

SEAM↗

NASA NTRS1998

The Marshall Engineering Thermosphere (MET) Model: Technical Description - Volume 1

Volume 1 presents a technical description of the Marshall Engineering Thermosphere (MET) model atmosphere and a summary of its historical development. Various programs developed to augment the original capability of the model are discussed in detail. The report also describes each of the individual subroutines developed to enhance the model. Computer codes for these subroutines are contained in four appendices.

Smith, R. E.↗

NASA NTRS2011

Model-Based Systems Engineering for Capturing Mission Architecture System Processes with an Application Case Study - Orion Flight Test 1

Model-based Systems Engineering (MBSE) is an emerging methodology that can be leveraged to enhance many system development processes. MBSE allows for the centralization of an architecture description that would otherwise be stored in various locations and formats, thus simplifying communication among the project stakeholders, inducing commonality in representation, and expediting report generation. This paper outlines the MBSE approach taken to capture the processes of two different, but related, architectures by employing the Systems Modeling Language (SysML) as a standard for architecture description and the modeling tool MagicDraw. The overarching goal of this study was to demonstrate the effectiveness of MBSE as a means of capturing and designing a mission systems architecture. The first portion of the project focused on capturing the necessary system engineering activities that occur when designing, developing, and deploying a mission systems architecture for a space mission. The second part applies activities from the first to an application problem - the system engineering of the Orion Flight Test 1 (OFT-1) End-to-End Information System (EEIS). By modeling the activities required to create a space mission architecture and then implementing those activities in an application problem, the utility of MBSE as an approach to systems engineering can be demonstrated.

Orion Flight Test 1 (OFT-1)↗

NASA NTRS1996

Heuristic Modeling for TRMM Lifetime Predictions

Analysis time for computing the expected mission lifetimes of proposed frequently maneuvering, tightly altitude constrained, Earth orbiting spacecraft have been significantly reduced by means of a heuristic modeling method implemented in a commercial-off-the-shelf spreadsheet product (QuattroPro) running on a personal computer (PC). The method uses a look-up table to estimate the maneuver frequency per month as a function of the spacecraft ballistic coefficient and the solar flux index, then computes the associated fuel use by a simple engine model. Maneuver frequency data points are produced by means of a single 1-month run of traditional mission analysis software for each of the 12 to 25 data points required for the table. As the data point computations are required only a mission design start-up and on the occasion of significant mission redesigns, the dependence on time consuming traditional modeling methods is dramatically reduced. Results to date have agreed with traditional methods to within 1 to 1.5 percent. The spreadsheet approach is applicable to a wide variety of Earth orbiting spacecraft with tight altitude constraints. It will be particularly useful to such missions as the Tropical Rainfall Measurement Mission scheduled for launch in 1997, whose mission lifetime calculations are heavily dependent on frequently revised solar flux predictions.

Jordan, P. S.↗

NASA NTRSDate not supplied

A Simple Model for Rotating Detonation Rocket Engine Sizing and Performance Estimates

A Rotating Detonation Rocket Engine (RDRE) model is described which characterizes the device as an essentially infinite number of circumferentially arranged, sequentially firing pulse detonation engine (PDE) tubes. The PDE tubes are in turn treated as lumped-parameter chambers, each of which executes a dynamic (i.e. time dependent) Atkinson cycle with an allowance for a finite Mach number during the refill portion of the cycle. The formulation results in several free parameters which are used to essentially tune the model such that it closely matches higher fidelity computational fluid dynamic RDRE simulations in terms of performance, flow rates, etc. The simplicity of the model allows for straightforward combination with component models (e.g. turbopumps, cooling jackets, etc.) as well as implementation within larger vehicle simulations. Such capability allows for realistic mission analyses and benefits studies which will help to determine the best application for the RDRE concept.

detonation↗

NASA NTRSDate not supplied

A Simple Model for Rotating Detonation Rocket Engine Sizing and Performance Estimates

detonation↗

NASA NTRS2012

Variable-Speed Simulation of a Dual-Clutch Gearbox Tiltrotor Driveline

This investigation explores the variable-speed operation and shift response of a prototypical two-speed dual-clutch transmission tiltrotor driveline in forward flight. Here, a Comprehensive Variable-Speed Rotorcraft Propulsion System Modeling (CVSRPM) tool developed under a NASA funded NRA program is utilized to simulate the drive system dynamics. In this study, a sequential shifting control strategy is analyzed under a steady forward cruise condition. This investigation attempts to build upon previous variable-speed rotorcraft propulsion studies by 1) including a fully nonlinear transient gas-turbine engine model, 2) including clutch stick-slip friction effects, 3) including shaft flexibility, 4) incorporating a basic flight dynamics model to account for interactions with the flight control system. Through exploring the interactions between the various subsystems, this analysis provides important insights into the continuing development of variable-speed rotorcraft propulsion systems.

DeSmidt, Hans↗

NASA NTRS2012

Design and Fabrication of DebriSat - A Representative LEO Satellite for Improvements to Standard Satellite Breakup Models

This paper discusses the design and fabrication of DebriSat, a 50 kg satellite developed to be representative of a modern low Earth orbit satellite in terms of its components, materials used, and fabrication procedures. DebriSat will be the target of a future hypervelocity impact experiment to determine the physical characteristics of debris generated after an on-orbit collision of a modern LEO satellite. The major ground-based satellite impact experiment used by DoD and NASA in their development of satellite breakup models was SOCIT, conducted in 1992. The target used for that experiment was a Navy transit satellite (~40 cm, 35 kg) fabricated in the 1960's. Modern satellites are very different in materials and construction techniques than those built 40 years ago. Therefore, there is a need to conduct a similar experiment using a modern target satellite to improve the fidelity of the satellite breakup models. To ensure that DebriSat is truly representative of typical LEO missions, a comprehensive study of historical LEO satellite designs and missions within the past 15 years for satellites ranging from 1 kg to 5000 kg was conducted. This study identified modern trends in hardware, material, and construction practices utilized in recent LEO missions. Although DebriSat is an engineering model, specific attention is placed on the quality, type, and quantity of the materials used in its fabrication to ensure the integrity of the outcome. With the exception of software, all other aspects of the satellite s design, fabrication, and assembly integration and testing will be as rigorous as that of an actual flight vehicle. For example, to simulate survivability of launch loads, DebriSat will be subjected to a vibration test. As well, the satellite will undergo thermal vacuum tests to verify that the components and overall systems meet typical environmental standards. Proper assembly and integration techniques will involve comprehensive joint analysis, including the precise torqueing of fasteners and thread locking. Finally, the implementation of process documentation and verification procedures is discussed to provide a comprehensive overview of the design and fabrication of this representative LEO satellite.

Clark, S.↗

NASA NTRS1999

STS-93 SSME Nozzle Tube Rupture Investigation

STS-93 was launched on July 23, 1999. There was an anomaly at the end of the launch in that the main engines shut down 0.16 second early because sensors detected a low level of oxidizer in the LOX tank (actually the duct from the tank to the vehicle). This resulted in a cutoff velocity for the vehicle that was 16 ft/sec low. It should have been 25872 ft/sec. The OMS engines were subsequently used to achieve the proper orbit. An investigation was immediately initiated into the cause of this LOX tank low level cutoff. It was noticed during the launch that the turbine temperatures for Main Engine 3 (E2019) were approximately 100 F higher than the preflight prediction. Linear Engine Model matching of the data indicated that a nozzle leak best fit the data. Post launch review of the data showed, that at approximately five seconds into the start, numerous parameters indicated small anomalous shifts. These shifts were all consistent with a rupture of nozzle tubes. Post launch review of the films showed that just after Space Shuttle Main Engine (SSME) ignition and just prior to liftoff a streak is seen in the exhaust plume of E2019. Just after liftoff the streak can be seen emanating from the nozzle wall. This photo confirmed that a leak was coming from the nozzle tubes. Based on the photo, the axial location of the leak was estimated to be 28 in. from the aft end of the nozzle and in line with nozzle coolant feed line #1. The streak continued to be visible during the launch.

Romine, W. Dennis↗

NASA NTRS2008

Predicting Chandra CCD Degradation with the Chandra Radiation Model

Not long after launch of the Chandra X-Ray Observatory, it was discovered that the Advanced CCD Imaging Spectrometer (ACIS) detector was rapidly degrading due to radiation. Analysis by Chandra personnel showed that this degradation was due to 10w energy protons (100 - 200 keV) that scattered down the optical path onto the focal plane. In response to this unexpected problem, the Chandra Team developed a radiation-protection program that has been used to manage the radiation damage to the CCDs. This program consists of multiple approaches - scheduled sating of the ACIS detector from the radiation environment during passage through radiation belts, real-time monitoring of space weather conditions, on-board monitoring of radiation environment levels, and the creation of a radiation environment model for use in computing proton flux and fluence at energies that damage the ACIS detector. This radiation mitigation program has been very successful. The initial precipitous increase in the CCDs' charge transfer inefficiency (CTI) resulting from proton damage has been slowed dramatically, with the front-illuminated CCDS having an increase in CTI of only 2.3% per year, allowing the ASIS detector's expected lifetime to exceed requirements. This paper concentrates on one aspect of the Chandra radiation mitigation program, the creation of the Chandra Radiation Model (CRM). Because of Chandra's highly elliptical orbit, the spacecraft spends most of its time outside of the trapped radiation belts that present the severest risks to the ACIS detector. However, there is still a proton flux environment that must be accounted for in all parts of Chandra's orbit. At the time of Chandra's launch there was no engineering model of the radiation environment that could be used in the outer regions of the spacecraft's orbit, so the CRM was developed to provide the flux environment of 100 - 200 keV protons in the outer magnetosphere, magnetosheath, and solar wind regions of geospace. This presentation describes CRM, its role in Chandra operations, and its prediction of the ACIS CTI increase.

Minow, Joseph I.↗

NASA NTRS2015

Geometry and Simulation Results for a Gas Turbine Representative of the Energy Efficient Engine (EEE)

This paper describes the geometry and simulation results of a gas-turbine engine based on the original EEE engine developed in the 1980s. While the EEE engine was never in production, the technology developed during the program underpins many of the current generation of gas turbine engines. This geometry is being explored as a potential multi-stage turbomachinery test case that may be used to develop technology for virtual full-engine simulation. Simulation results were used to test the validity of each component geometry representation. Results are compared to a zero-dimensional engine model developed from experimental data. The geometry is captured in a series of Initial Graphical Exchange Specification (IGES) files and is available in a supplement to this report.

Gas turbine engines↗

NASA NTRS2015

Geometry and Simulation Results for a Gas Turbine Representative of the Energy Efficient Engine (EEE)

Gas turbine engines↗