Organizational Unit:

Aerospace Systems Design Laboratory (ASDL)

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https://hdl.handle.net/1853/70744

Parent Organization

Organizational Unit

Daniel Guggenheim School of Aerospace Engineering

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https://finding-aids.library.gatech.edu/agents/corporate_entities/1135

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Publication Search Results

Now showing 1 - 10 of 149

Crew Launch Vehicle (CLV) Independent Performance Evaluation

(Georgia Institute of Technology, 2005-11) Young, David Anthony ; Krevor, Zachary C. ; Tanner, Christopher ; Thompson, Robert W. ; Wilhite, Alan W.

The crew launch vehicle is a new NASA launch vehicle design proposed by the Exploration Systems Architecture Study (ESAS) to provide reliable transportations of humans and cargo from the earth’s surface to low earth orbit (LEO). ESAS was charged with the task of looking at the options for returning to the moon in support of the Vision for Space Exploration. The ESAS results, announced in September 2005, favor the use of shuttle-derived launch vehicles for the goals of servicing the International Space Station after the retirement of the STS and supporting the proposed lunar exploration program. The first launch vehicle to be developed is the Crew Launch Vehicle (CLV), which will be operational by 2012, and will be derived from a four-segment Shuttle Solid Rocket Booster (SRB) and an upper-stage powered by an expendable version of the Space Shuttle Main Engine (SSME). The CLV will be capable of sending approximately 60,000 lbs to LEO in the form of a Crew Exploration Vehicle (CEV) as well as a Service Module (SM) to support the CEV. The purpose of this paper is to compare the published CLV numbers with those computed using the design methodology currently used in the Space System Design Laboratory (SSDL) at The Georgia Institute of Technology. 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 CLV. The final performance, reliability, and cost information are then compared with the original ESAS results and the discrepancies are analyzed. Once the design process was completed, a parametric Excel based model is created from the point design. This model can be used to resize CLV for changing system metrics (such as payload) as well as changing technologies.
Preliminary Design of a 2D Supersonic Inlet to Maximize Total Pressure Recovery

(Georgia Institute of Technology, 2005-09) Ran, Hongjun ; Mavris, Dimitri N.

This paper provides a method of preliminary design for a two-dimensional, mixed compression, two-ramp supersonic inlet to maximize total pressure recovery and match the mass flow demand of the engine. For an on-design condition, the total pressure recovery is maximized according to the optimization criterion, and the dimensions of the inlet in terms of ratios to the engine face diameter are calculated. The optimization criterion is defined such that in a system of (n-1) oblique shocks and one normal shock in two dimensions, the maximum shock pressure recovery is obtained when the shocks are of equal strength. This paper also provides a method to estimate the total pressure recovery for an off-design condition for the specified inlet configuration. For an off-design condition, conservative estimation of the total pressure recovery is given so that performance of the engine at the off-design condition can be estimated. To match the mass flow demand of the engine, the second ramp angle is adjusted and the open/close schedule of a bypass door is determined. The effects of boundary layer are not considered for the supersonic part of the inlet, however friction and expansion losses are considered for the subsonic diffuser.
Entry System Options for Human Return from the Moon and Mars

(Georgia Institute of Technology, 2005-08) Putnam, Zachary R. ; Braun, Robert D. ; Rohrschneider, Reuben R. ; Dec, John A.

Earth entry system options for human return missions from the Moon and Mars were analyzed and compared to identify trends among the configurations and trajectory options and to facilitate informed decision making at the exploration architecture level. Entry system options included ballistic, lifting capsule, biconic, and lifting body configurations with direct entry and aerocapture trajectories. For each configuration and trajectory option, the thermal environment, deceleration environment, crossrange and downrange performance, and entry corridor were assessed. In addition, the feasibility of a common vehicle for lunar and Mars return was investigated. The results show that a low lift-to-drag ratio (L/D = 0.3) vehicle provides sufficient performance for both lunar and Mars return missions while providing the following benefits: excellent packaging efficiency, low structural and TPS mass fraction, ease of launch vehicle integration, and system elegance and simplicity. Numerous configuration options exist that achieve this L/D.
LASSO - Lunar Architecture Stochastic Simulator and Optimizer

(Georgia Institute of Technology, 2005-08) Alemany, Kristina ; Street, David C.

The Lunar Architecture Stochastic Simulator and Optimizer (LASSO) is a simulation-based capability, based upon discrete event simulation (DES), for evaluating and optimizing flight element options for lunar transportation architectures. This simulation capability improves the ability to rapidly measure cost, reliability, and schedule impacts of various top-level architecture decisions and individual elements within an architecture. The ability to probabilistically simulate and even optimize an overall transportation approach represents a significant enhancement over current deterministic analysis capabilities for top-level decision making. LASSO integrates a database of flight elements in Microsoft Excel® with architecture models in Rockwell Software’s Arena®. The Arena models are further integrated into Phoenix Integration’s ModelCenter® to allow optimization of the overall architecture by selecting various combinations of elements from the database. Sample results are presented for an expendable and a reusable lunar transportation architecture to illustrate the capabilities of LASSO for top-level decision making.
An Experimental and Analytical Study of High-Energy-Density propellants for Liquid Rocket Engines

(Georgia Institute of Technology, 2005-07) Kokan, Timothy Salim ; Olds, John R.

There exists wide ranging research interest in high-energy-density matter (HEDM) propellants as a potential replacement of existing industry standard fuels (LH2, RP-1, MMH, UDMH) for liquid rocket engines. The U.S. Air Force Research Laboratory, the U.S. Army Research Lab, and the NASA Marshall Space Flight Center each have ongoing programs in the synthesis and development of these potential new propellants. The thermophysical understanding of HEDM propellants is necessary to model their performance in the conceptual design of liquid rocket engines. Most industry standard powerhead design tools (e.g. NPSS, ROCETS, and REDTOP-2) require several thermophysical properties of a given propellant over a wide range of temperature and pressure. These properties include enthalpy, entropy, density, internal energy, specific heat, viscosity, and thermal conductivity. For most of these potential new HEDM propellants, this thermophysical data either does not exist or is incomplete over the range of temperature and pressure necessary for liquid rocket engine design and analysis. The work presented is a technique for obtaining enthalpy and density data for new propellants through the use of a combination of analytical/computational methods (quantum mechanics and molecular dynamics) and experimental investigations. Details of this technique and its application to an example HEDM fuel currently of interest are provided.
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.
Nuclear Gas Turbine Propulsion System for Long Endurance Titan Exploration

(Georgia Institute of Technology, 2005-07) Colby, Luke S. ; Prakash, Ravi ; Braun, Robert D.

An innovative propulsion system concept that enables powered flight on Saturn’s largest moon, Titan, is discussed. This propulsion system concept uses waste heat from a NASA Multi-Mission Radioisotopic Thermoelectric Generator (MMRTG) to power a gas turbine engine. The propulsion system captures MMRTG waste heat by utilizing the in-situ resources of Titan’s cold dense nitrogen atmosphere as a working fluid, passed through a heat exchanger. The heated gas is then run through a turbine to extract electrical power significantly greater than that available from the MMRTG’s thermoelectric effect. In addition to analysis, an experimental system was constructed to validate the feasibility of the proposed concept. This investigation compares the results obtained with this experimental system to analytic predictions. Experimental system performance exceeding 500 watts of measured power output was achieved. This propulsive performance enables consideration of a robotic vertical takeoff and landing vehicle with an altitude ceiling of 15 km, range of 50 km, endurance of 3-4 months, payload capacity of 25 kg, and a gross mass of 400 kg as a future Titan aerial platform.
Artemis: A Reusable Excursion Vehicle Concept for Lunar Exploration

(Georgia Institute of Technology, 2005-07) Young, David Anthony ; Olds, John R. ; Hutchinson, Virgil L., Jr. ; Krevor, Zachary C. ; Young, James J.

Artemis is a reusable excursion vehicle for lunar landing missions. It is intended to transport a notional CEV vehicle from low lunar orbit (LLO) to the lunar surface. It can be reused by refueling the vehicle in LLO. Artemis is nominally sized to carry a 10 MT payload to the lunar surface and then return it to LLO. Artemis is powered by four liquid oxygen and liquid hydrogen fueled RL-10 engines. These RL-10 engines provide the necessary thrust and allow the Artemis lander to complete its nominal mission with two engines inoperative. The Artemis lander has volume margin built into its propellant tanks. This volume margin combined with an innovative cross-feed system allows Artemis to complete its ascent from the lunar surface with a propellant tank failure. This cross-feed system also allows Artemis to adjust the center of gravity (cg) of the vehicle by transferring propellant among the propellant tanks. Artemis lands on the moon with six articulating legs. This provides redundancy against a leg failure on landing and provides Artemis with the ability to land on uneven terrain. This vehicle is designed to be launched by a heavy-lift evolved expendable launch vehicle (EELV). This design constraint results in the distinct shape of the lander. Artemis is launched as a compact cylinder in the EELV payload shroud, and then autonomously assembles itself via robotic arms similar to those currently used by the shuttle program. Details of the conceptual design process used for Artemis are included in this paper. The disciplines used in the design include configuration, propulsion design and selection, trajectory, mass properties, structural design, cost, operations, and 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 team process and used to minimize the gross weight of the Artemis. Once the design process was completed, a parametric Excel based model was created from the point design. This model can be used to resize Artemis for changing system metrics (such as payload) as well as changing technologies. The Artemis recurring and non-recurring costs were also computed. The total development cost including the design, development, testing and evaluation (DDT&E) cost is $2.17 B FY'04. The theoretical first unit (TFU) cost is $303 M FY'04. Trade studies on life cycle costs (LCC) vs. fuel cost to LLO as well as flight rate are also discussed. A summary of design disciplines as well as the economic results are included.
Sonic Boom Minimization Using Improved Linearized Tools and Probabilistic Propagation

(Georgia Institute of Technology, 2005-06-10) Rallabhandi, Sriram Kishore ; Mavris, Dimitri N.

Sonic boom modelling is multidisciplinary involving aerodynamic and aero-acoustics analyses. The near field pressure signature is first obtained using either linearized or non-linear methods. This is then converted into a F-function, which is then propagated to the ground using aero-acoustic routines. Existing linearized methods operate on simple approximations of true geometry. Using improved linearized tools that operate on unstructured water-tight geometries, the accuracy and efficacy of shape optimization can be greatly improved. The sonic boom minimization technique is reformulated as an optimization problem and boom propagation is carried out in a probabilistic fashion. A bi-level reverse optimization is conducted to design aircraft to meet low sonic boom requirements under atmospheric uncertainty.
Probabilistic Matching of Turbofan Engine Performance Models to Test Data

(Georgia Institute of Technology, 2005-06-06) Roth, Bryce Alexander ; Doel, David L. ; Cissell, Jeffrey J.

This paper describes the development of an improved method for reliable, repeatable, and accurate matching of engine performance models to test data. The centerpiece of this approach is a minimum variance estimator algorithm with a priori estimates which addresses both deterministic and probabilistic aspects of the problem. Specific probabilistic aspects include uncertainty in the measurements, prior expectations on model matching parameters, and noise in the power setting parameters. The algorithm is able to produce optimal results using any number of measurements and model matching parameters and can therefore take advantage of all measured data to produce the best possible match. This improves on current matching algorithms which require that the number of measured parameters be equal to the number of model matching parameters. This algorithm has been implemented in the Numerical Propulsion System Simulation (NPSS) and tested on a generic high-bypass turbofan model typical of those used in commercial service. The baseline engine model and simulated test data are described in detail. Several exercises are discussed to illustrate results available from this algorithm including the matching of a typical power calibration data set and matching of a typical production engine data set.