Edwin L. Fasanella

Water Impact Test and Simulation of a Composite Energy Absorbing Fuselage Section

In March 2002, a 25-ft/s vertical drop test of a composite fuselage section was conducted onto water. The purpose of the test was to obtain experimental data characterizing the structural response of the fuselage section during water impact for comparison with two previous drop tests that were performed onto a rigid surface and soft soil. For the drop test, the fuselage section was configured with ten 100-lb. lead masses, five per side, that were attached to seat rails mounted to the floor. The fuselage section was raised to a height of 10-ft. and dropped vertically into a 15-ft. diameter pool filled to a depth of 3.5-ft. with water. Approximately 70 channels of data were collected during the drop test at a 10-kHz sampling rate. The test data were used to validate crash simulations of the water impact that were developed using the nonlinear, explicit transient dynamic codes, MSC.Dytran and LS-DYNA. The fuselage structure was modeled using shell and solid elements with a Lagrangian mesh, and the water was modeled with both Eulerian and Lagrangian techniques. The fluid-structure interactions were executed using the fast general coupling in MSC.Dytran and the Arbitrary Lagrange-Euler (ALE) coupling in LS-DYNA. Additionally, the smooth particle hydrodynamics (SPH) meshless Lagrangian technique was used in LS-DYNA to represent the fluid. The simulation results were correlated with the test data to validate the modeling approach. Additional simulation studies were performed to determine how changes in mesh density, mesh uniformity, fluid viscosity, and failure strain influence the test-analysis correlation.

Development of a Scale Model Composite Fuselage Concept for Improved Crashworthiness

A composite fuselage concept for light aircraft has been developed to provide improved crashworthiness. The fuselage consists of a relatively rigid upper section, or passenger cabin, including a stiff structural e oor and a frangible lower section that encloses the crash energy management structure. The crashworthy performance of the fuselage concept was evaluated through impact testing of a one-e fth-scale model fuselage section. The impact design requirement for the scale model fuselage is to achieve a 125- g average e oor-level acceleration for a 31ft/s vertical impact onto a rigid surface. The energy absorption behavior of two different sube oor cone gurations was determined through quasi-static crushing tests. For the dynamic evaluation, each sube oor cone guration was incorporated into a one-e fth-scale model fuselage section, which was dropped from a height of 15 ft to achieve a 31-ft/s vertical velocity at impact. The experimental data demonstrate that the fuselage section with a foam-block sube oor cone guration satise ed the impact design requirement. A second drop test was performed to evaluate the energy absorption performanceofthefuselageconceptfor an off-axis impactcondition. The experimental data are correlated with analytical predictions from a e nite element model developed using the nonlinear, explicit transient dynamic code MSC/DYTRAN.

Full-Scale Crash Test and Simulation of a Composite Helicopter

Abstract A finite element model of the Sikorsky Advanced Composite Airframe Program (ACAP) helicopter was developed using the non-linear, explicit transient dynamic code, MSC.Dytran. Analytical predictions were correlated with experimental data obtained from a full-scale crash test of the Sikorsky ACAP helicopter flight test article that was conducted in June 1999 at the Impact Dynamics Research Facility of NASA Langley Research Centre, Hampton, Virginia, USA. The helicopter was impacted at 11.58 m/s vertical and 9.9-m/s forward velocity with an attitude of 6.25° pitch (nose up) and 3.5° left roll. Due to the relatively long crash pulse duration, a rigid-body helicopter model with an energy absorbing landing gear model was executed initially. Prior to fuselage contact, a deformable structural model was executed with the rigid-body nodal displacements and velocities used as initial conditions.

Impact Testing and Simulation of a Crashworthy Composite Fuselage Section with Energy-Absorbing Seats and Dummies

AbstractA 25-ft/s vertical drop test of a composite fuselage section was conducted with two energy-absorbing seats occupied by an-thropomorphic dummies to evaluate the crashworthy features of the fuselage section and to determine its interaction with theseats and dummies. The 5-ft. diameter fuselage section consists of a stiff structural floor and an energy-absorbing subfloorconstructed of Rohacel foam blocks. The experimental data from this test were analyzed and correlated with predictions froma crash simulation developed using the nonlinear, explicit transient dynamic computer code, MSC.Dytran. The anthropo-morphic dummies were simulated using the Articulated Total Body (ATB) code, which is integrated into MSC.Dytran.IntroductionA research program was conducted at NASA Langley Re-search Center to develop an innovative and cost-effectivecrashworthy fuselage concept for light aircraft and rotor-craft [1-3]. The composite fuselage concept was designedto meet structural and flight-load requirements and toprovide improved crash protection. The two primary de-sign goals for crashworthiness are to limit the impactforces transmitted to the occupants, and to maintain thestructural integrity of the fuselage to ensure a minimumsafe occupant volume. To meet these objectives, an air-craft or rotorcraft fuselage must be designed for high stiff-ness and strength to prevent structural collapse during acrash. Yet, the fuselage design must not be so stiff that ittransmits or amplifies high impact loads to the occupants.ldeally, the design should contain some crushable ele-ments to help limit the loads transmitted to the occupantto survivable or non-injurious levels.The fuselage concept, shown in Figure 1, consists of astiff upper fuselage, a structural floor, and an energy-absorbing subfloor. The upper section of the fuselagecabin is fabricated using a composite sandwich construc-tion and is designed to provide a protective shell thatencloses the occupants in the event of a crash. The en-ergy-absorbing subfloor is designed to dissipate kineticenergy through stable crushing. Finally, a key feature ofthe fuselage concept is the stiff structural floor. Thestructural floor is designed to react the loads generated bycrushing of the subfloor, and to provide a stable platformfor seat and restraint attachment.During the first year of the research program, a 12-in.diameter, 1/5-scale model composite fuselage was de-signed, fabricated, and tested to verify structural andflight-load requirements [3]. During the second year ofthe research program, energy-absorbing subfloor configu-rations were evaluated using quasi-static testing and finiteelement simulation to determine the best design for use inthe 1/5-scale model fuselage concept [4, 5]. During thethird year of the program, a full-scale version of the fuse-lage concept was fabricated, and a vertical drop test wasconducted to validate the scaling process [6]. Test, analy-sis, and correlation with finite element models were per-formed for each test in the series. For the 1/5- and earlyfull-scale drop tests, the inertial loading that normallywould be provided by seats and occupants was representedwith lead weights. In April 2001, a full-scale fuselagesection was tested with two energy-absorbing seats, eachwith an anthropomorphic dummy occupant. The objec-tive of the drop test was to demonstrate the crashworthi-ness of the fuselage concept for a more realistic loadingenvironment using seats and dummies. The data from thedrop test and the development of an integrated crashsimulation are the focus of this paper.Since the completion of the initial research, the compositefuselage section has been used as a test bed for conductingother crash-related experiments. In 2000, two drop testsof a composite fuselage section were performed for thespecific goal of examining test and analysis correlationapproaches for detailed finite element crash simulations[7]. One test was performed from a drop height of 1.75inches to excite the linear frequency response, and testdata were correlated with an MSC.Nastran analysis. Thesecond test was performed for an impact velocity of 25ft/s, and the test data were correlated with a nonlinear,transient dynamic crash simulation. For both tests, thefuselage section was loaded symmetrically using leadmasses that were attached to the floor through seat tracks.The total floor mass was approximately 1000 lbs. The 25ft/s impact test described in Ref. 7 is of particular interestbecause it was performed at the same initial vertical veloc-ity as the fuselage test with seats and dummies describedin this paper. In addition, the fuselage section describedin Ref. 7 had nearly the same floor loading; however,only lead masses were attached to the floor in that test.

Simulation of Aircraft Landing Gears with a Nonlinear Dynamic Finite Element Code

Recent advances in computational speed have made aircraft and spacecraft crash simulations using an explicit, nonlinear, transient-dynamic,e niteelement analysiscodemore feasible.This paperdescribes thedevelopment of a simplelanding-gearmodel, which accurately simulatestheenergy absorbed by thegearwithoutaddingsubstantial complexity to the model. For a crash model the landing gear response is approximated with a spring where the force applied to the fuselage is computed in a user-written subroutine. Helicopter crash simulations using this approach are compared with previously acquired experimental data from a full-scale crash test of a composite helicopter.

IMPACT RESPONSE OF COMPOSITE FUSELAGE FRAMES

Graphite-epoxy frames were drop tested onto a concrete floor to simulate crash loadings. The frames have Z-shaped cross sections typical of designs often proposed for fuselage structure of advanced composite transports. A diameter of six feet for the frames was chosen to reduce specimen fabrication costs and to facilitate testing. Accelerometer, strain gage, and photographic measurements are presented which characterize the impact behavior of frames with differing masses to represent structural or seat/occupant masses. Failures of the graphite-epoxy frames involved complete separations through the cross section. All damage to the lightly loaded composite frames was confined to an area close to the impact point. Subsequent failures left and right of the impact point occurred for the more heavily loaded specimens.

Finite Element Simulation of a Full-Scale Crash Test of a Composite Helicopter

A finite element model of the Sikorsky Advanced Composite Airframe Program (ACAP) helicopter was developed using the nonlinear, explicit transient dynamic code, MSC.Dytran. Analytical predictions were correlated with experimental data obtained from a full-scale crash test of the Sikorsky ACAP helicopter flight test article that was conducted at the Impact Dynamics Research Facility of NASA Langley Research Center. The helicopter was impacted at 38-ft/s vertical and 32.5-ft/s forward velocity with an attitude of 6.25° pitch (nose up) and 3.5°-left roll. The objective of the crash simulation was to evaluate the capabilities of a commercially available transient dynamic code in predicting the response of a composite airframe subjected to impact loading. The model was developed from an existing MSC.Nastran modal-vibration model of the helicopter. Considerable modifications were made in converting the original modal-vibration model to a model for crash simulation. Following conversion of the model, a two-stage modeling approach was used to generate analytical predictions. Due to the relatively long pulse duration, a rigid structural model containing a fairly complex landing gear model was executed from initial contact through landing gear stroke. Prior to fuselage contact, the nodal displacements and velocities were output to a file. Then, a flexible structural model was executed with the nodal displacements and velocities used as initial conditions. This paper describes the development of the finite element crash model, the two-stage modeling approach, and the correlation of the analytical predictions with the experimental data from the full-scale crash test of the ACAP helicopter. ________________________ Presented at the American Helicopter Society 56 Annual Forum, Virginia Beach, Virginia, May 2-4, 2000. This paper is a work of the U. S. Government and is therefore in the public domain. Introduction An important aspect of crashworthiness research is the demonstration and validation of analytical/computational tools for accurate simulation of airframe structural response to crash impacts. The “validation of numerical simulations” was identified as one of five key technology shortfalls during the Workshop on Computational Methods for Crashworthiness [1], which was held at NASA Langley Research Center in 1992. Crash simulation codes can be used during the airframe design phase to certify seats and aircraft to dynamic crash loads, to predict seat and occupant response to impact with the probability of injury, and to evaluate numerous crash scenarios not economically feasible with full-scale crash testing. The US Army has been active in supporting the development and utilization of crash modeling and simulation codes for many decades. More than 25 years ago, the US Army sponsored the initial development of a kinematic crash analysis code, KRASH [2], by the Lockheed-California Company. Kinematic codes employ a semiempirical modeling approach using lumpedmasses, beams, and nonlinear springs to represent the airframe structure. These codes rely heavily on test data for definition of spring properties to characterize the crushing behavior of the subfloor and other structural components. Good correlation between the analytical and experimental data is usually obtained for global parameters, such as engine or landing gear response. However, these codes would be unable to predict localized responses, e.g. the stress level in an airframe component at a particular time during a crash event. Currently, a new generation of crash analysis codes have been developed that will accurately simulate the nonlinear, transient dynamic response of airframe structures. These finite element codes, such as LS-DYNA [3], MSC.Dytran [4], and PAM-CRASH [5], use an explicit solver that eliminates the need to repetitively decompose large global stiffness matrices as is required for implicit codes. Explicit codes require an extremely small time step, typically less than a microsecond, whose duration is controlled by the smallest element in the model. Thus, impact simulations having a pulse duration on the order of 30-40 milliseconds can require several CPU hours to solve on an engineering workstation. Consequently, efficient beam, shell, and solid elements are needed to achieve quick run times for very large models. The new codes are capable of modeling nonlinear geometric behavior including large structural deformations. In addition, these codes are very effective in modeling materials, such as metals, that deform plastically and that have well known failure mechanisms. However, the use of light weight, high strength composite materials for aircraft construction brings with it difficulties in modeling material response and failure behavior. Structural composite materials can exhibit a wide range of material responses from linear elastic to completely nonlinear anisotropic behavior, depending on the individual fiber and matrix properties and laminate stacking sequence. Also, laminated composite materials exhibit a wide variety of failure modes including matrix cracking, fiber failure, and delamination that can occur singly or in combination. These failure modes can change depending on the type and rate of loading. In general, the initial failure event in a single ply of a composite laminate does not produce catastrophic failure. Consequently, the capability to model the progressive failure of composite materials, from initial damage to ultimate failure, is needed. With the increased application of composite materials in the construction of advanced aircraft and rotorcraft, it is important to build confidence in the computational capabilities of these codes through analytical/experimental validation. In 1996, the US Army initiated a fourphase research program to evaluate the capabilities of commercial crash simulation codes for modeling the impact response of a composite helicopter. As part of this program, a finite element crash model of the Sikorsky Advanced Composite Airframe Program (ACAP) helicopter [6, 7] was developed. In this effort, an existing MSC.Nastran [8] modal-vibration model of the helicopter was converted to an MSC.Dytran model for the crash simulation. In 1999, a full-scale crash test of an ACAP helicopter was performed at the Impact Dynamics Research Facility (IDRF) [9] of NASA Langley Research Center to generate experimental data for correlation with the simulation. The present paper will describe: (1) the development of the helicopter crash model, (2) modifications made to the crash model to better represent the test article, (3) a summary of the experimental results from the full-scale crash test of the ACAP helicopter, and (4) the validation of the crash simulation through analytical/experimental correlation. Finite Element Model Development Description of MSC.Dytran Finite Element Code The commercial code, MSC.Dytran, was used to perform the crash simulation of the ACAP helicopter. MSC.Dytran is a three-dimensional finite element code for simulating highly nonlinear transient response of solids, structures, and fluids. The MSC.Patran [10] preand postprocessing code was used with the MSC.Dytran “Preference” to build the finite element crash model and to post-process the results. Description of the Sikorsky Modal-Vibration Model An MSC.Nastran model of the ACAP helicopter that was originally developed for correlation with modal vibration data [11] was obtained from Sikorsky Aircraft. The model, shown in Figure 1, had approximately 5000 nodes, 9,500 elements, 219 material models including many different composite materials, and over 700 different property cards. The elements included 5,453 shell elements; 1,956 beam elements; 1,770 rod elements; and 372 concentrated masses. Because this model was originally used for modal analysis, extensive modifications were required to convert it for a crash analysis. Figure 1. Sikorsky Modal-vibration model of the ACAP helicopter. Conversion of the modal-vibration model to a

Crash Simulation of a Vertical Drop Test of a Commuter-Class Aircraft

Abstract A finite element model of an ATR42-300 commuter-class aircraft was developed and a crash simulation was executed. Analytical predictions were correlated with data obtained from a 30 ft/s (9.14 m/s) vertical drop test of the aircraft. The purpose of the test was to evaluate the structural response of the aircraft when subjected to a severe, but survivable, impact. The aircraft was configured with seats, dummies, luggage, and other ballast. The wings were filled with 8,700 lb. (3,946 kg) of water to represent the fuel. The finite element model, which consisted of 57,643 nodes and 62,979 elements, was developed from direct measurements of the airframe geometry. The seats, dummies, luggage, fuel, and other ballast were represented using concentrated masses. The model was executed in LS-DYNA, a commercial code for performing explicit transient dynamic simulations. Predictions of structural deformation and selected time-history responses were generated. The simulation was successfully validated through extensive test-analysis correlation.

Impact evaluation of composite floor sections

Graphite-epoxy floor sections representative of aircraft fuselage construction were statically and dynamically tested to evaluate their response to crash loadings. These floor sections were fabricated using a frame-stringer design typical of present aluminum aircraft without features to enhance crashworthiness. The floor sections were tested as part of a systematic research program developed to study the impact response of composite components of increasing complexity. The ultimate goal of the research program is to develop crashworthy design features for future composite aircraft. Initially, individual frames of six-foot diameter were tested both statically and dynamically. The frames were then used to construct built-up floor sections for dynamic tests at impact velocities of approximately 20 feet/sec to simulate survivable crash velocities. In addition, static tests were conducted to gain a better understanding of the failure mechanisms seen in the dynamic tests.

A Summary of DOD-Sponsored Research Performed at NASA Langley's Impact Dynamics Research Facility

The Impact Dynamics Research Facility (IDRF) is a 240-ft.-high gantry structure located at NASA Langley Research Center in Hampton, Virginia. The IDRF was originally built in the early 1960s for use as a Lunar Landing Research Facility. As such, the facility was configured to simulate the reduced gravitational environment of the Moon, allowing the Apollo astronauts to practice lunar landings under realistic conditions. In 1985, the IDRF was designated a National Historic Landmark based on its significant contributions to the Apollo Moon Landing Program. In the early 1970s the facility was converted into its current configuration as a full-scale crash test facility for light aircraft and rotorcraft. Since that time, the IDRF has been used to perform a wide variety of impact tests on full-scale aircraft, airframe components, and space vehicles in support of the General Aviation (GA) aircraft industry, the U.S. Department of Defense (DOD), the rotorcraft industry, and the NASA Space program. The objectives of this paper are twofold: to describe the IDRF facility and its unique capabilities for conducting structural impact testing, and to summarize the impact tests performed at the IDRF in support of the DOD. These tests cover a time period of roughly 2 1/2 decades, beginning in 1975 with the full-scale crash test of a CH-47 Chinook helicopter, and ending in 1999 with the external fuel system qualification test of a UH-60 Black Hawk helicopter. NASA officially closed the IDRF in September 2003; consequently, it is important to document the past contributions made in improved human survivability and impact tolerance through DOD-sponsored research performed at the IDRF.