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Space Systems Design Laboratory (SSDL)

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https://hdl.handle.net/1853/70747

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Daniel Guggenheim School of Aerospace Engineering

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Publication Search Results

Now showing 1 - 10 of 10

High Mass Mars Entry, Descent, and Landing Architecture Assessment

(Georgia Institute of Technology, 2009-09) Steinfeldt, Bradley A. ; Theisinger, John E. ; Korzun, Ashley M. ; Clark, Ian G. ; Grant, Michael J. ; Braun, Robert D.

As the nation sets its sight on returning humans to the Moon and going onward to Mars, landing high mass payloads (>/= 2 t) on the Mars surface becomes a critical technological need. Viking heritage technologies (e.g., 70degrees sphere-cone aeroshell, SLA-561V thermal protection system, and supersonic disk-gap-band parachutes) that have been the mainstay of the United States' robotic Mars exploration program do not provide sufficient capability to land such large payload masses. In this investigation, a parametric study of the Mars entry, descent, and landing design space has been conducted. Entry velocity, entry vehicle configuration, entry vehicle mass, and the approach to supersonic deceleration were varied. Particular focus is given to the entry vehicle shape and the supersonic deceleration technology trades. Slender bodied vehicles with a lift-to-drag ratio (L=D) of 0.68 are examined alongside blunt bodies with L=D = 0.30. Results demonstrated that while the increased L=D of a slender entry configuration allows for more favorable terminal descent staging conditions, the greater structural efficiencies of blunt body systems along with the reduced acreage required for the thermal protection system affords an inherently lighter vehicle. The supersonic deceleration technology trade focuses on inflatable aerodynamic decelerators (IAD) and supersonic retropropulsion, as supersonic parachute systems are shown to be excessively large for further consideration. While entry masses (the total mass at the top of the Mars atmosphere) between 20 and 100 t are considered, a maximum payload capability of 37.3 t results. Of particular note, as entry mass increases, the gain in payload mass diminishes. It is shown that blunt body vehicles provide sufficient vertical L=D to decelerate all entry masses considered through the Mars atmosphere with adequate staging conditions for the propulsive terminal descent. A payload mass fraction penalty of approximately 0.3 exists for the use of slender bodied vehicles. Another observation of this investigation is that the increased aerothermal and aerodynamic loads induced from a direct entry trajectory (velocity ~6.75 km/s) reduce the payload mass fraction by approximately 15% compared to entry from orbital velocity (~4 km/s). It should be noted that while both IADs and supersonic retropropulsion were evaluated for each of the entry masses, configurations, and velocities, the IAD proved to be more mass-efficient in all instances. The sensitivity of these results to modeling assumptions was also examined. The payload mass of slender body vehicles was observed to be approximately four times more sensitive to modeling assumptions and uncertainty than blunt bodies.
Aerodynamic and Aeroelastic Characteristics of a Tension Cone Inflatable Aerodynamic Decelerator

(Georgia Institute of Technology, 2009-05) Clark, Ian G. ; Cruz, Juan R. ; Hughes, Monica F. ; Ware, Joanne S. ; Madlangbayan, Albert ; Braun, Robert D.

The supersonic aerodynamic and aeroelastic characteristics of a tension cone inflatable aerodynamic decelerator were investigated by wind tunnel testing. Two sets of tests were conducted: one using rigid models and another using textile models. Tests using rigid models were conducted over a Mach number range from 1.65 to 4.5 at angles of attack from -12 to 20 degrees. The axial, normal, and pitching moment coefficients were found to be insensitive to Mach number over the tested range. The axial force coefficient was nearly constant (C^A = 1.45 +/- 0.05) with respect to angle of attack. Both the normal and pitching moment coefficients were nearly linear with respect to angle of attack. The pitching moment coefficient showed the model to be statically stable about the reference point. Schlieren images and video showed a detached bow shock with no evidence of large regions of separated flow and/or embedded shocks at all Mach numbers investigated. Qualitatively similar static aerodynamic coefficient and flow visualization results were obtained using textile models at a Mach number of 2.5. Using inflatable textile models the torus pressure required to maintain the model in the fully-inflated configuration was determined. This pressure was found to be sensitive to details in the structural configuration of the inflatable models. Additional tests included surface pressure measurements on rigid models and deployment and inflation tests with inflatable models.
Computational Aerodynamic Analysis of a Tension Cone Supersonic Inflatable Aerodynamic Decelerator

(Georgia Institute of Technology, 2009-03) Clark, Ian G. ; Braun, Robert D.

The 2009 Mars Science Laboratory mission has brought renewed awareness to the difficulty of landing large payloads on the surface of Mars. As a result, a new suite of decelerator technologies is being investigated for future robotic and human-precursor missions. One such technology is the supersonic inflatable aerodynamic decelerator (IAD). Previous studies have shown that a supersonic IAD can provide sizable increases in landed mass versus traditional parachute based systems, particularly for near-term robotic mission. This is due to the ability of an IAD to deploy at higher Mach numbers and dynamic pressures than a parachute, thus allowing for greater deceleration earlier in the entry sequence. 1 2 As part of the Program to Advance Inflatable Decelerators for Atmospheric Entry, one particular configuration, the tension cone, has undergone a series of wind tunnel experiments designed to acquire a full characterization of the aerodynamic performance of a particular tension cone geometry. One test objective entailed the acquisition of a data set useful for validating computational tools for later IAD analysis efforts. This paper presents a summary of the work performed in investigating two separate computational fluid dynamics codes for their suitability in predicting tension cone performance. The first code, NASCART-GT, is a solution adaptive, Cartesian grid code that is used for rapid inviscid analysis of axisymmetric geometries. The second code, Overflow was used for Navier-Stokes analysis of three-dimensional geometries. These codes were evaluated for their ability to match measured pressure distributions, static force and moment coefficients, and observed flowfield characteristics. Overflow is also used to investigate flow features that were not observed during testing, such as the aft body recirculation region. Additional investigation into the aerodynamic performance of a tension cone was performed through a parametric analysis of multiple tension cone geometries. Three primary shape parameters were varied with the goal of identifying undesirable flowfield characteristics such as shocks attached to the surface of the tension shell and to provide insight into the sensitivity of drag to tension cone geometry.
Supersonic Inflatable Aerodynamics Decelerators for Use on Future Robotic Missions To Mars

(Georgia Institute of Technology, 2008-03) Clark, Ian G. ; Hutchings, Allison L. ; Tanner, Christopher L. ; Braun, Robert D.

The 2009 Mars Science Laboratory mission is being designed to place an 850 kg rover on the surface of Mars at an altitude of at least one kilometer [1]. This is being accomplished using the largest aeroshell and supersonic parachute ever flown on a Mars mission. Future missions seeking to place more massive payloads on the surface will be constrained by aeroshell size and deployment limitations of supersonic parachutes [2],[3]. Inflatable aerodynamic decelerators (IADs) represent a technology path that can relax those constraints and provide a sizeable increase in landed mass. This mass increase results from improved aerodynamic characteristics that allow IADs to be deployed at higher Mach numbers and dynamic pressures than can be achieved by current supersonic parachute technology. During the late 1960’s and early 1970’s preliminary development work on IADs was performed. This included initial theoretical shape and structural analysis for a variety of configurations as well as wind tunnel and atmospheric flight tests for a particular configuration, the Attached Inflatable Decelerator (AID). More recently, the Program to Advance Inflatable Decelerators for Atmospheric Entry (PAI-DAE) has been working to mature a second configuration, the supersonic tension cone decelerator, for use during atmospheric entry. 1,2 This paper presents an analysis of the potential advantages of using a supersonic IAD on a proposed 2016 Mars mission. Conclusions drawn are applicable to both the Astrobiology Field Laboratory and Mars Sample Return mission concepts. Two IAD configurations, the AID and tension cone, are sized and traded against their system-level performance impact. Analysis includes preliminary aerodynamic drag estimates for the different configurations, trajectory advantages provided by the IADs, and preliminary geometric and mass estimates for the IAD subsystems. Entry systems utilizing IADs are compared against a traditional parachute system as well as a system employing an IAD in the supersonic regime and a parachute in the subsonic regime. Key sensitivities in IAD design are included to highlight areas of importance in future technology development programs.
Lazarus: A SSTO Hypersonic Vehicle Concept Utilizing RBCC and HEDM Propulsion Technologies

(Georgia Institute of Technology, 2006-11) Young, David Anthony ; Kokan, Timothy Salim ; Clark, Ian G. ; Tanner, Christopher ; Wilhite, Alan W.

Lazarus is an unmanned single stage reusable launch vehicle concept utilizing advanced propulsion concepts such as rocket based combined cycle engine (RBCC) and high energy density material (HEDM) propellants. These advanced propulsion elements make the Lazarus launch vehicle both feasible and viable in today's highly competitive market. The Lazarus concept is powered by six rocket based combined cycle engines. These engines are designed to operate with HEDM fuel and liquid oxygen (LOX). During atmospheric flight the LOX is augmented by air traveling through the engines and the resulting propellant mass fractions make single stage to orbit (SSTO) possible. A typical hindrance to SSTO vehicles are the large wings and landing gear necessary for takeoff of a fully fueled vehicle. The Lazarus concept addresses this problem by using a sled to take off horizontally. This sled accelerates the vehicle to over 500 mph using the launch vehicle engines and a propellant cross feed system. This propellant feed system allows the vehicle to accelerate using its own propulsion system without carrying the necessary fuel required while it is attached to the sled. Lazarus is designed to deliver 5,000 lbs of payload to a 100 nmi x 100 nmi x 28.5° orbit due East out of Kennedy Space Center (KSC). This mission design allows for rapid redeployment of small orbital assets with little launch preparation. Lazarus is also designed for a secondary strike mission. The high speed and long range inherent in a SSTO launch vehicle make it an ideal global strike platform. Details of the conceptual design process used for Lazarus are included in this paper. The disciplines used in the design include aerodynamics, configuration, propulsion design, trajectory, mass properties, cost, operations, reliability and safety. Each of these disciplines was computed using a conceptual design tool similar to that used in industry. These disciplines were then combined into an integrated design process and used to minimize the gross weight of the Lazarus design.
Mars Entry, Descent, and Landing Parametric Sizing and Design Space Visualization Trades

(Georgia Institute of Technology, 2006-08) Alemany, Kristina ; Wells, Grant William ; Theisinger, John ; Clark, Ian G. ; Braun, Robert D.

Entry, descent, and landing (EDL) is a multidimensional, complex problem, which is difficult to visualize in simple plots. The purpose of this work is to develop a systematic visualization scheme that could capture Mars EDL trades as a function of a limited number of variables, such that programmatic design decisions could be effectively made with insight of the design space. Using the Mars Science Laboratory (MSL) as a basis, contour plots have been generated for key EDL figures of merit, such as maximum landed elevation and landed mass as a function of four input parameters: entry mass, entry velocity, entry flight path angle, and vehicle L/D. Additionally, sensitivity plots have been generated in an attempt to capture the effects of varying the fixed input parameters. This set of EDL visualization data has been compiled into a Mars EDL handbook to aid in pre-phase A design space exploration and decision making.
An Evaluation of Ballute Entry Systems for Lunar Return Missions

(Georgia Institute of Technology, 2006-08) Clark, Ian G. ; Braun, Robert D. ; Theisinger, John ; Wells, Grant William

This study investigates the advantages and feasibility of using ballutes for Earth entry at lunar return velocities. Using analysis methods suitable for conceptual design and assuming a CEV type entry vehicle, multiple entry strategies were investigated. Entries that jettison the ballute after achieving low Earth orbit conditions were shown to reduce heating rates to within reusable thermal protection system limits. Deceleration was mitigated to approximately four g's when a moderate amount of lift was applied subsequent to ballute jettison. Primary ballute size drivers are the thermal limitations and areal densities of the ballute material. Performance requirements for both of those metrics were generated over a range of total ballute system masses. Lastly, preliminary investigation of a lower mass cargo variant of the CEV allowed for additional reduction of ballute system mass. However, ballute system mass as a percentage of the total entry mass was shown to be relatively independent of the entry mass.
An Evaluation of Ballute Entry Systems for Lunar Return Missions

(Georgia Institute of Technology, 2006-05-07) Clark, Ian G.

A study was undertaken to assess the advantages and feasibility of using ballutes for Earth entry at lunar return velocities. Using analysis methods suitable for conceptual design, multiple entry strategies were investigated. Entries that jettison the ballute after achieving orbit were shown to reduce heating rates to within reusable thermal protection system limits and deceleration was mitigated to approximately four g’s when a moderate amount of lift was applied post-jettison. Ballute size drivers were demonstrated to be the thermal limitations and areal densities of the ballute material. Performance requirements for both of those metrics were generated over a range of total ballute system masses.
Tempest: Crew Exploration Vehicle Concept

(Georgia Institute of Technology, 2005-07) Hutchinson, Virgil L., Jr. ; Olds, John R. ; Alemany, Kristina ; Christian, John A., III ; Clark, Ian G. ; Crowley, John ; Krevor, Zachary C. ; Rohrschneider, Reuben R. ; Thompson, Robert W. ; Young, David Anthony ; Young, James J.

Tempest is a reusable crew exploration vehicle (CEV) for transferring crew from the Earth to the lunar surface and back. Tempest serves as a crew transfer module that supports a 4-person crew for a mission duration of 18 days, which consists of 8 days total transit duration and 10-day surface duration. Primary electrical power generation and on-orbit maneuvering for Tempest is provided by an attached Power and Propulsion Module (PPM). Hydrogen (H2)/oxygen (O2) fuel cells and a high energy-density matter (HEDM)/liquid oxygen (LOX) propellant reaction control system (RCS) provide power and reaction control respectively during Tempest’s separation from the PPM. Tempest is designed for a lifting entry and is equipped with parachutes for a soft landing. Tempest is part of an overall lunar transportation architecture. The 60,731 lbs combination of Tempest and the PPM are launched atop the notional Centurion C-1 heavylift launch vehicle (HLLV) and delivered to a 162 nmi, 28.5º circular orbit. After separating from the C-1 upper stage, the Tempest/PPM autonomously rendezvous with Manticore, an expendable trans-lunar injection (TLI) stage pre-positioned in the current orbit, and transfer to a lunar trajectory. After entering a 54 nmi polar circular lunar orbit, the Tempest/PPM separate from Manticore. Tempest separates from the PPM and is ferried to/from the lunar surface by Artemis, a reusable lunar lander. Upon return from the lunar surface, Tempest reconnects with the PPM, and the PPM provides the trans-earth injection (TEI) burn required to return to low earth orbit (LEO). Prior to atmospheric entry, Tempest separates from the PPM and subsequently executes a lifting entry trajectory. Crushable thermal foam attached to the lower surface of Tempest serves as an ablative thermal protection system (TPS) and the impact absorber of the parachute landing. Details of the conceptual design process used for Tempest are included in this paper. The disciplines used in the design include: configuration, aerodynamics, propulsion, trajectory, mass properties, environmental control life support system (ECLSS), entry aeroheating and TPS, terminal landing system (TLS), cost, operations, and reliability & safety. Each of these disciplines was computed using a conceptual design tool similar to that used in industry. These disciplines were then combined and optimized for the minimum gross weight of the Tempest CEV. The total development cost including the design, development, testing and evaluation (DDT&E) cost was determined to be $2.9 B FY’04. The theoretical first unit (TFU) cost for the Tempest CEV was $479 M FY’04. A summary of design disciplines as well as the economic results are included.
Reusable Exploration Vehicle (REV): Orbital Space Tourism Concept

(Georgia Institute of Technology, 2005-05) Clark, Ian G. ; Francis, Scott R. ; Otero, Richard E. ; Wells, Grant William

On the heels of the recent success of the X-Prize, sub-orbital space tourism is nearly a reality. Though the requirements are significantly tougher, orbital space tourism is the next logical step. The Reusable Exploration Vehicle (REV) concept is an economically feasible design capable of making this next step. Centered around a lenticular lifting body, the REV concept relies on commercial launch vehicles to reduce DDT&E expenditures. Capable of ferrying five passengers and one crew member for three orbits, the REV is shown to be capable of keeping maximum debt exposure to less than $250M while attaining an IRR of 70% with an estimated market capture of 66%.