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At least 865 records · Page 48

Effect of Impact Damage and Open Holes on the Compression Strength of Tough Resin/high Strain Fiber Laminates

Tests to assess the performance of composite structures damaged by impact or containing local discontinuities such as cutouts are discussed. In particular, two plate specimen test configurations which were developed to define the serverity of compression strength reduction are described; one for impact damage and the other open hole specimens. The test technique for impact specimens involves damaging the plate at selected energies, measuring the size of damage by ultrasonic C-scan techniques and measuring the residual strength in a compression load test. Open-hole specimen compression tests are conducted for several different hole diameters and the failure strain and load and mode of failure recorded. The plate specimen used in these tests is designed with length, width, thickness and laminate stiffness to ensure that overall plate buckling is not responsible for initiating failure.

Williams, J. G.↗

Residual Strength Analyses of Riveted Lap-Splice Joints

The objective of this paper was to analyze the crack-linkup behavior in riveted-stiffened lap-splice joint panels with small multiple-site damage (MSD) cracks at several adjacent rivet holes. Analyses are based on the STAGS (STructural Analysis of General Shells) code with the critical crack-tip-opening angle (CTOA) fracture criterion. To account for high constraint around a crack front, the "plane strain core" option in STAGS was used. The importance of modeling rivet flexibility with fastener elements that accurately model load transfer across the joint is discussed. Fastener holes are not modeled but rivet connectivity is accounted for by attaching rivets to the sheet on one side of the cracks that simulated both the rivet diameter and MSD cracks. Residual strength analyses made on 2024-T3 alloy (1.6-mm thick) riveted-lap-splice joints with a lead crack and various size MSD cracks were compared with test data from Boeing Airplane Company. Analyses were conducted for both restrained and unrestrained buckling conditions. Comparison of results from these analyses and results from lap-splice-joint test panels, which were partially restrained against buckling indicate that the test results were bounded by the failure loads predicted by the analyses with restrained and unrestrained conditions.

Seshadri, B. R.↗

Testing and Analyses of Advanced Composite Tow-Steered Shells with Cutouts

The structural response of two composite tow-steered shells with small cutouts in end compression is assessed using analyses and experimental measurements. The cylindrical shells were manufactured without cutouts using an automated fiber placement system, where the shells’ fiber orientation angles vary continuously around the shell circumference from ±10 degrees on the axially stiff crown and keel, to ±45 degrees on the shear-stiff sides. The first shell with overlaps has laminate thickness variations on the crown and keel that result from application of all 24 tows during each pass of the fiber placement system. The second shell without overlaps uses the fiber placement system’s tow drop/add capability to achieve a more uniform shell wall thickness. These shells without cutouts were tested in axial compression and buckled elastically. An unreinforced cutout representative of a passenger door on a commercial aircraft fuselage is machined into one side of each shell. Results from nonlinear finite element analysis are compared with their corresponding load, displacement, and strain measurements in prebuckling through stable postbuckling. Results from tests and analyses of shells with large cutouts are also discussed. The very good agreement observed in this study increases confidence in applying tow-steered composites to operational vehicles.

Tow-Steered↗

UltraSail CubeSat Solar Sail Flight Experiment

UltraSail is a next-generation, highrisk, high-payoff sail system for the launch, deployment, stabilization, and control of very large (km2 class) solar sails enabling high payload mass fractions for interplanetary and deep space spacecraft. UltraSail is a non-traditional approach to propulsion technology achieved by combining propulsion and control systems developed for formation- flying microsatellites with an innovative solar sail architecture to achieve controllable sail areas approaching 1 km2, sail subsystem area densities approaching 1 g/m2, and thrust levels many times those of ion thrusters used for comparable deep space missions. UltraSail can achieve outer planetary rendezvous, a deep-space capability now reserved for high-mass nuclear and chemical systems. There is a twofold rationale behind the UltraSail concept for advanced solar sail systems. The first is that sail-andboom systems are inherently size-limited. The boom mass must be kept small, and column buckling limits the boom length to a few hundred meters. By eliminating the boom, UltraSail not only offers larger sail area, but also lower areal density, allowing larger payloads and shorter mission transit times. The second rationale for UltraSail is that sail films present deployment handling difficulties as the film thickness approaches one micrometer. The square sail requires that the film be folded in two directions for launch, and similarly unfolded for deployment. The film is stressed at the intersection of two folds, and this stress varies inversely with the film thickness. This stress can cause the film to yield, forming a permanent crease, or worse, to perforate. By rolling the film as UltraSail does, creases are prevented. Because the film is so thin, the roll thickness is small. Dynamic structural analysis of UltraSail coupled with dynamic control analysis shows that the system can be designed to eliminate longitudinal torsional waves created while controlling the pitch of the blades, while using solar photon pressure to slew the spin axis. Vacuum tests have also verified that electrostatic and molecular adhesion forces can substantially be eliminated by making the film electrically conductive, reducing the peel force of the film off the storage roll to levels of 100s of micro-N. The innovation demonstrated the capability of deploying a six-micron aluminum- coated film from a reel through a slit in vacuum. The innovation also demonstrated a spin-stabilized method for deploying a long reel of solar sail film using solar pressure to spin-up and orbit raise the satellite, and also a gravity gradient method for deploying a long reel of solar sail film using solar pressure to orbit raise the satellite. The solar sail mass fraction of 25% is consistent with high specific impulse ion systems, but without the added weight and cost of a power source and processing unit. The large sail area, coupled with low film density, is giving UltraSail a high payload fraction. The UltraSail deployment scheme unrolls a micrometerscale reflection-coated polyimide film from a storage mandrel to a maximum length of several kilometers with the aid of a blade tip satellite.

Carroll, David↗

Panel flutter in a low-density atmosphere

The problem of panel flutter is considered in the case of a free-molecule flow, the effect of the aerodynamic shear stress being incorporated into the structural model through distributed longitudinal and bending moment loads. This kind of analysis is relevant in the case of hypersonic flight vehicles, like the NASP, especially because in these conditions the shear stress can be expected to be considerably larger that the pressure at a given point. The aerodynamic loading is derived assuming a quasi-steady approximation. Two important parameters are given by alpha-m, the 'momentum accommodation coefficient', and Theta, the temperature ratio between the panel temperature and the temperature of the undisturbed flow. For high enough values of Theta the variation of the linear flutter parameter with alpha-m is close to linear. Comparison with continuum theory, given by linear piston theory, depends on the values of Theta and alpha-m considered. Finally, it is shown that buckling with respect to a uniform distributed longitudinal load is stabilizing.

Resende, Hugo B.↗

Computer design synthesis of a below knee-Syme prosthesis

A detailed design synthesis analysis of the BK Syme prosthesis is provided, to determine the socket's cutout orientation size and shape, cutout fillet shape, socket wall thickness distribution and the reinforced fiber distribution in the socket wall, for a minimally stressed structurally safe lightweight prosthesis. For analysis purposes, the most adverse socket loading is obtained at the push-off stage of gait; this loading is idealized as an axial in-plane loading on the bottom edge of the circular cylindrical socket shell whose top edge is considered fixed. Finite element stress analysis of the socket shell (with uniform and graded wall thickness) are performed for various orientations of the cutout and for various types of corner fillets. A lateral cutout with a streamline fillet is recommended. The wall material (i.e., thickness) distribution is determined so as to minimize the stresses, while ensuring that the wall material's stress limits are not exceeded. For such a maximally stressed lightweight socket shell, the panels in the neighborhood of the cutout are checked to ensure that they do not buckle under their acquired stresses. A fiber-reinforced laminated composite socket shell is also analyzed in order to recommend optimum variables in orientations and densities of reinforcing fibers.

Elangovan, P. T.↗

Design of Fiber Reinforced Foam Sandwich Panels for Large Ares V Structural Applications

The preliminary design of three major structural components within NASA's Ares V heavy lift vehicle using a novel fiber reinforced foam composite sandwich panel concept is presented. The Ares V payload shroud, interstage, and core intertank are designed for minimum mass using this panel concept, which consists of integral composite webs separated by structural foam between two composite facesheets. The HyperSizer structural sizing software, in conjunction with NASTRAN finite element analyses, is used. However, since HyperSizer does not currently include a panel concept for fiber reinforced foam, the sizing was performed using two separate approaches. In the first, the panel core is treated as an effective (homogenized) material, whose properties are provided by the vendor. In the second approach, the panel is treated as a blade stiffened sandwich panel, with the mass of the foam added after completion of the panel sizing. Details of the sizing for each of the three Ares V components are given, and it is demonstrated that the two panel sizing approaches are in reasonable agreement for thinner panel designs, but as the panel thickness increases, the blade stiffened sandwich panel approach yields heavier panel designs. This is due to the effects of local buckling, which are not considered in the effective core property approach.

Bednarcyk, Brett A.↗

Innovative design of composite structures: The use of curvilinear fiber format in structural design of composites

The influence is investigated of a curvilinear fiber format on load carrying capacity of a layered fiber reinforced plate with a centrally located hole. A curvilinear fiber format is descriptive of layers in a laminate having fibers which are aligned with the principal stress directions in those layers. Laminates of five curvilinear fiber format designs and four straightline fiber format designs are considered. A quasi-isotropic laminate having a straightline fiber format is used to define a baseline design for comparison with the other laminate designs. Four different plate geometries are considered and differentiated by two values of hole diameter/plate width equal to 1/6 and 1/3, and two values of plate length/plate width equal to 2 and 1. With the plates under uniaxial tensile loading on two opposing edges, alignment of fibers in the curvilinear layers with the principal stress directions is determined analytically by an iteration procedure. In-plane tensile load capacity is computed for all of the laminate designs using a finite element analysis method. A maximum strain failure criterion and the Tsai-Wu failure criterion are applied to determine failure loads and failure modes. Resistance to buckling of the laminate designs to uniaxial compressive loading is analyzed using the commercial code Engineering Analysis Language. Results indicate that the curvilinear fiber format laminates have higher in-plane tensile load capacity and comparable buckling resistance relative to the straightline fiber format laminates.

Charette, R. F.↗

Innovative design of composite structures: Design, manufacturing, and testing of plates utilizing curvilinear fiber trajectories

As a means of improving structural design, the concept of fabricating flat plates containing holes by incorporating curvilinear fiber trajectories to transmit loads around the hole is studied. In the present discussion this concept is viewed from a structural level, where access holes, windows, doors, and other openings are of significant size. This is opposed to holes sized for mechanical fasteners. Instead of cutting the important load-bearing fibers at the hole edge, as a conventional straightline design does, the curvilinear design preserves the load-bearing fibers by orienting them in smooth trajectories around the holes, their loading not ending abruptly at the hole edge. Though the concept of curvilinear fiber trajectories has been studied before, attempts to manufacture and test such plates have been limited. This report describes a cooperative effort between Cincinnati Milacron Inc., NASA Langley Research Center, and Virginia Polytechnic Institute and State University to design, manufacture, and test plates using the curvilinear fiber trajectory concept. The paper discusses details of the plate design, details of the manufacturing, and a summary of results from testing the plates with inplane compressive buckling loads and tensile loads. Comparisons between the curvilinear and conventional straightline fiber designs based on measurements and observation are made. Failure modes, failure loads, strains, deflections, and other key responses are compared.

Hyer, M. W.↗

Residual Strength Analyses of Monolithic Structures

Finite-element fracture simulation methodology predicts the residual strength of damaged aircraft structures. The methodology uses the critical crack-tip-opening-angle (CTOA) fracture criterion to characterize the fracture behavior of the material. The CTOA fracture criterion assumes that stable crack growth occurs when the crack-tip angle reaches a constant critical value. The use of the CTOA criterion requires an elastic- plastic, finite-element analysis. The critical CTOA value is determined by simulating fracture behavior in laboratory specimens, such as a compact specimen, to obtain the angle that best fits the observed test behavior. The critical CTOA value appears to be independent of loading, crack length, and in-plane dimensions. However, it is a function of material thickness and local crack-front constraint. Modeling the local constraint requires either a three-dimensional analysis or a two-dimensional analysis with an approximation to account for the constraint effects. In recent times as the aircraft industry is leaning towards monolithic structures with the intention of reducing part count and manufacturing cost, there has been a consistent effort at NASA Langley to extend critical CTOA based numerical methodology in the analysis of integrally-stiffened panels.In this regard, a series of fracture tests were conducted on both flat and curved aluminum alloy integrally-stiffened panels. These flat panels were subjected to uniaxial tension and during the test, applied load-crack extension, out-of-plane displacements and local deformations around the crack tip region were measured. Compact and middle-crack tension specimens were tested to determine the critical angle (wc) using three-dimensional code (ZIP3D) and the plane-strain core height (hJ using two-dimensional code (STAGS). These values were then used in the STAGS analysis to predict the fracture behavior of the integrally-stiffened panels. The analyses modeled stable tearing, buckling, and crack branching at the integral stiffener using different values of critical CTOA for different material thicknesses and orientation. Comparisons were made between measured and predicted load-crack extension, out-of-plane displacements and local deformations around the crack tip region. Simultaneously, three-dimensional capabilities to model crack branching and to monitor stable crack growth of multiple cracks in a large thick integrally-stiffened flat panels were implemented in three-dimensional finite element code (ZIP3D) and tested by analyzing the integrally-stiffened panels tested at Alcoa. The residual strength of the panels predicted from STAGS and ZP3D code compared very well with experimental data. In recent times, STAGS software has been updated with new features and now one can have combinations of solid and shell elements in the residual strength analysis of integrally-stiffened panels.

Forth, Scott↗

Application of multivariable search techniques to structural design optimization

Multivariable optimization techniques are applied to a particular class of minimum weight structural design problems: the design of an axially loaded, pressurized, stiffened cylinder. Minimum weight designs are obtained by a variety of search algorithms: first- and second-order, elemental perturbation, and randomized techniques. An exterior penalty function approach to constrained minimization is employed. Some comparisons are made with solutions obtained by an interior penalty function procedure. In general, it would appear that an interior penalty function approach may not be as well suited to the class of design problems considered as the exterior penalty function approach. It is also shown that a combination of search algorithms will tend to arrive at an extremal design in a more reliable manner than a single algorithm. The effect of incorporating realistic geometrical constraints on stiffener cross-sections is investigated. A limited comparison is made between minimum weight cylinders designed on the basis of a linear stability analysis and cylinders designed on the basis of empirical buckling data. Finally, a technique for locating more than one extremal is demonstrated.

Jones, R. T.↗

Discovery of Interstellar Anions in Cepheus and Auriga

We report the detection of microwave emission lines from the hydrocarbon anion C6H(-) and its parent neutral C6H in the star-forming region LI251 A (in Cepheus), and the pre-stellar core LI512 (in Auriga). The carbon chain-bearing species C4H, HC3N, HC5N, HC7N, and C3S are also detected in large abundances. The observations of L1251A constitute the first detections of anions and long-chain polyynes and cyanopolyynes (with more than five carbon atoms) in the Cepheus Flare star-forming region, and the first detection of anions in the vicinity of a protostar outside of the Taurus molecular cloud complex, indicating a possible wider importance for anions in the chemistry of star formation. Rotational excitation temperatures have been derived from the HC3N hyperfine structure lines and are found to be 6.2 K for L1251A and 8.7 K for LI5l2. The anion-to-neutral ratios are 3.6% and 4.1%, respectively, which are within the range of values previously observed in the interstellar medium, and suggest a relative uniformity in the processes governing anion abundances in different dense interstellar clouds. This research contributes toward the growing body of evidence that carbon chain anions are relatively abundant in interstellar clouds throughout the Galaxy, but especially in the regions of relatively high density and high depletion surrounding pre-stellar cores and young, embedded protostars.

Cordiner, M. A.↗

Carbon-Nanotube-Carpet Heat-Transfer Pads

Microscopic thermal-contact pads that include carpet-like arrays of carbon nanotubes have been invented for dissipating heat generated in integrated circuits and similarly sized single electronic components. The need for these or other innovative thermal-contact pads arises because the requisite high thermal conductances cannot be realized by scaling conventional macroscopic thermal-contact pads down to microscopic sizes. Overcoming limitations of conventional thermal-contact materials and components, the carbon-nanotube thermal-contact pads offer the high thermal conductivities needed to accommodate the high local thermal power densities of modern electronic circuits, without need for large clamping pressures, extreme smoothness of surfaces in contact, or gap-filling materials (e.g., thermally conductive greases) to ensure adequate thermal contact. Moreover, unlike some conventional thermal-contact components, these pads are reusable. The figure depicts a typical pad according to the invention, in contact with a rough surface on an electronic component that is to be cooled. Through reversible bending and buckling of carbon nanotubes at asperities on the rough surface, the pad yields sufficiently, under relatively low contact pressure, that thermal contact is distributed to many locations on the surface to be cooled, including valleys where contact would not ordinarily occur in conventional clamping of rigid surfaces. Hence, the effective thermal-contact area is greater than that achievable through scaling down of a macroscopic thermal-contact pad. The extremely high longitudinal thermal conductivities of the carbon nanotubes are utilized to conduct heat away from potential hot spots on the surface to be cooled. The fibers protrude from a layer of a filler material (Cu, Ag, Au, or metal-particle- filled gels), which provides both mechanical support to maintain the carbon nanotubes in alignment and thermal conductivity to enhance the diffusion of concentrated heat from the nanotubes into the larger adjacent volume of a heat sink. The array of carbon nanotubes, the filler material, and the heat sink are parts of a unitary composite structure that is fabricated as follows: 1. Using techniques that have been reported previously, the array of substantially perpendicularly oriented carbon nanotubes is grown on a metal, silicon, or other suitable thermally conductive substrate that is intended to become the heat sink. 2. By means of chemical vapor deposition, physical vapor deposition, plasma deposition, ion sputtering, electrochemical deposition, or casting from a liquid phase, some or all of the interstitial volume between carbon nanotubes is filled with the aforementioned layer of mechanically supporting, thermally conductive material. 3. To cause the carbon nanotubes to protrude the desired length from the filler material, an outer layer of filler is removed by mechanical polishing, chemical mechanical polishing, wet chemical etching, electrochemical etching, or dry plasma etching.

Li, Jun↗

Blended-Wing-Body (BWB) Fuselage Structural Design for Weight Reduction

Structural analysis and design of efficient pressurized fuselage configurations for the advanced Blended-Wing-Body (BWB) flight vehicle is a challenging problem. Unlike a conventional cylindrical pressurized fuselage, stress level in a box type BWB fuselage is an order of magnitude higher, because internal pressure primarily results in bending stress instead of skin-membrane stress. In addition, resulting deformation of aerodynamic surface could significantly affect performance advantages provided by lifting body. The pressurized composite conformal multi-lobe tanks of X-33 type space vehicle also suffered from similar problem. In the earlier BWB design studies, Vaulted Ribbed Shell (VLRS), Flat Ribbed Shell (FRS); Vaulted shell Honeycomb Core (VLHC) and Flat sandwich shell Honeycomb Core (FLHC) concepts were studied. The flat and vaulted ribbed shell concepts were found most efficient. In a recent study, a set of composite sandwich panel and cross-ribbed panel were analyzed. Optimal values of rib and skin thickness, rib spacing, and panel depth were obtained for minimal weight under stress and buckling constraints. In addition, a set of efficient multi-bubble fuselage (MBF) configuration concept was developed. The special geometric configuration of this concept allows for balancing internal cabin pressure load efficiently, through membrane stress in inner-stiffened shell and inter-cabin walls, while the outer-ribbed shell prevents buckling due to external resultant compressive loads. The initial results from these approximate finite element analyses indicate progressively lower maximum stresses and deflections compared to the earlier study. However, a relative comparison of the FEM weight per unit floor area of the segment unit indicates that the unit weights are still relatively higher that the conventional B777 type cylindrical or A380 type elliptic fuselage design. Due to the manufacturing concern associated with multi-bubble fuselage, a Y braced box-type fuselage alternative with special resin-film injected (RFI) stitched carbon composite with foam-core was designed by Boeing under a NASA research contract for the 480 passenger version. It is shown that this configuration can be improved to a modified multi-bubble fuselage which has better stress distribution, for same material and dimension.

Mukhopadhyay, V.↗

Tow-Steered Panels With Holes Subjected to Compression or Shear Loads

Tailoring composite laminates to vary the fiber orientations within a fiber layer of a laminate to address non-uniform stress states and provide structural advantages such as the alteration of principal load paths has potential application to future low-cost, light-weight structures for commercial transport aircraft. Evaluation of this approach requires the determination of the effectiveness of stiffness tailoring through the use of curvilinear fiber paths in flat panels including the reduction of stress concentrations around the holes and the increase in load carrying capability. Panels were designed through the use of an optimization code using a genetic algorithm and fabricated using a tow-steering approach. Manufacturing limitations, such as the radius of curvature of tows the machine could support, avoidance of wrinkling of fibers and minimization of gaps between fibers were considered in the design process. Variable stiffness tow-steered panels constructed with curvilinear fiber paths were fabricated so that the design methodology could be verified through experimentation. Finite element analysis where each element s stacking sequence was accurately defined is used to verify the behavior predicted based on the design code. Experiments on variable stiffness flat panels with central circular holes were conducted with the panels loaded in axial compression or shear. Tape and tow-steered panels are used to demonstrate the buckling, post-buckling and failure behavior of elastically tailored panels. The experimental results presented establish the buckling performance improvements attainable by elastic tailoring of composite laminates.

Jegley, Dawn C.↗

Sizing-stiffened composite panels loaded in the postbuckling range

Stiffened panels are widely used in aircraft structures such as wing covers, fuselages, control surfaces, spar webs, bulkheads, and floors. The detailed sizing of minimum-weight stiffened panels involves many considerations. Use of composite materials introduces additional complexities. Many potential modes of failure exist. Analyses for these modes are often not trivial, especially for those involving large out-of-plane displacements. Accurate analyses of all potential failure modes are essential. Numerous practical constraints arise from manufacturing/cost considerations and from damage tolerance, durability, and stiffness requirements. The number of design variables can be large when lamina thicknesses and stacking sequence are being optimized. A significant burden is placed on the sizing code due to the complex analyses, practical constraints, and number of design variables. On the other hand, sizing weight-efficient panels without the aid of an automated procedure is almost out of the question. The sizing code postbuckled Open-Stiffener Optimum Panels (POSTOP) has been developed to aid in the design of minimum-weight panels subject to the considerations mentioned above. Developed for postbuckled composite panels, POSTOP may be used for buckling resistant panels and metallic panels as well. The COPES/CONMIN optimizer is used in POSTOP although other options such as those in the ADS system could be substituted with relative ease. The basic elements of POSTOP are shown. Some of these elements and usage of the program are described.

Biggers, S. B.↗

Finite Element Modeling for Compression Strength Prediction of Honeycomb Cores with Geometric Imperfections Measured using X-ray CT Imaging

Aluminum honeycomb cores have been used extensively in composite sandwich panels due to having high bending rigidity while maintaining low density. During manufacturing of the aluminum honeycomb, the thin metal walls are susceptible to imperfections that deviate from an ideal honeycomb shape. Computational tools to quantify imperfections and investigate their effects on quasi-static compression and impact response can inform design criteria to construct safer and lighter launch vehicle structures. This paper presents the results of a study on finite element modeling of metallic honeycomb cores (HCC) with geometric imperfections measured using X-ray Computer Tomography (CT), to predict the compression response. Finite element (FE) models are constructed with measured imperfections with appropriate boundary conditions. It is shown that the average behavior of larger 10x11 cell models can be predicted by sampling single cell models from the larger domain. Amplification of imperfections for cells at the center of sampled specimens are mode switched from the larger out-of-plane imperfection to the mode shape of the adjacent cells resulting in strengthening of the center specimens.

aluminum honeycomb↗

Finite Element Modeling and Analysis of Mars Entry Aeroshell Baseline Concept

The structure that is developed and analyzed in this project must be able to survive all the various load conditions that it will encounter along its course to Mars with the minimal amount of weight and material. At this stage, the goal is to study the capability of the structure using a finite element model (FEM). This FEM is created using a python script, and is numerically solved in Nastran. The purpose of the model is to achieve an optimization of mass given specific constraints on launch and entry. The generation and analysis of the baseline Rigid Mid-Range Lift to Drag Ratio Aeroshell model is a continuation and an improvement on previous work done for the FEM. The model is generated using Python programming with the axisymmetric placement of nodes for beam and shell elements. The shells are assigned a honeycomb sandwich material with an aluminum honeycomb core and composite face sheets, and the beams are assigned the same material as the shell face sheets. There are two load cases assigned to the model: Earth launch and Mars entry. The Earth launch case consists of pressure, gravity, and vibration loads, and the Mars entry case consists of just pressure and gravity loads. The Earth launch case was determined to be the driving case, though the analyses are performed for both cases to ensure the constraints are satisfied. The types of analysis performed with the model are design optimization, statics, buckling, normal modes, and frequency response, the last of which is only for the Earth launch load case. The final results indicated that all of the requirements are satisfied except the thermal limits, which could not yet be tested, and the normal modes for the Mars entry. However, the frequency limits during Mars entry are expected to be much higher than the lower frequency limits set for the analysis. In addition, there are still improvements that can be made in order to reduce the weight while still meeting all requirements.

Ahmed, Samee W.↗