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Recent Radar Observations of the Sub-Centimeter Orbital Debris Environment
Get PDFThe NASA Orbital Debris Program Office (ODPO) has conducted radar observations of the orbital debris environment since the early 1990s to provide measurement data that supports orbital debris models and risk mitigation activities in support of NASA mission objectives. Orbital debris radar observations are a unique mode for radar operation, employing a fixed beam configuration to statistically sample the environment. An advantage of conducting operations in this fashion is that it enables observations of smaller classes of orbital debris than would otherwise be available from the same sensor operating in a traditional tracking mode. Orbital debris-mode radar observations are used to fill in the gaps, which exist in the currently available data from the Space Surveillance Network (SSN), on small size orbital debris populations that represent significant risk to NASA programs. These gaps have typically covered orbital debris with characteristic sizes less than approximately 10 cm down to approximately 3 mm in low Earth orbit (LEO) depending upon the altitude and sensor configuration. The value of orbital debris radar measurements lies in the ability to extract partial orbital element information about orbital debris in the centimeter to several millimeter size regimes in low Earth orbit which are not available from other measurement sources. This paper will discuss observations of this smaller class of orbital debris observed in recent years from the radars at the MIT Haystack Observatory in Westford, Massachusetts, and the Goldstone Solar System Radar near Barstow, California. The former radar is able to observe orbital debris down to approximately 5 mm, and the latter, orbital debris with characteristic sizes near 3 mm at altitudes less than 1000 km. The characteristics and inferences about the current LEO orbital debris environment, and the different subpopulations that are identifiable in the observations are highlighted
Lightning Density and Thunderstorm Initiation in the Lake Victoria Region in East Africa
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X-56A Structural Dynamics Ground Testing Overview and Lessons Learned
Get PDFThe X-56A Multi-Utility Technology Testbed (MUTT) is a subscale, fixed-wing aircraft designed for high-risk aeroelastic flight demonstration and research. Structural dynamics ground testing for model validation was especially important for this vehicle because the structural model was directly used in the development of a flight control system with active flutter suppression capabilities. Structural dynamics ground tests of the X-56A MUTT with coupled rigid-body and structural modes provided a unique set of challenges. An overview of the ground vibration test (GVT) and moment of inertia (MOI) test setup and execution is presented. The series of GVTs included the wing by itself attached to a strongback and complete vehicle at two mass conditions: empty and full fuel. Two boundary conditions for the complete-vehicle test were studied: on landing gear and suspended free-free. Pitch MOI tests were performed using a compound pendulum method and repeated with two different pendulum lengths for independent verification. The original soft-support test configuration for the GVT used multiple bungees, resulting in unforeseen coupling interactions between the soft-support bungees and the vehicle structural modes. To resolve this problem, the soft-support test setup underwent multiple iterations. The various GVT configurations and boundary-condition modifications are highlighted and explained. Lessons learned are captured for future consideration when performing structural dynamics testing with similar vehicles
Modeling Heatshield Erosion Due to Dust Particle Impacts for Martian Entries
Get PDFHeatshield design for spacecraft entering the atmosphere of Mars may be affected by the presence of atmospheric dust. Particle impacts with sufficient kinetic energy can cause spallation damage to the heatshield that must be estimated. The dust environment in terms of particle size distribution and number density can be inferred from ground-based or atmospheric observations at Mars. Using a Lagrangian approach, the particle trajectories through the shock layer can be computed using a set of coupled ordinary differential equations. The dust particles are small enough that non-continuum effects must be accounted for when computing the drag coefficient and heat transfer to the particle surface. Surface damage correlations for impact crater diameter and penetration depth are presented for fused-silica, AVCOAT, Shuttle tiles, cork, and Norcoat Lige. The cork and Norcoat Lige correlations are new and were developed in this study. The modeling equations presented in this paper are applied to compute the heatshield erosion due to dust particle impacts on the ExoMars Schiaparelli entry capsule during dust storm conditions
Application of Risk Informed Decision Making to Highly Reliable Three Dimensionally Woven Thermal Protection System for Mars Sample Return
Get PDFThe NASA Risk Informed Decision Making process is used to assess a trade space of three dimensionally woven thermal protection systems for application to the Mars Sample Return Earth Entry Vehicle. Candidate architectures are assessed based on mission assurance, technical development, cost, and schedule risk. Assessment methodology differed between the architectures, utilizing a four-point quantitative scale for mission assurance and technical development and highly tailored PERT techniques for cost and schedule. Risk results are presented, in addition to a review of RIDM effectiveness for this application
Mid Lift-to-Drag Ratio Rigid Vehicle 6-DoF Performance for Human Mars Entry, Descent, and Landing: A Fractional Polynomial Powered Descent Guidance Approach
Get PDFDefining a feasible vehicle design and mission architecture capable of reliably delivering apayload of 20 metric tons (mt) or more is a great challenge for landing humans on Mars. TheMid Lift-to-Drag Rigid Vehicle (MRV), a rigid decelerator studied in NASAs Entry, Descent,and Landing Architecture Study (EDLAS), has shown to be a viable vehicle candidate forfuture human Mars missions. As the vehicle concept matures, models of increasing fidelity areadded to the six-degree-of-freedom (6DoF) EDL simulation. This paper presents 6DoFsimulation results using model updates for vehicle mass properties, fineness ratio, andaerodynamic-propulsive interactions. Additionally, an assessment of the Fractional-Polynomial Powered Descent Guidance (FP2DG) performance is presented, and the vehicleperformance is compared with the Tunable Apollo Powered Descent Guidance (TAPDG).Finally, Monte Carlo results of the vehicle design trades are presented
Preliminary Computational Assessment of Disk Rotating Detonation Engine Configurations
Get PDFA rotating detonation engine (RDE) configuration whereby the working fluid enters and exits in a predominantly radial manner is examined using a quasi-two-dimensional computational fluid dynamic simulation. The simulation, based on a Cartesian coordinate system, was originally developed to examine the physics and performance of the more typical annular RDE. Modifications required to accommodate the radial and circumferential flowfield are discussed. The centripetal forces that arise in this disk RDE (DRDE) configuration create a different wave structure than that seen in the annular RDE. They also give rise to markedly different fluid behavior depending on whether the flow is radially inward or radially outward. Using an entropy-based measure of pressure gain, it is found that for the preliminary idealized calculations performed in this paper, the inward flowing DRDE outperforms the outward flowing variant. The inward flowing DRDE is further shown to outperform the equivalent annular RDE. The effects on performance of several parameters are examined, including inner-to-outer diameter ratio, inner-to-outer cross-sectional area ratio, and inlet throat-to-channel area ratio
Free Space Optical Link Utilizing a Modulated Retro-Reflector Intended for Planetary Duplex Communication Links Between an Orbiter and Surface Unit
Get PDFPresented are simulation and experimental results that provide duplex optical-free space communication links with minimal power and pointing requirements by using a modulated retro-reflector (MRR) for planetary communications. The design is the MRR resides on the surface of a planet or moon, where energy is scarce, while the source of the communication laser resides on an orbiter to achieve satellite-to-ground communications. Also, a simulated scenario using the Mars Reconnaissance Orbiter (MRO) is provided for real world potential results. The information sent through this communication path can range from raw scientific data to multimedia files such as videos and pictures. Bidirectional communications is established with the MRR by using a nested pulse position modulation (PPM) structure. This modulation scheme is then evaluated for its validity in a proof-of-concept experiment. Initial results indicate a promising return-link performance of at least 300 kbps in the nested arrangement
Pterodactyl: Control System Design for Deployable Entry Vehicles
Get PDFAs space exploration missions continue to get more complex, greater demands for mass and volume are being placed on planetary entry vehicles. At the same time, the desire for precision entry targeting to enable surface missions at very specific locations is becoming common. Deployable aeroshells are one type of technology that is being developed to increase the down-mass capability of planetary entry vehicles while complying with current launch vehicle volume constraints. Propulsive control can be used for precision entry targeting but can require a large percentage of the vehicle's total mass or volume for propellant. The Pterodactyl project is studying alternative entry control systems on a Deployable Entry Vehicle to potentially introduce a mass and volume efficient entry system that can achieve precision targeting. This paper presents the mechanical design work undertaken to support the Pterodactyl trade study of three entry vehicle control system options: 1) Flaps, 2) Mass Movement, and 3) Reaction Control System
Thermal Evolution and Core Formation on Asteroid 4 Vesta in the Magma Ocean Regime
Get PDFGeochemical observations of the eucrite and diogenite meteorites, together with observations made by NASAs Dawn spacecraft while orbiting asteroid 4 Vesta, indicate that Vesta has differentiated to form a crust, mantle, and core. Eucrite and diogenite petrology is best explained by solidification of the crust from a magma ocean constituting 60-70% of Vestas silicates [3], or a temperature of ~1550 C. The abundances of moderately siderophile elements (Ni, Co, Mo, W, and P) in eucrites require that essentially all of the metallic phase in Vesta segregated to form a core prior to eucrite formation and likely reached a temperature of 1450- 1575 C. These observations provide important constraints on Vestas thermal evolution. The high inferred temperature indicates that convective heat transport must have been important during part of Vestas thermal evolution. In this study, we model Vestas thermal evolution in the magma ocean regime