Development of finite element model for approximating composite structures crushing characteristics

A finite element model of a composite keel beam web was developed to approximate its crushing behavior and to provide estimates of its energy absorption characteristics.The model development was done as part of NASA Langley Impact DynamicsResearch Facility development of an energy absorptive composite keel beam for use in general aviation aircraft subfloors intended to increase occupant survivability in crashimpacts with vertical sink rates up to 30 ft/s2. The finite element model was developed by modeling the impact test of a block core keel beam web specimen referred to asB1.4 [8] and establishing correlation with energy absorption parameters measured during testing. SDRC Ideas® Master Series was used to generate the specimen geometry and finite element mesh. MSC/Dytran™ was used to define the test kinematics and dynamics, the specimen material constitutive relationships and failure criteria and to solve the model. Initially the glass fabric/epoxy and kevlar fabric/epoxy material wrapping the foam core was modeled as an orthotropic laminate and assigned a Tsai-Wu failure criteria. The foam was modeled as crushable material with a stress-strain and failure relationship defined by the foam stress-strainThe impact dynamics was modeled as a free fall onto a rigidly supported specimen.This specimen/test representation resulted in premature failure of the laminate and collapse of the specimen model. The premature failure was due to unrealistically high accelerations and impact forces being generated by the free fall impact. The free fall curve.Ill Impact was eliminated and the impact acceleration was defined to be that of the actual acceleration pulse generated in the test. With this representation the model’s crushing behavior similar to that of the test. However the energy absorption parameter magnitudes were low. The model crash initiation load was 2.8 kLb (12.5 kN) and its sustained crash load was .25 kLb (1.1 kN) which were 63.7% and 91.9% below recorded test values respectively. The low magnitudes were a result of the orthotropic constitutive relationship not representing the material nonlinear shear stress-shear strain behavior and the brittle failure mode of the Tsai-Wu failure criteria. Constitutive relationship and failure criteria that better approximates the glass fabric/epoxy and kevlar fabric/epoxy material are needed to improve the model\u27s estimation of energy absorptive parameters. An approximation of these materials behavior was made by representing them as a laminate of isotropic elastic-plastic laminate with strain based failure criteria. With this representation the model crashinitiation load was 6.7 kLb (30.0\u27kN) and its sustained crash load was 3.1 kLb (13.8kN) which are within 5% and 1.5% of recorded test values respectively. Thus the model closely approximates the specimen crashing behavior and its energy absorption parameters for the defined impact acceleration when constitutive/failure relationships similar to the material are used

Trajectory optimization and performance sensitivity studies of NASA Langley\u27s proposed small satellite launch system

Optimal ascent and booster glide back trajectories were determined for NASA Langley\u27s proposed small satellite launcher, SSL-1, for a given polar mission, vehicle configuration, propulsion system, aerodynamic characteristics, structural characteristics and trajectory constraints. The optimal ascent and glide back trajectories were determined for a launch from Vandenberg Air Force Base launch pad SLC-2W and booster glide back to Vandenberg Air Force Base runway 30. The SSL-1 ascent and glide back trajectories were simulated and optimized in POST, Program to Optimize Simulated Trajectories. Inertial pitch angles relative to a inertial launch frame were specified as independent variables in the ascent trajectory and optimized to yield maximum weight to orbit. Aerodynamic angles were specified as independent variables in the booster glide back trajectory and optimized to yield maximum altitude at a heading alignment cylinder six nautical miles south of runway 30. The SSL-1 could not perform an ascent trajectory that satisfies the constraint of gliding the booster back to a heading alignment cylinder for runway 30. The optimal SSL-1 ascent trajectory results in 1022 lb of total weight and 384 lb of payload being inserted into a 150 nautical mile polar orbit. However, a booster glide back that achieves a desired altitude goal of 18800 ft at a heading alignment cylinder for runway 30 could not be performed from the separation point of the optimal ascent for the given aerodynamic and structural limits. The separation Mach number could not be reduced to a point where the booster could attain a desired glide back altitude using reductions in booster size alone since the booster size could not be reduced more than 3% and meet the dynamic pressure at separation constraint of 300 Ib/ft2. The glide back altitude goal can be obtained if the structural normal force limit is increased to 3g loads or the aerodynamic constraint on dynamic pressure at separation is increased to 400 Ib/ft2. The altitude goal will likely be obtained if a high angle of attack drag maneuver is performed between Mach numbers 3.2 and 1.2. The maximum allowable angle of attack for stable flight in this speed range and the corresponding lift/drag characteristics are needed to quantify the obtainable altitude. To achieve the desired altitude goal, modifications in the aerodynamic and/or structural limitations are needed. Weight to orbit performance is influenced by the dynamic pressure at separation constraint but is not sensitive to it. The weight to orbit ranges from 384 lb to 400 lb for dynamic pressure limits of 300 Ib/ft2 to 500 Ib/f2. The glide back altitude is sensitive to the dynamic pressure at separation constraint. Glide back altitude at the HAC ranges from 11995 ft to 24600 ft for dynamic pressure limits of 300 Ib/ft2 to 500 Ib/ft2. Both ascent and glide back performance is insensitive to atmospheric winds. Mean winds reduce payload by 2 lb and increase altitude at the heading alignment cylinder 515 ft. The SSL-1 weight to orbit performance is insensitive to movements in the vehicle\u27s C.G. Movements up to 7% of the reference length result in a 2 lb change in payload. The glide back is sensitive to structural normal force limits Increasing the limit from 2.5g to 5.0g increases altitude at the heading alignment cylinder from 11995 ft to 23410 ft

Trajectory-Based Loads for the Ares I-X Test Flight Vehicle

In trajectory-based loads, the structural engineer treats each point on the trajectory as a load case. Distributed aero, inertial, and propulsion forces are developed for the structural model which are equivalent to the integrated values of the trajectory model. Free-body diagrams are then used to solve for the internal forces, or loads, that keep the applied aero, inertial, and propulsion forces in dynamic equilibrium. There are several advantages to using trajectory-based loads. First, consistency is maintained between the integrated equilibrium equations of the trajectory analysis and the distributed equilibrium equations of the structural analysis. Second, the structural loads equations are tied to the uncertainty model for the trajectory systems analysis model. Atmosphere, aero, propulsion, mass property, and controls uncertainty models all feed into the dispersions that are generated for the trajectory systems analysis model. Changes in any of these input models will affect structural loads response. The trajectory systems model manages these inputs as well as the output from the structural model over thousands of dispersed cases. Large structural models with hundreds of thousands of degrees of freedom would execute too slowly to be an efficient part of several thousand system analyses. Trajectory-based loads provide a means for the structures discipline to be included in the integrated systems analysis. Successful applications of trajectory-based loads methods for the Ares I-X vehicle are covered in this paper. Preliminary design loads were based on 2000 trajectories using Monte Carlo dispersions. Range safety loads were tied to 8423 malfunction turn trajectories. In addition, active control system loads were based on 2000 preflight trajectories using Monte Carlo dispersions

Ares I-X Separation and Reentry Trajectory Analyses

The Ares I-X Flight Test Vehicle was launched on October 28, 2009 and was the first and only test flight of NASA s two-stage Ares I launch vehicle design. The launch was successful and the flight test met all of its primary and secondary objectives. This paper discusses the stage separation and reentry trajectory analysis that was performed in support of the Ares I-X test flight. Pre-flight analyses were conducted to assess the risk of stage recontact during separation, to evaluate the first stage flight dynamics during reentry, and to define the range safety impact ellipses of both stages. The results of these pre-flight analyses were compared with available flight data. On-board video taken during flight showed that the flight test vehicle successfully separated without any recontact. Reconstructed trajectory data also showed that first stage flight dynamics were well characterized by pre-flight Monte Carlo results. In addition, comparisons with flight data indicated that the complex interference aerodynamic models employed in the reentry simulation were effective in capturing the flight dynamics during separation. Finally, the splash-down locations of both stages were well within predicted impact ellipses

Ares I-X Range Safety Simulation Verification and Analysis IV and V

NASA s ARES I-X vehicle launched on a suborbital test flight from the Eastern Range in Florida on October 28, 2009. NASA generated a Range Safety (RS) flight data package to meet the RS trajectory data requirements defined in the Air Force Space Command Manual 91-710. Some products included in the flight data package were a nominal ascent trajectory, ascent flight envelope trajectories, and malfunction turn trajectories. These data are used by the Air Force s 45th Space Wing (45SW) to ensure Eastern Range public safety and to make flight termination decisions on launch day. Due to the criticality of the RS data in regards to public safety and mission success, an independent validation and verification (IV&V) effort was undertaken to accompany the data generation analyses to ensure utmost data quality and correct adherence to requirements. Multiple NASA centers and contractor organizations were assigned specific products to IV&V. The data generation and IV&V work was coordinated through the Launch Constellation Range Safety Panel s Trajectory Working Group, which included members from the prime and IV&V organizations as well as the 45SW. As a result of the IV&V efforts, the RS product package was delivered with confidence that two independent organizations using separate simulation software generated data to meet the range requirements and yielded similar results. This document captures ARES I-X RS product IV&V analysis, including the methodology used to verify inputs, simulation, and output data for an RS product. Additionally a discussion of lessons learned is presented to capture advantages and disadvantages to the IV&V processes used

Use of Flexible Body Coupled Loads in Assessment of Day of Launch Flight Loads

A Day of Launch flight loads assessment technique that determines running loads calculated from flexible body coupled loads was developed for the Ares I-X Flight Test Vehicle. The technique was developed to quantify DOL flight loads in terms of structural load components rather than the typically used q-alpha metric to provide more insight into the DOL loads. In this technique, running loads in the primary structure are determined from the combination of quasi-static aerodynamic loads and dynamic loads. The aerodynamic loads are calculated as a function of time using trajectory parameters passed from the DOL trajectory simulation and are combined with precalculated dynamic loads using a load combination equation. The potential change in aerodynamic load due to wind variability during the countdown is included in the load combination. In the event of a load limit exceedance, the technique allows the identification of what load component is exceeded, a quantification of how much the load limit is exceeded, and where on the vehicle the exceedance occurs. This technique was used to clear the Ares I-X FTV for launch on October 28, 2009. This paper describes the use of coupled loads in the Ares I-X flight loads assessment and summarizes the Ares I-X load assessment results

Upper Atmospheric Monitoring for Ares I-X Ascent Loads and Trajectory Evaluation on the Day-of-Launch

During the launch countdown of the Ares I-X test vehicle, engineers from Langley Research Center will use profiles of atmospheric density and winds in evaluating vehicle ascent loads and controllability. A schedule for the release of balloons to measure atmospheric density and winds has been developed by the Natural Environments Branch at Marshall Space Flight Center to help ensure timely evaluation of the vehicle ascent loads and controllability parameters and support a successful launch of the Ares I-X vehicle

Ares I-X Range Safety Flight Envelope Analysis

Ares I-X was the first test flight of NASA's Constellation Program's Ares I Crew Launch Vehicle designed to provide manned access to low Earth orbit. As a one-time test flight, the Air Force's 45th Space Wing required a series of Range Safety analysis data products to be developed for the specified launch date and mission trajectory prior to granting flight approval on the Eastern Range. The range safety data package is required to ensure that the public, launch area, and launch complex personnel and resources are provided with an acceptable level of safety and that all aspects of prelaunch and launch operations adhere to applicable public laws. The analysis data products, defined in the Air Force Space Command Manual 91-710, Volume 2, consisted of a nominal trajectory, three sigma trajectory envelopes, stage impact footprints, acoustic intensity contours, trajectory turn angles resulting from potential vehicle malfunctions (including flight software failures), characterization of potential debris, and debris impact footprints. These data products were developed under the auspices of the Constellation's Program Launch Constellation Range Safety Panel and its Range Safety Trajectory Working Group with the intent of beginning the framework for the operational vehicle data products and providing programmatic review and oversight. A multi-center NASA team in conjunction with the 45th Space Wing, collaborated within the Trajectory Working Group forum to define the data product development processes, performed the analyses necessary to generate the data products, and performed independent verification and validation of the data products. This paper outlines the Range Safety data requirements and provides an overview of the processes established to develop both the data products and the individual analyses used to develop the data products, and it summarizes the results of the analyses required for the Ares I-X launch

Ares-I-X Vehicle Preliminary Range Safety Malfunction Turn Analysis

Ares-I-X is the designation given to the flight test version of the Ares-I rocket (also known as the Crew Launch Vehicle - CLV) being developed by NASA. As part of the preliminary flight plan approval process for the test vehicle, a range safety malfunction turn analysis was performed to support the launch area risk assessment and vehicle destruct criteria development processes. Several vehicle failure scenarios were identified which could cause the vehicle trajectory to deviate from its normal flight path, and the effects of these failures were evaluated with an Ares-I-X 6 degrees-of-freedom (6-DOF) digital simulation, using the Program to Optimize Simulated Trajectories Version 2 (POST2) simulation framework. The Ares-I-X simulation analysis provides output files containing vehicle state information, which are used by other risk assessment and vehicle debris trajectory simulation tools to determine the risk to personnel and facilities in the vicinity of the launch area at Kennedy Space Center (KSC), and to develop the vehicle destruct criteria used by the flight test range safety officer. The simulation analysis approach used for this study is described, including descriptions of the failure modes which were considered and the underlying assumptions and ground rules of the study, and preliminary results are presented, determined by analysis of the trajectory deviation of the failure cases, compared with the expected vehicle trajectory

A Proposed Ascent Abort Flight Test for the Max Launch Abort System

The NASA Engineering and Safety Center initiated the Max Launch Abort System (MLAS) Project to investigate alternate crew escape system concepts that eliminate the conventional launch escape tower by integrating the escape system into an aerodynamic fairing that fully encapsulates the crew capsule and smoothly integrates with the launch vehicle. This paper proposes an ascent abort flight test for an all-propulsive towerless escape system concept that is actively controlled and sized to accommodate the Orion Crew Module. The goal of the flight test is to demonstrate a high dynamic pressure escape and to characterize jet interaction effects during operation of the attitude control thrusters at transonic and supersonic conditions. The flight-test vehicle is delivered to the required test conditions by a booster configuration selected to meet cost, manufacturability, and operability objectives. Data return is augmented through judicious design of the boost trajectory, which is optimized to obtain data at a range of relevant points, rather than just a single flight condition. Secondary flight objectives are included after the escape to obtain aerodynamic damping data for the crew module and to perform a high-altitude contingency deployment of the drogue parachutes. Both 3- and 6-degree-of-freedom trajectory simulation results are presented that establish concept feasibility, and a Monte Carlo uncertainty assessment is performed to provide confidence that test objectives can be met