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This slide presentation reviews the importance of Orbital debris engineering models. They are mathematical tools to assess orbital debris flux. It briefly reviews the history of the orbital debris engineering models, and reviews the new features in the current model (i.e., ORDEM2010).
An Engineering Model (EM) 8 cm Ion Thruster Propulsion Subsystem was developed for operation at a thrust level 5 mN (1.1 mlb) at a specific impulse 1 sub sp = 2667 sec with a total system input power P sub in = 165 W. The system dry mass is 15 kg with a mercury-propellant-reservoir capacity of 8.75 kg permitting uninterrupted operation for about 12,500 hr. The subsystem can be started from a dormant condition in a time less than or equal to 15 min. The thruster has a design lifetime of 20,000 hr with 10,000 startup cycles. A gimbal unit is included to provide a thrust vector deflection capability of + or - 10 degrees in any direction from the zero position. The EM subsystem development program included thruster optimization, power-supply circuit optimization and flight packaging, subsystem integration, and subsystem acceptance testing including a cyclic test of the total propulsion package.
The Meteoroid Engineering Model (MEM) version 3 is NASA’s most current and accurate model of the meteoroid environment. MEM 3 supersedes all previous versions of MEM, including MEM Release 2.0 (MEMR2), MEM Release 1.0c (MEMR1c), and previously internally controlled and released versions of MEMCxP v2.0 and LunarMEM v2.0. Earlier versions of MEM superseded older models of the meteoroid environment such as the Grün model and its derivative, Technical Memo 4527 (hereafter abbreviated as TM 4527) [1]. Prior to the establishment of the NASA Meteoroid Environment Office (MEO), NASA’s meteoroid environment models relied on a simple empirical expression derived from [2], as described in [3] and later in [1]. This expression describes the meteoroid flux incident on a flat plate near 1 au. TM 4527 assumes an isotropic environment, making the orientation of the plate irrelevant [4]. The flux was combined with scale factors to account for the reduction in flux occurring when the Earth shields the spacecraft from a portion of the meteoroid environment and the enhancement in flux due to the focusing effect of Earth’s gravitational field. TM 4527 also introduced a crude, piecewise meteoroid speed distribution with an average velocity of 19 km/s for an orbiting spacecraft based on [5]. Finally, TM 4527 assumed a three-step density distribution in which dust particles smaller than 10−6 g have a density of 2 g/cu cm, micrometeoroids between 10−6 g and 0.01 g have a density of 1 g/cu cm, and meteoroids larger than 0.01 g have a density of 0.5 g/cm3. Thus, the meteoroid model presented in TM 4527 was assembled from multiple independent sources. The model of TM 4527 was also used for years in Space Station risk assessments, and is described in Space Station Specification (SSP) 30425.
A thermal vacuum test was performed on the engineering model of the outer planets atmospheric entry probe. Steady state runs at three simulated radioisotope heating unit loads and one transient run simulating the pre-entry power profile were made to determine the thermal characteristics of the engineering model. An analytic simulation of the model was correlated to the test data. Several steady state and one transient run were made with the model attached to the spacecraft adapter to determine the thermal interface between the model and the adapter.
The Controls and Dynamics Technology Branch at NASA Glenn Research Center are interested in combining a solid oxide fuel cell (SOFC) to operate in conjunction with a gas turbine engine. A detailed engine model currently exists in the Matlab/Simulink environment. The idea is to incorporate a SOFC model within the turbine engine simulation and observe the hybrid system's performance. The fuel cell will be heated to its appropriate operating condition by the engine s combustor. Once the fuel cell is operating at its steady-state temperature, the gas burner will back down slowly until the engine is fully operating on the hot gases exhausted from the SOFC. The SOFC code is based on a steady-state model developed by The U.S. Department of Energy (DOE). In its current form, the DOE SOFC model exists in Microsoft Excel and uses Visual Basics to create an I-V (current-voltage) profile. For the project's application, the main issue with this model is that the gas path flow and fuel flow temperatures are used as input parameters instead of outputs. The objective is to create a SOFC model based on the DOE model that inputs the fuel cells flow rates and outputs temperature of the flow streams; therefore, creating a temperature profile as a function of fuel flow rate. This will be done by applying the First Law of Thermodynamics for a flow system to the fuel cell. Validation of this model will be done in two procedures. First, for a given flow rate the exit stream temperature will be calculated and compared to DOE SOFC temperature as a point comparison. Next, an I-V curve and temperature curve will be generated where the I-V curve will be compared with the DOE SOFC I-V curve. Matching I-V curves will suggest validation of the temperature curve because voltage is a function of temperature. Once the temperature profile is created and validated, the model will then be placed into the turbine engine simulation for system analysis.
Pulse Detonation Engine Technology is currently being investigated at Glenn for both airbreathing and rocket propulsion applications. The potential for both mechanical simplicity and high efficiency due to the inherent near-constant-volume combustion process, may make Pulse Detonation Engines (PDE's) well suited for a number of mission profiles. Assessment of PDE cycles requires a simulation capability that is both fast and accurate. It should capture the essential physics of the system, yet run at speeds that allow parametric analysis. A quasi-one-dimensional, computational-fluid-dynamics-based simulation has been developed that may meet these requirements. The Euler equations of mass, momentum, and energy have been used along with a single reactive species transport equation, and submodels to account for dominant loss mechanisms (e.g., viscous losses, heat transfer, and valving) to successfully simulate PDE cycles. A high-resolution numerical integration scheme was chosen to capture the discontinuities associated with detonation, and robust boundary condition procedures were incorporated to accommodate flow reversals that may arise during a given cycle. The accompanying graphs compare experimentally measured and computed performance over a range of operating conditions for a particular PDE. Experimental data were supplied by Fred Schauer and Jeff Stutrud from the Air Force Research Laboratory at Wright-Patterson AFB and by Royce Bradley from Innovative Scientific Solutions, Inc. The left graph shows thrust and specific impulse, Isp, as functions of equivalence ratio for a PDE cycle in which the tube is completely filled with a detonable hydrogen/air mixture. The right graph shows thrust and specific impulse as functions of the fraction of the tube that is filled with a stoichiometric mixture of hydrogen and air. For both figures, the operating frequency was 16 Hz. The agreement between measured and computed values is quite good, both in terms of trend and magnitude. The error is under 10 percent everywhere except for the thrust value at an equivalence ratio of 0.8 in the left figure, where it is 14 percent. The simulation results shown were made using 200 numerical cells. Each cycle of the engine, approximately 0.06 sec, required 2.0 min of CPU time on a Sun Ultra2. The simulation is currently being used to analyze existing experiments, design new experiments, and predict performance in propulsion concepts where the PDE is a component (e.g., hybrid engines and combined cycles).
Development of a 30-cm mercury electron bombardment Engineering Model ion thruster has successfully brought the thruster from the status of a laboratory experimental device to a point approaching flight readiness. This paper describes the development progress of the Engineering Model (EM) thruster in four areas: (1) design features and fabrication approaches, (2) performance verification and thruster to thruster variations, (3) structural integrity, and (4) interface definition. The design of major subassemblies, including the cathode-isolator-vaporizer (CIV), main isolator-vaporizer (MIV), neutralizer isolator-vaporizer (NIV), ion optical system, and discharge chamber/outer housing is discussed along with experimental results.
The topics are presented in viewgraph form and include the following: nuclear thermal propulsion (NTP) & detailed nuclear engine modeling; modeling and engineering simulation of nuclear thermal rocket systems; nuclear thermal rocket simulation system; INSPI-NTVR core axial flow profiles; INSPI-NTRV core axial flow profiles; specific impulse vs. chamber pressure; turbine pressure ratio vs. chamber pressure; NERVA core axial flow profiles; P&W XNR2000 core axial flow profiles; pump pressure rise vs. chamber pressure; streamline of jet-induced flow in cylindrical chamber; flow pattern of a jet-induced flow in a chamber; and radiative heat transfer models.
The design, fabrication, and testing of a radiative cooler are described. This cooler is an engineering model suitable for bench testing in the laboratory as a part of the 10-micrometer wavelength engineering model receiver, and conforms to the standard radiative cooler configuration, except that the inner stage and its support system were redesigned to accommodate the larger, heavier SAT detector. This radiative cooler will cool the detector to cryogenic temperature levels when the receiver is in a space environment or in a suitable thermal vacuum chamber. Equipment specifications are given along with the results of thermal tests, vibration tests, and electrical integrity tests.
The Routh approximation technique is applied in the frequency domain to a 16th order state variable turbofan engine model. The results obtained motivate the extension of the frequency domain formulation of the Routh method to the time domain to handle the state variable formulation directly. The time domain formulation is derived, and a characterization, which specifies all possible Routh similarity transformations, is given. The characterization is computed by the solution of two eigenvalue eigenvector problems. The application of the time domain Routh technique to the state variable engine model is described, and some results are given.
The Routh approximation technique is applied in the frequency domain to a 16th order state variable turbofan engine model. The results obtained motivate the extension of the frequency domain formulation of the Routh method to the time domain to handle the state variable formulation directly. The time domain formulation is derived and a new characterization, which specifies all possible Routh similarity transformations, is given. This characterization is computed by the solution of two eigenvalue-eigenvector problems. The application of the time domain Routh technique to the state variable engine model is described and some results are given. Additional computational problems are discussed including an optimization procedure which can improve the approximation accuracy by taking advantage of the transformation characterization.
The NASA Orbital Debris Program Office at Johnson Space Center has developed a new computer-based orbital debris engineering model, ORDEM2000, which describes the orbital debris environment in the low Earth orbit region between 200 and 2000 km altitude. The model is appropriate for those engineering solutions requiring knowledge and estimates of the orbital debris environment (debris spatial density, flux, etc.). ORDEM2000 can also be used as a benchmark for ground-based debris measurements and observations. We incorporated a large set of observational data, covering the object size range from 10 mm to 10 m, into the ORDEM2000 debris database, utilizing a maximum likelihood estimator to convert observations into debris population probability distribution functions. These functions then form the basis of debris populations. We developed a finite element model to process the debris populations to form the debris environment. A more capable input and output structure and a user-friendly graphical user interface are also implemented in the model. ORDEM2000 has been subjected to a significant verification and validation effort. This document describes ORDEM2000, which supersedes the previous model, ORDEM96. The availability of new sensor and in situ data, as well as new analytical techniques, has enabled the construction of this new model. Section 1 describes the general requirements and scope of an engineering model. Data analyses and the theoretical formulation of the model are described in Sections 2 and 3. Section 4 describes the verification and validation effort and the sensitivity and uncertainty analyses. Finally, Section 5 describes the graphical user interface, software installation, and test cases for the user.
Results are presented from measurements of unsteady loads during performance tests of a generic high-speed engine model, which were made using high-frequency pressure gages installed in existing calorimeter ports of the engine model and recessed into the interior wall surface in order to reduce thermal flux to the gage diaphragm. In was found that the boundary layer pressure spectra at the model wall start to deviate from their flat plate counterpart at a short distance into the model inlet, which suggests the contributions to the spectra from the shock/boundary layer interaction. It was also found that significant levels of combustion noise propagate up through the subsonic portion of the boundary layer well into the inlet region. At the condition of an unstart, the combustion noise apparently couples with the acoustic modes of the model to cause acoustic 'hot spots' well upstream of the combustor.
The concept employed in an existing Trajectory Control Sensor (TCS) breadboard is being developed into an engineering model to be considered for flight on the Shuttle as a Detailed Test Objective (DTO). The sensor design addresses the needs of Shuttle/SSF docking/berthing by providing relative range and range rate to 1500 meters as well as the perceived needs of AR&C by relative attitude measurement over the last 100 meters. Range measurement is determined using a four-tone ranging technique. The Doppler shift on the highest frequency tone will be used to provide direct measurement of range rate. Bearing rate and attitude rates will be determined through back differencing of bearing and attitude, respectively. The target consists of an isosceles triangle configuration of three optical retroreflectors, roughly one meter and one-half meter in size. After target acquisition, the sensor continually updates the positions of the three retros at a rate of about one hertz. The engineering model is expected to weigh about 25 pounds, consume 25-30 watts, and have an envelope of about 1.25 cubic feet. The following concerns were addressed during the presentation: are there any concerns with differentiating attitude and bearing to get attitude and bearing rates? Since the docking scenario has low data bandwidth, back differencing is a sufficient approximation of a perfect differentiator for this application. Could range data be obtained if there were no retroreflectors on the target vehicle? Possibly, but only at close range. It would be dependent on target characteristics.
Development of an Engineering Model 8-cm Mercury Ion Thruster System for Satellite Control has been successfully completed. This system operates at a specific impulse in excess of 2600 sec, produces a thrust of 5 mN with a total input power of 165 W; it has a dry mass of 16.6 kg and a mercury-propellant-reservoir capacity of 8.75 kg. This paper summarizes the development work, the system characteristics and performance, and the testing undertaken to verify the design.
The NASA Orbital Debris Engineering Model (ORDEM) 3.1 Model Verification and Validation (V&V) document accompanies the delivery of the latest ORDEM 3.1 model (Vavrin & Manis, 2019) and provides a detailed description of the V&V activities used to verify that the model was built correctly and validate the model against independent, real world sources of data obtained from sampling the orbital debris environment. This ORDEM 3.1 Model V&V document, along with the related ORDEM 3.1 Model Process document – which covers details of the mathematical, statistical, and physical basis of the model – are intended to inform credibility assessments, risk analyses, uncertainty characterizations, and other applications derived from use of the model by the ORDEM 3.1 user community.
The HEMAP (Human Engineering Modeling and Performance) Lab is a joint effort between the Industrial and Human Engineering group and the KAVE (Kennedy Advanced Visualiations Environment) group. The lab consists of sixteen camera system that is used to capture human motions and operational tasks, through te use of a Velcro suit equipped with sensors, and then simulate these tasks in an ergonomic software package know as Jac, The Jack software is able to identify the potential risk hazards.
Various models for free piston Stirling engines are reviewed. Initial models were developed primarily for design purposes and to predict operating parameters, especially efficiency. More recently, however, such models have been used to predict engine stability. Free piston Stirling engines have no kinematic constraints and stability may not only be sensitive to the load, but also to various nonlinear loss and spring constraints. The present understanding is reviewed of various loss mechanisms for free piston Stirling engines and how they have been incorporated into engine models is discussed.