Houston, TX, United States

Odyssey Space Research

www.odysseysr.com
Houston, TX, United States

Odyssey Space Research, LLC is a small business based in Houston, Texas near NASA Lyndon B. Johnson Space Center providing engineering research and analysis services. This start-up in the space industry founded in November 2003 has already won major contracts and is the only private company working on the 5 next human-rated spacecraft . Wikipedia.


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Lu P.,San Diego State University | Brunner C.W.,Odyssey Space Research | Stachowiak S.J.,NASA | Mendeck G.F.,NASA | And 2 more authors.
Journal of Guidance, Control, and Dynamics | Year: 2017

The process, methodology, and results of a two year effort are presented in this paper on verification of an advanced entry guidance algorithm, called Fully Numerical Predictor-corrector Entry Guidance (FNPEG). FNPEGis a modelbased numerical guidance algorithm capable of performing both direct (orbital or suborbital) entry and skip entry missions. Few vehicle-dependent adjustments are necessary, and no reference trajectory or mission-dependent planning is required. The algorithm is applicable to a wide range of vehicles with different lift-to-drag ratios and includes state-of-the-art capability to effectively control g load and damp out phugoid oscillations, without adversely affecting the guidance precision. FNPEG has undergone extensive testing and evaluation in the high-fidelity simulation environment for the Orion spacecraft at NASA Johnson Space Center. In this paper, the verification methodology and process are described. The metrics for verification are defined. Extensive testing and simulation results onFNPEGand the comparison with the primary entry guidance algorithm for Orion, PredGuid, are provided. The outcome of this effort has clearly demonstrated the capability, strong robustness, and excellent performance of FNPEG, even in the presence of dispersions and uncertainties significantly higher than the design level.


Brunner C.W.,Iowa State University | Brunner C.W.,Odyssey Space Research | Lu P.,Iowa State University | Lu P.,Shanghai JiaoTong University
AIAA Guidance, Navigation, and Control Conference | Year: 2010

The dramatic increase in computational power since the Apollo program has enabled the development of numerical predictor-corrector (NPC) entry guidance algorithms that allow on-board accurate determination of a vehicle's trajectory. These algorithms are sufficiently mature to be flown. NPC algorithms are highly adaptive, especially in the face of extreme dispersion and off-nominal situations. The performance and reliability of entry guidance are critical to successful missions. This paper compares the performance of a recently developed NPC skip entry guidance algorithm with that of the Apollo skip entry guidance. Through extensive dispersion testing, it is clearly demonstrated that the Apollo skip entry guidance algorithm would be severely inadequate in meeting the landing precision requirement for missions with medium and long ranges under moderate dispersions. In the presence of large dispersions, a significant number of failures occur even for short-range missions. The NPC algorithm, on the other hand, is able to ensure high landing precision in all cases tested. Copyright © 2010 by Christopher Brunner and Ping Lu.


Brunner C.W.,Odyssey Space Research | Lu P.,Iowa State University
Journal of the Astronautical Sciences | Year: 2012

The dramatic increase incomputational power since the Apollo program has enabled the development of numerical predictor-corrector (NPC) entry guidance algorithms that allow on-board accurate determination of a vehicle's trajectory. These algorithms are sufficiently mature to be flown. They are highly adaptive, especially in the face of extreme dispersion and off-nominal situationscompared with reference-trajectory following algorithms. The performance and reliability of entry guidance are critical to mission success. This papercompares the performance of a recently developed fully numerical predictor-corrector entry' guidance (FNPEG) algorithm with that of the Apollo skip entry guidance. Through extensive dispersion testing, it is clearly demonstrated that the Apollo skip entry guidance algorithm would be inadequate in meeting the landing precision requirement for missions with medium (4000-7000 km) and long (>7000 km) downrange capability requirements under moderate dispersions chiefly due to poor modeling of atmospheric drag. In the presence of large dispersions, a significant number of failures occur even for short-range missions due to the deviation from planned reference trajectories. The FNPEG algorithm, on the other hand, is able to ensure high landing precision in all cases tested. All factors considered, a strong case is made for adopting fully numerical algorithms for future skip entry missions.


Thompson B.F.,Odyssey Space Research
Journal of Guidance, Control, and Dynamics | Year: 2011

The article presents a simple method for enhancing existing Lambert targeting methods by using the hodograph to compute velocity for any transfer angle. The 180° transfer orbit can be constrained to lie in the plane defined by these two vectors. The result is no orbit plane change for the transfer, which has the advantage of minimizing spacecraft propellant consumption in most cases. The velocity hodograph is the locus of the tip of the velocity vector drawn in the spacecraft-centered radial and transverse coordinate plane. The origin is inside the hodograph circle for hyperbolic orbits. The geometrical properties and associated equations are the same for hodographs of all orbit types. The minimum energy transfer is not a true Lambert Problem because the time of flight is dependent on the specific transfer conditions. However, the minimum energy transfer is fundamental to many astrodynamics applications, and can be solved by the hodograph method.


Goodman J.L.,Odyssey Space Research
Advances in the Astronautical Sciences | Year: 2016

The 1950s era reference trajectory and correlated guidance techniques (Delta Minimum, Delta, and Q) were not capable of supporting space missions envisioned by 1960. The development of digital flight computers enabled explicit guidance algorithms to be developed using results from the calculus of variations. The Iterative Guidance Mode (IGM) and E Guidance were explicit guidance schemes that were successfully developed for and flown in the Apollo Program. Hypersurface targeting provided constraints to IGM for the Trans Lunar Injection burn. Powered Explicit Guidance (PEG) was developed later and successfully flew on the Space Shuttle from 1981 to 2011. PEG was a more capable algorithm that could support demanding Space Shuttle abort profiles.


Goodman J.L.,Odyssey Space Research
Advances in the Astronautical Sciences | Year: 2016

Dr. Robert Goddard performed the first successful flight demonstration of gyroscopically controlled vertical rocket flight on March 28, 1935 near Roswell, New Mexico. German Army research into missile guidance started in the early 1930s and resulted in the LEV-3 system that was flown on the V-2 (A-4) during World War II. The LEV-3 and other developments at Peenemünde led directly to guidance systems developed in Huntsville, Alabama for the Redstone, Jupiter, Pershing I, and Saturn vehicles.


Senent J.S.,Odyssey Space Research
Advances in the Astronautical Sciences | Year: 2010

If a pure numerical iterative approach is used, targeting entry interface (EI) conditions for nominal and abort return trajectories or for correction maneuvers can be computationally expensive. This paper describes an algorithm to obtain an optimal impulsive maneuver that generates a trajectory satisfying a set of EI targets: inequality constraints on longitude, latitude and azimuth and a fixed flight-path angle. Most of the calculations require no iterations, making it suitable for real-time applications or large trade studies. This algorithm has been used to generate initial guesses for abort trajectories during Earth-Moon transfers.


Williams J.,Inc. Engineering and Science Contract Group | Senent J.S.,Odyssey Space Research | Lee D.E.,NASA
Advances in the Astronautical Sciences | Year: 2012

Copernicus is a software tool for advanced spacecraft trajectory design and optimization. The latest version (v3.0.1) was released in October 2011. It is available at no cost to NASA centers, government contractors, and organizations with a contractual affiliation with NASA. This paper provides a brief overview of the recent development history of Copernicus. An overview of the evolution of the software is given, along with a discussion of significant new features and improvements (such as gravity assist maneuvers, halo orbits, and a new impulsive to finite burn conversion wizard). Some examples of how Copernicus is used to design spacecraft missions are also shown.


Grant
Agency: National Aeronautics and Space Administration | Branch: | Program: SBIR | Phase: Phase I | Award Amount: 99.98K | Year: 2011

Odyssey Space Research proposes to develop a modular navigation software package to provide precise state information for offline analysis and real-time applications. This navigation package will use particle filter methodology to process discrete observation data and maintain an accurate state. This navigation system will leverage several NASA products to rapidly prototype and demonstrate the feasibility of this software during Phase I, including the General Mission Analysis Tool (GMAT) and Trick, taking it from TRL 2 past TRL 3. Phase II will deliver an expanded modular software product integrated into several other software packages demonstrating different estimation capabilities (TRL 5-6). This system will function as a standalone estimation package that can be easily integrated into other software packages, or as the basis for embedded flight software algorithms.This navigation package will be designed to meet the position, velocity, and time estimation requirements for space missions. It will contain an expanded state vector used to estimate non-Gaussian forcing functions perturbing the vehicle's dynamics. This navigator will integrate the measurements from diverse sensors running at different rates. And it will demonstrate accurate estimation of uncertain dynamics parameters that are affecting the vehicle's state such as the gravitation field of small bodies.


Grant
Agency: National Aeronautics and Space Administration | Branch: | Program: SBIR | Phase: Phase I | Award Amount: 125.00K | Year: 2015

Odyssey proposes a new fault management planning and design tool and methodology that uses state-based simulations with programmable dynamic state definitions to provide early assessments of fault management system scope and cost. The tool will utilize models developed in SysML to capture system characteristics and relationships between system components as well as mapping of functionality to requirements and mission objectives, and a probabilistic state-based simulation to determine requirements compliance, fault probabilities, and the number of fault paths in the system. The tool will provide useful visualization of the FM design and fault paths, including the dynamic aspect of the system, as well as visual representations of system complexity. In addition, the tool will provide automated means to estimate the complexity of the FM design based on system characteristics and simulation results. The tool will provide engineers and managers with the ability to scope the fault management (FM) effort from requirements development through verification at a point early in the design process. Phase I will focus on proof of concept and demonstration of key aspects of the tool, with full tool development and scaling to complex systems in Phase II.

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