ADEPT, A Mechanically Deployable Re-Entry Vehicle System, Enabling Interplanetary CubeSat and Small Satellite Missions

There is growing interest for utilizing Small Satellites beyond low Earth orbit. A number of secondary CubeSat payload missions are planned at Mars, cis-Lunar Space, near Earth objects, and moons of the Gas Giants. Use of smaller systems may enable utilization of otherwise unused capacity of larger “host” missions. Development of re-entry systems that leverage and accommodate Small Satellite technology will substantially expand the range of mission applications by offering the capability for high speed entry or aerocapture at destinations with atmospheres. Deployable entry vehicles (DEVs) offer benefits over traditional rigid aeroshells including volume, mass and payload form factor. The Adaptive Deployable Entry and Placement Technology (ADEPT) offers such a delivery capability for Small Sat or CubeSat orbiter(s), in-situ elements, or landers. The ADEPT system can package with off the shelf CubeSat deployment systems (1U-16U) to offer a delivery capability for a single CubeSat or constellations. Furthermore, ADEPT can deliver the same science payload to a destination with a stowed diameter a factor of 3-4 times smaller than an equivalent rigid aeroshell, alleviating volumetric constraints on the secondary payload accommodation or primary carrier spacecraft bus. This paper will describe ADEPT’s current development status and define various interplanetary mission concepts in order to provide guidelines for potential Small Satellite payload developers and mission implementers

Enabling and Enhancing Science Exploration Across the Solar System: Aerocapture Technology for SmallSat to Flagship Missions

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Implementing Nonlinear Buoyancy and Excitation Forces in the WEC-Sim Wave Energy Converter Modeling Tool

Wave energy converters (WECs) are commonly designed and analyzed using numerical models that combine multibody dynamics with hydrodynamic models based on the Cummins equation and linearized hydrodynamic coefficients. These modeling methods are attractive design tools because they are computationally inexpensive and do not require the use of high-performance computing resources necessitated by high-fidelity methods, such as Navier-Stokes computational fluid dynamics. Modeling hydrodynamics using linear coefficients assumes that the device undergoes small motions and that the wetted surface area of the devices is approximately constant. WEC devices, however, are typically designed to undergo large motions to maximize power extraction, calling into question the validity of assuming that linear hydrodynamic models accurately capture the relevant fluid-structure interactions. In this paper, we study how calculating buoyancy and Froude-Krylov forces from the instantaneous position of a WEC device changes WEC simulation results compared to simulations that use linear hydrodynamic coefficients. First, we describe the WEC-Sim tool used to perform simulations and how the ability to model instantaneous forces was incorporated into WEC-Sim. We then use a simplified one-body WEC device to validate the model and to demonstrate how accounting for these instantaneously calculated forces affects the accuracy of simulation results, such as device motions, hydrodynamic forces, and power generation. Other aspects of WEC-Sim code development and verification are presented in a companion paper [1] that is also being presented at OMAE2014

Development of an Earth Smallsat Flight Test to Demonstrate Viability of Mars Aerocapture

Aerocapture has long been considered a compelling orbital maneuver that significantly reduces the cost of a wide variety of Mars orbital missions, as well as potential missions to Venus, Neptune, Titan, and return to Earth. Numerous conceptual studies have advanced the technical maturity of aerocapture to help enable its use on missions to different atmospheric worlds. Despite this technological readiness, the lack of an integrated aerocapture flight system demonstration has often been cited as rationale for not employing the technique on a flight mission. In order to facilitate a potential flight test, recent research has focused on the use of drag modulation-based techniques to greatly reduce the complexity of aerocapture systems. Drag modulation systems utilize changes in a vehicle\u27s drag area during flight to effect control over the vehicle\u27s ballistic coefficient, and therefore its final trajectory. Compared to traditional bank-to-steer lifting methods, these techniques enable use of extremely simple avionics algorithms, sensors and actuators, and eliminate the need for cg offset and an on-board propulsive reaction control system. Due to its simplicity, drag modulation-based aerocapture is a technique that can be readily tested via a smallsat, with results that can be applied to a variety of different planetary missions. This investigation is focused on the development of a conceptual smallsat mission that will demonstrate the feasibility of an aerocapture system. The spacecraft will be deployed as a secondary payload from a GTO; the conditions provided by the high-energy Earth-return trajectory from GTO will help minimize aerocapture targeting and post-maneuver delta-V requirements. Upon entering the atmosphere, the smallsat will employ hypersonic drag-modulation techniques to control the aerocapture maneuver: after enough energy has been lost via atmospheric drag, the spacecraft will jettison a drag shield, thereby modifying its ballistic coefficient and enabling capture into a relevant orbit after the atmospheric pass. Conceptual mission development has been focused on the design of the smallsat and the aerocapture device, as well as the modeling and selection of the spacecraft\u27s atmospheric trajectory. The difference in ballistic coefficients before and after the drag shield is jettisoned will be the primary source of control authority over the aerocapture maneuver; as such, sizing both the main spacecraft body and the drag shield directly drives numerous other mission design aspects. Aerocapture system design is being accomplished through means of a comprehensive trade study, with focus will placed on control requirements, simplicity, and the applicability of the system to other missions. Three-degree-of-freedom numeric simulation can be used to model the spacecraft\u27s trajectory, enabling quick analysis of peak heating and post-aerocapture orbits. Focus has also been placed on the scalability of results to different types of missions at Mars and other targets. This investigation will culminate in a baseline small satellite mission concept that is fully documented, including technical approach, management plan, cost estimation and risk assessment. A successful mission will show that drag modulation-based aerocapture can be used as an effective means of orbit insertion at Earth and other atmospheric worlds, with scalable applications to both large and small spacecraft