Paul V. Tartabini

Multidisciplinary Analysis of a Lifting Body Launch Vehicle

As part of phase 2 of the X-33 Program, NASA selected an integrated lifting body/aerospike engine configuration as the study vehicle for the conceptual analysis of a single-stage-to-orbit reusable launch vehicle. A team at NASA Langley Research Center participated in the screening and evaluation of a number of proposed vehicle configurations in the early phases of the conceptual design process. The performance analyses that supported these studies were conducted to assess the effect of the vehicles lifting capability, linear aerospike engine, and metallic thermal protection system on the weight and performance of the vehicle. These performance studies were conducted in a multidisciplinary fashion that indirectly linked the trajectory optimization with weight estimation and aerothermal analysis tools. This approach was necessary to develop optimized ascent and entry trajectories that met all vehicle design constraints. Significant improvements in ascent performance were achieved when the vehicle flew a lifting trajectory and varied the engine mixture ratio during flight. Also, a considerable reduction in empty weight was possible by adjusting the total oxidizer-to-fuel and liftoff thrust-to-weight ratios. However, the optimal ascent flight profile had to be altered to ensure that the vehicle could be trimmed in pitch using only the flow diverting capability of the aerospike engine. Likewise, the optimal entry trajectory had to be tailored to meet thermal protection system heating rate and transition constraints while satisfying a crossrange requirement.

Aerodynamic Characteristics and Glide-Back Performance of Langley Glide-Back Booster

NASA-Langley Research Center is conducting system level studies on an-house concept of a small launch vehicle to address NASA’s needs for rapid deployment of small payloads to Low Earth Orbit. The vehicle concept is a three-stage system with a reusable first stage and expendable upper stages. The reusable first stage booster, which glides back to launch site after staging around Mach 3 is named the Langley Glide-Back Booster (LGBB). This paper discusses the aerodynamic characteristics of the LGBB from subsonic to supersonic speeds, development of the aerodynamic database and application of this database to evaluate the glide back performance of the LGBB. The aerodynamic database was assembled using a combination of wind tunnel test data and engineering level analysis. The glide back performance of the LGBB was evaluated using a trajectory optimization code and subject to constraints on angle of attack, dynamic pressure and normal acceleration.

Ares I-X Trajectory Reconstruction: Methodology and Results

The Ares I-X trajectory reconstruction produced best-estimated trajectories of the flight-test vehicle ascent through stage separation and of the first- and upper-stage entries after separation. The trajectory-reconstruction process combines onboard, ground-based, and atmospheric measurements to produce the trajectory estimates, using an iterated extended Kalman filter algorithm. The Ares I-X vehicle had a number of onboard and ground-based sensors that were available, including inertial measurement units, radar, air data, and weather balloons. However, due to problems with calibrations and/or data, not all of the sensor data were used. This paper describes the methodology and results of the trajectory-reconstruction process, including flight-data preprocessing and input uncertainties, trajectory-estimation algorithms and dynamic models, output transformations, and comparisons with preflight predictions. The results of the reconstruction indicate nominal vehicle performance that is well within the range o...

A Multidisciplinary Performance Analysis of a Lifting-Body Single-Stage-to-Orbit Vehicle

Lockheed Martin Skunk Works (LMSW) is currently developing a single-stage-to-orbit reusable launch vehi- cle called VentureStar™. A team at NASA Langley Re- search Center participated with LMSW in the screening and evaluation of a number of early VentureStar™ con- figurations. The performance analyses that supported these initial studies were conducted to assess the effect of a lifting body shape, linear aerospike engine and me- tallic thermal protection system (TPS) on the weight and performance of the vehicle. These performance studies were performed in a multidisciplinary fashion that indi- rectly linked the trajectory optimization with weight es- timation and aerothermal analysis tools. This approach was necessary to develop optimized ascent and entry tra- jectories that met all vehicle design constraints. Significant improvements in ascent performance were achieved when the vehicle flew a lifting trajectory and varied the engine mixture ratio during flight. Also, a considerable reduction in empty weight was possible by adjusting the total oxidizer-to-fuel and liftoff thrust-to- weight ratios. However, the optimal ascent flight profile had to be altered to ensure that the vehicle could be trimmed in pitch using only the flow diverting capability of the aerospike engine. Likewise, the optimal entry tra- jectory had to be tailored to meet TPS heating rate and transition constraints while satisfying a crossrange re- quirement.

ASCENT, STAGE SEPARATION AND GLIDEBACK PERFORMANCE OF A PARTIALLY REUSABLE SMALL LAUNCH VEHICLE

An integrated analysis is presented for ascent, stage separation and glide back performance of a small, partially reusable launch vehicle sized for a payload of about 330 lbs to a 150 nm polar orbit. The altitude margin was used a performance metric for the glideback performance. Aerodynamic databases for each of these three phases of flight were developed using a combination of engineering level code, free stream and proximity wind tunnel test data and Euler CFD results. The ascent and glideback trajectories were generated using POST and the stage separation simulation was done using the in-house software SepSim as a front end to the commercially available multi-body dynamic simulation code ADAMS ® . The payload to the designated polar orbit was optimized subject to the constraints imposed by stage separation and adequate performance reserve for the glideback booster in addition to the usual ascent trajectory constraints. Nomenclature A N normal acceleration, g’s ! angle of attack, deg ∀! relative difference in angle of attack, deg

Development and evaluation of an operational aerobraking strategy for Mars Odyssey

The Mars 2001 Odyssey Orbiter successfully completed the aerobraking phase of its mission on 11 January 2002. The support provided by NASA’s Langley Research Center to the navigation team at the Jet Propulsion Laboratory, California Institute of Technology, in the planning and operational support of Mars Odyssey aerobraking is discussed. Specifically, the development of a three-degree-of-freedom aerobraking trajectory simulation and its application to both preflight planning activities and operations is described. The importance of running the simulation in a Monte Carlo fashion to capture the effects of mission and atmospheric uncertainties is demonstrated, and the utility of including predictive logic within the simulation that could mimic operational maneuver decision making is shown. A description is also provided of how the simulation was adapted to support flight operations as both a validation and risk reduction tool and as a means of obtaining a statistical basis for maneuver strategy decisions. This latter application was the first use of Monte Carlo trajectory analysis in an aerobraking mission.

Constraint Force Equation Methodology for Modeling Multi-Body Stage Separation Dynamics

This paper discusses a generalized approach to the multi-body separation problems in a launch vehicle staging environment based on constraint force methodology and its implementation into the Program to Optimize Simulated Trajectories II (POST2), a widely used trajectory design and optimization tool. This development facilitates the inclusion of stage separation analysis into POST2 for seamless end-to-end simulations of launch vehicle trajectories, thus simplifying the overall implementation and providing a range of modeling and optimization capabilities that are standard features in POST2. Analysis and results are presented for two test cases that validate the constraint force equation methodology in a stand-alone mode and its implementation in POST2.

Modeling Multibody Stage Separation Dynamics Using Constraint Force Equation Methodology

This paper discusses the application of the constraint force equation methodology and its implementation for multibody separation problems using three specially designed test cases. The first test case involves two rigid bodies connected by a fixed joint, the second case involves two rigid bodies connected with a universal joint, and the third test case is that of Mach 7 separation of the X-43A vehicle. For the first two cases, the solutions obtained using the constraint force equation method compare well with those obtained using industry- standard benchmark codes. For the X-43A case, the constraint force equation solutions show reasonable agreement with the flight-test data. Use of the constraint force equation method facilitates the analysis of stage separation in end-to-end simulations of launch vehicle trajectories

Mach 10 Stage Separation Analysis for the X43-A

This paper describes the pre-flight stage separation analysis that was conducted in sup- port of the final flight of the X-43A. In that flight, which occurred less than eight months af- ter the successful Mach 7 flight, the X-43A Research Vehicle attained a peak speed of Mach 9.6. Details are provided on how the lessons learned from the Mach 7 flight affected separa- tion modeling and how adjustments were made to account for the increased flight Mach number. Also, the procedure for defining the feedback loop closure and feed-forward pa- rameters employed in the separation control logic are described, and their effect on separa- tion performance is explained. In addition, the range and nominal values of these parame- ters, which were included in the Mission Data Load, are presented. Once updates were made, the nominal pre-flight trajectory and Monte Carlo statistical results were determined and stress tests were performed to ensure system robustness. During flight the vehicle per- formed within the uncertainty bounds predicted in the pre-flight analysis and ultimately set the world record for airbreathing powered flight.

Simulation and Analyses of Multi-Body Separation in Launch Vehicle Staging Environment

The development of methodologies, techniques, and tools for analysis and simulation of multi-body separation is critically needed for successful design and operation of next generation launch vehicles. As a part of this activity, ConSep simulation tool is being developed. ConSep is a generic MATLAB-based front-and-back-end to the commercially available ADAMS  solver, an industry standard package for solving multi-body dynamic problems. This paper discusses the 3-body separation capability in ConSep and its application to the separation of the Shuttle Solid Rocket Boosters (SRBs) from the External Tank (ET) and the Orbiter. The results are compared with STS-1 flight data.