Organizational Unit:

Space Systems Design Laboratory (SSDL)

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

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Organizational Unit

Daniel Guggenheim School of Aerospace Engineering

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https://finding-aids.library.gatech.edu/agents/corporate_entities/1138

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Publication Search Results

Now showing 1 - 10 of 47

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.
Performance Characterization of Supersonic Retropropulsion for Application to High-Mass Mars Entry, Descent, and Landing

(Georgia Institute of Technology, 2009-08) Korzun, Ashley M. ; Braun, Robert D.

Prior high-mass Mars EDL systems studies have neglected aerodynamic-propulsive interactions and performance impacts during the supersonic phase of descent. The goal of this investigation is to accurately evaluate the performance of supersonic retropropulsion with increasing vehicle ballistic coefficient across a range of initiation conditions relevant for future high-mass Mars landed systems. Past experimental work has established supersonic retropropulsion trends in static aerodynamics as a function of retropropulsion configuration, freestream conditions, and thrust. From this experimental database, an aerodynamic-propulsive interactions model is created. EDL system performance results are developed with the potential aerodynamic drag preservation included and excluded during this phase of flight for comparison against prior studies. The results of this investigation demonstrate the significance of aerodynamic drag preservation as a function of retropropulsion initiation conditions, characterize mass optimal trajectories utilizing supersonic retropropulsion, and compare propulsion system requirements with existing propulsion systems and systems under development for future exploration missions.
Experimental Determination of Material Properties for Inflatable Aeroshell Structures

(Georgia Institute of Technology, 2009-05) Hutchings, Allison L.. ; Braun, Robert D. ; Masuyama, Kento ; Welch, Joseph V.

As part of a deployable aeroshell development effort, system design, materials evaluation, and analysis methods are under investigation. One specific objective is to validate finite element analysis techniques used to predict the deformation and stress fields of aeroshell inflatable structures under aerodynamic loads. In this paper, we discuss the results of an experimental mechanics study conducted to ensure that the material inputs to the finite element models accurately predict the load elongation characteristics of the coated woven fabric materials used in deployable aeroshells. These coated woven fabrics exhibit some unique behaviors under load that make the establishment of a common set of test protocols difficult. The stiffness of a woven fabric material will be influenced by its biaxial load state. Uniaxial strip tensile testing although quick and informative, may not accurately capture the needed structural model inputs. Woven fabrics, when loaded in the bias direction relative to the warp and fill axes, have a resultant stiffness that is quite low as compared with the warp and fill directional stiffness. We evaluate the experimental results from two load versus elongation test devices. Test method recommendations are made based on the relevance and accuracy of these devices. Experimental work is conducted on a sample set of materials, consisting of four fabrics of varying stiffness and strength. The building blocks of a mechanical property database for future aeroshell design efforts are constructed.
The State of Problem Decomposition in Engineering Design

(Georgia Institute of Technology, 2009-05) Otero, Richard E. ; Braun, Robert D.

As engineers examine larger coupled systems, computational complexity, available resources and the lack of expert intuition create a need for advancing the state of automated decomposition methods. Larger coupled problems may, for instance, expand beyond an expert's experience in manual decomposition. This paper describes the current state of automated decomposition within engineering and explores several promising ideas that could be utilized to advance the state of the practice. It is proposed that the explicit modeling of dependencies is both possible and available for a new avenue of advancement in how problem decomposition is applied to engineering.
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.
Ablative Thermal Response Analysis Using the Finite Element Method

(Georgia Institute of Technology, 2009-01) Dec, John A. ; Braun, Robert D.

A review of the classic techniques used to solve ablative thermal response problems is presented. The advantages and disadvantages of both the finite element and finite difference methods are described. As a first step in developing a three dimensional finite element based ablative thermal response capability, a one dimensional computer tool has been developed. The finite element method is used to discretize the governing differential equations and Galerkin's method of weighted residuals is used to derive the element equations. A code to code comparison between the current 1-D finite element tool and the 1-D Fully Implicit Ablation and Thermal response program (FIAT), a NASA-standard finite difference tool, has been performed.
Trajectory Reconstruction of a Martian Planetary Probe Mission: Reconstruction of the Spirit Mars Exploration Rover Entry Descent Landing Performance

(Georgia Institute of Technology, 2008-10) Wells, Grant W. ; Braun, Robert D.

Accurate post-flight reconstruction of a vehicle’s trajectory during entry into a planetary atmosphere can produce a wide array of valuable information. Data collected through the reconstruction of entry, descent, and landing system performance enables the quantification of performance margins for future systems. Beyond the engineering knowledge gained through trajectory reconstruction, the results may also be used by planetary scientists to generate an accurate atmospheric profile. This paper provides a reconstruction of the trajectory, vehicle orientation, and atmospheric density profile for the hypersonic and supersonic phases of the Spirit Mars Exploration Rover spacecraft.
Design Space Pruning Heuristics and Global Optimization Method for Conceptual Design of Low-Thrust Asteroid Tour Missions

(Georgia Institute of Technology, 2008-09) Alemany, Kristina ; Braun, Robert D.

Electric propulsion has recently become a viable technology for spacecraft, enabling shorter flight times, fewer required planetary gravity assists, larger payload masses, and/or smaller launch vehicles. With the maturation of this technology, however, comes a new set of challenges in the area of trajectory design. In 2006, the 2nd Global Trajectory Optimization Competition (GTOC2) posed a difficult mission design problem: to design the best possible low-thrust trajectory, in terms of final mass and total mission time, that would rendezvous with one asteroid in each of four pre-defined groups. Even with recent advances in low-thrust trajectory optimization, a full enumeration of this problem was not possible. This work presents a two-step methodology for determining the optimum solution to a low-thrust, combinatorial asteroid rendezvous problem. First is a pruning step that uses a heuristic sequence to quickly reduce the size of the design space. Second, a multi-level genetic algorithm is combined with a low-thrust trajectory optimization method to locate the best solutions of the reduced design space. The proposed methodology is then validated by applying it to a problem with a known solution.
Multiobjective Hypersonic Entry Aeroshell Shape Optimization

(Georgia Institute of Technology, 2008-09) Theisinger, John E. ; Braun, Robert D.

A capability has been developed that utilizes multiobjective optimization to identify hypersonic entry aeroshell shapes that will increase landed mass capability. Aeroshell shapes are parameterized using non-uniform rational B-splines to generate complete aeroshell surfaces. Hypersonic aerodynamic objectives and constraints are computed by numerically integrating pressure coefficient distributions obtained using Newtonian flow theory. An integrated optimization environment is created using iSIGHT with single- and multiobjective evolutionary algorithms. Results are presented based on optimization using constraints derived from the aeroshell for the Mars Science Laboratory mission. Resulting solutions clearly demonstrate the trade-offs between drag-area, static stability, and volumetric efficiency for this particular mission.