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NASA NTRS2003

Optimal Design and Damage Tolerance Verification of an Isogrid Structure for Helicopter Application

A composite isogrid panel design for application to a rotorcraft fuselage is presented. An optimum panel design for the lower fuselage of the rotorcraft that is subjected to combined in-plane compression and shear loads was generated using a design tool that utilizes a smeared-stiffener theory in conjunction with a genetic algorithm. A design feature was introduced along the edges of the panel that facilitates introduction of loads into the isogrid panel without producing undesirable local bending gradients. A low-cost manufacturing method for the isogrid panel that incorporates these design details is also presented. Axial compression tests were conducted on the undamaged and low-speed impact damaged panels to demonstrate the damage tolerance of this isogrid panel. A combined loading test fixture was designed and utilized that allowed simultaneous application of compression and shear loads to the test specimen. Results from finite element analyses are presented for the isogrid panel designs and these results are compared with experimental results. This study illustrates the isogrid concept to be a viable candidate for application to the helicopter lower fuselage structure.

Baker, Donald J.↗

NASA NTRS1989

Experimental measurement of structural power flow on an aircraft fuselage

An experimental technique is used to measure the structural power flow through an aircraft fuselage with the excitation near the wing attachment location. Because of the large number of measurements required to analyze the whole of an aircraft fuselage, it is necessary that a balance be achieved between the number of measurement transducers, the mounting of these transducers, and the accuracy of the measurements. Using four transducers mounted on a bakelite platform, the structural intensity vectors at locations distributed throughout the fuselage are measured. To minimize the errors associated with using a four transducers technique the measurement positions are selected away from bulkheads and stiffeners. Because four separate transducers are used, with each transducer having its own drive and conditioning amplifiers, phase errors are introduced in the measurements that can be much greater than the phase differences associated with the measurements. To minimize these phase errors two sets of measurements are taken for each position with the orientation of the transducers rotated by 180 deg and an average taken between the two sets of measurements. Results are presented and discussed.

Cuschieri, J. M.↗

NASA NTRS1991

Experimental measurement of structural power flow on an aircraft fuselage

An experimental technique is used to measure structural intensity through an aircraft fuselage with an excitation load applied near one of the wing attachment locations. The fuselage is a relatively large structure, requiring a large number of measurement locations to analyze the whole of the structure. For the measurement of structural intensity, multiple point measurements are necessary at every location of interest. A tradeoff is therefore required between the number of measurement transducers, the mounting of these transducers, and the accuracy of the measurements. Using four transducers mounted on a bakelite platform, structural intensity vectors are measured at locations distributed throughout the fuselage. To minimize the errors associated with using the four transducer technique, the measurement locations are selected to be away from bulkheads and stiffeners. Furthermore, to eliminate phase errors between the four transducer measurements, two sets of data are collected for each position, with the orientation of the platform with the four transducers rotated by 180 degrees and an average taken between the two sets of data. The results of these measurements together with a discussion of the suitability of the approach for measuring structural intensity on a real structure are presented.

Cuschieri, J. M.↗

NASA NTRS1946

Stresses in and General Instability of Monocoque Cylinders with Cutouts I : Experimental Investigation of Cylinders with a Symmetric Cutout Subjected to Pure Bending

Ten 24S-T alclad cylinders of 20-inch diameter, 45- or 58-inch length, and 0.012-inch wall thickness, reinforced with 24S-T aluminum alloy stringers and rings were tested in pure bending. In the middle of the compression side of the cylinders there was a cutout extending over 19 inches in the longitudinal direction, and over an angle of 45 degrees, 90 degrees, or 135 degrees in the circumferential direction. The strain in the stringers and in the sheet covering was measured with metal electric strain gages. The stress distribution in the cylinders deviate considerably from the linear law valid for cylinders without a cutout. The maximum strain measured was about four-thirds of the value calculated from the Mc/I formula when I was taken as the moment of inertia of the cross section of the portion of the cylinder where the cutout was situated. A diagram is presented containing the strain factors defined as the ratios of measured strain to strain calculated with the Mc/I formula. All the 10 cylinders tested failed in general instability. Two symmetric and one antisymmetric pattern of buckling were observed and the buckling load appeared to be independent of the method of manufacture and the length of the cylinder. The buckling load of the cylinders having cutouts extending over 45 degrees, 90 degrees, and 135 degrees was 66, 47, and 31 percent, respectively, of the buckling load of the cylinder without a cutout.

CYLINDERS, MONOCOQUE-BUCKLING↗

NASA NTRS1991

Preliminary results on the fracture analysis of multi-site cracking of lap joints in aircraft skins

This paper presents results of a fracture mechanics analysis relevant to fatigue crack growth at rivets in lap joints of aircraft skins are presented. Multi-site damage (MSD) is receiving increased attention within the context of problems of aging aircraft. Fracture analyses previously carried out include small-scale modeling of rivet/skin interactions, larger-scale two-dimensional models of lap joints similar to that developed here, and full scale three-dimensional models of large portions of the aircraft fuselage. Fatigue testing efforts have included flat coupon specimens, two-dimensional lap joint tests, and full scale tests on specimens designed to closely duplicate aircraft sections. Most of this work is documented in the proceedings of previous symposia on the aging aircraft problem. The effect MSD has on the ability of skin stiffeners to arrest the growth of long skin cracks is a particularly important topic that remains to be addressed. One of the most striking features of MSD observed in joints of some test sections and in the joints of some of the older aircraft fuselages is the relative uniformity of the fatigue cracks from rivet to rivet along an extended row of rivets. This regularity suggests that nucleation of the cracks must not be overly difficult. Moreover, it indicates that there is some mechanism which keeps longer cracks from running away from shorter ones, or, equivalently, a mechanism for shorter cracks to catch-up with longer cracks. This basic mechanism has not been identified, and one of the objectives of the work is to see to what extent the mechanism is revealed by a fracture analysis of the MSD cracks. Another related aim is to present accurate stress intensity factor variations with crack length which can be used to estimate fatigue crack growth lifetimes once cracks have been initiated. Results are presented which illustrate the influence of load shedding from rivets with long cracks to neighboring rivets with shorter cracks. Results will also be included for the effect of residual stress due to the riveting process itself. All the results presented in this paper should be regarded as preliminary in the sense that further work by ourselves and others will be required to validate them and to establish their sensitivity to more embellished modelling assumptions.

Aging Aircraft↗

NASA NTRS2006

Mode I Toughness Measurements of Core/Facesheet Bonds in Honeycomb Sandwich Structures

Composite sandwich structures will be used in many future applications in aerospace, marine and offshore industries due to the fact that the strength and stiffness to mass ratios surpass any other structural type. Sandwich structure also offers advantages over traditional stiffened panels such as ease of manufacturing and repair. During the last three decades, sandwich structure has been used extensively for secondary structure in aircraft (fuselage floors, rudders and radome structure). Sandwich structure is also used as primary structure in rotorcraft, the most common example being the trailing edge of rotor blades. As with other types of composite construction, sandwich structure exhibits several types of failure mode such as facesheet wrinkling, core crushing and sandwich buckling. Facesheet/core debonding has also been observed in the marine and aerospace industry. During this failure mode, peel stresses applied to an existing facesheet/core debond or an interface low in toughness, results in the facesheet being peeled from the core material, possibly leading to a significant loss in structural integrity of the sandwich panel. In an incident during a test on a liquid hydrogen fuel tank of the X-33 prototype vehicle, the outer graphite/epoxy facesheet and honeycomb core became debonded from the inner facesheet along significant areas, leading to failure of the tank. As a consequence of the accident; significant efforts were made to characterize the toughness of the facesheet/core bond. Currently, the only standardized method available for assessing the quality of the facesheet/core interface is the climbing drum peel test (ASTM D1781). During this test a sandwich beam is removed from a panel and the lip of one of the facesheets is attached to a drum, as shown in Fig. 1. The drum is then rotated along the sandwich beam, causing the facesheet to peel from the core. This method has two major drawbacks. First, it is not possible to obtain quantitative fracture data from the test and so the results can only be used in a qualitative manner. Second, only sandwich structure with thin facesheets can be tested (to facilitate wrapping of the facesheet around the climbing drum). In recognition of the need for a more quantitative facesheet/core fracture test, several workers have devised experimental techniques for characterizing the toughness of the facesheet/core interface. In all of these cases, the tests are designed to yield a mode I-dominated fracture toughness of the facesheet/core interface in a manner similar to that used to determine mode I fracture toughness of composite laminates. In the current work, a modified double cantilever beam is used to measure the mode I-dominated fracture toughness of the interface in a sandwich consisting of glass/phenolic honeycomb core reinforced with graphite epoxy facesheets. Two specimen configurations were tested as shown in Fig 2. The first configuration consisted of reinforcing the facesheets with aluminum blocks (Fig. 2a). In the second configuration unreinforced specimens were tested (Fig. 2b). Climbing drum peel tests were also conducted to compare the fracture behavior observed between this test and the modified double cantilever beam. This paper outlines the test procedures and data reduction strategies used to compute fracture toughness values from the tests. The effect of specimen reinforcement on fracture toughness of the facesheet/core interface is discussed.

Nettles, Alan T.↗

NASA NTRS1997

Novel Composites for Wing and Fuselage Applications: Speedy Nonlinear Analysis of Postbuckled Panels in Shear (SNAPPS)

A simple and rapid analysis method, consisting of a number of modular, 'strength-of-materials-type' models, is presented for predicting the nonlinear response and stiffener separation of postbuckled, flat, composite, shear panels. The analysis determines the maximum principal tensile stress in the skin surface layer under to toe. Failure is said to occur when this stress reaches the mean transverse tensile strength of the layer. The analysis methodology consists of a number of closed-form equations that can easily be used in a 'hand analysis. For expediency, they have been programmed into a preliminary design code called SNAPPS (Speedy Nonlinear Analysis of Postbuckled Panels in Shear), which rapidly predicts postbuckling response of the panel for each value of the applied shear load. SNAPPS response and failure predictions were found to agree well with test results for three panels with widely different geometries, laminates and stiffnesses. Design guidelines are given for increasing the load-carrying capacity of stiffened, composite shear panels.

Sharp, Dave↗

NASA NTRS2013

Preliminary Design and Analysis of an In-plane PRSEUS Joint

As part of the National Aeronautics and Space Administration's (NASA's) Environmentally Responsible Aviation (ERA) program, the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) has been designed, developed and tested. However, PRSEUS development efforts to date have only addressed joints required to transfer bending moments between PRSEUS panels. Development of in-plane joints for the PRSEUS concept is necessary to facilitate in-plane transfer of load from PRSEUS panels to an adjacent structure, such as from a wing panel into a fuselage. This paper presents preliminary design and analysis of an in-plane PRSEUS joint for connecting PRSEUS panels at the termination of the rod-stiffened stringers. Design requirements are provided, the PRSEUS blade joint concept is presented, and preliminary design changes and analyses are carried out to examine the feasibility of the proposed in-plane PRSEUS blade joint. The study conducted herein focuses mainly on the PRSEUS structure on one side of the joint. In particular, the design requirements for the rod shear stress and bolt bearing stress are examined. A PRSEUS blade joint design was developed that demonstrates the feasibility of this in-plane PRSEUS joint concept to terminate the rod-stiffened stringers. The presented design only demonstrates feasibility, therefore, some areas of refinement are presented that would lead to a more optimum and realistic design.

Lovejoy, Andrew E.↗

NASA NTRS2016

Acoustic Emission of Large PRSEUS Structures (Pultruded Rod Stitched Efficient Unitized Structure)

In the role of structural health monitoring (SHM), Acoustic Emission (AE) analysis is being investigated as an effective method for tracking damage development in large composite structures under load. Structures made using Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) for damage tolerant, light, and economical airframe construction are being pursued by The Boeing Company and NASA under the Environmentally Responsible Aircraft initiative (ERA). The failure tests of two PRSEUS substructures based on the Boeing Hybrid Wing Body fuselage concept were conducted during third quarter 2011 and second quarter 2015. One fundamental concern of these tests was determining the effectiveness of the stitched integral stiffeners to inhibit damage progression. By design, severe degradation of load carrying capability should not occur prior to Design Ultimate Load (DUL). While minor damage prior to DUL was anticipated, the integral stitching should not fail since this would allow a stiffener-skin delamination to progress rapidly and alter the transfer of load into the stiffeners. In addition, the stiffeners should not fracture because they are fundamental to structural integrity. Getting the best information from each AE sensor is a primary consideration because a sparse network of sensors is implemented. Sensitivity to stiffener-contiguous degradation is supported by sensors near the stiffeners, which increases the coverage per sensor via AE waveguide actions. Some sensors are located near potentially critical areas or "critical zones" as identified by numerical analyses. The approach is compared with the damage progression monitored by other techniques (e.g. ultrasonic C-scan).

Horne, Michael R.↗

NASA NTRS2023

Assessing Flow Formability of Aerospace Aluminum Alloys via DIC Tensile Testing

Over the past decade, NASA Langley Research Center (LaRC) has championed integrally stiffened cylinder (ISC) technology for single-piece, cryogenic tank barrels on launch vehicles. The current investigation aims to extend the hybrid spin/shear/flow forming process to aircraft fuselage structures, a damage tolerance critical application. The WF Maschinenbau VUD-600® vertical spin/flow forming facility recently established at LaRC represents a reasonable sub-scale facsimile of the ISC deformation process for research and development (R&D) purposes. The objective of this study is to explore whether tensile testing with digital image correlation (DIC) is an effective way to rank the formability of candidate aerospace Al alloys and expedite empirical forming trials. Specific tensile data, such as reduction of area, strain hardening exponent, and modulus of resilience, are used as formability metrics for a variety of alloy/temper/product form combinations. Results from high-strength aluminum alloys AA 2139, AA 2050, AA 2043 and AA 2219 are compared with the medium-strength, highly formable AA 6061 benchmark. Rolled, forged and cast preform materials in both the -O temper (fully annealed) and -T4 temper (solution-treated, quenched, and naturally aged) conditions are evaluated. AA 2139 plate in the -T4 temper emerges as the top-ranked material, based on the criteria selected. Starting with preforms in the -T4 temper will result in flow-formed material exhibiting mechanical properties closer to aircraft fuselage requirements. The optimum balance between strength and damage tolerance may also be achieved via post-forming procedures that avoid quenching and stretching.

Aluminum alloys↗

NASA NTRS2012

Hybrid Wing-Body (HWB) Pressurized Fuselage Modeling, Analysis, and Design for Weight Reduction

This paper describes the interim progress for an in-house study that is directed toward innovative structural analysis and design of next-generation advanced aircraft concepts, such as the Hybrid Wing-Body (HWB) and the Advanced Mobility Concept-X flight vehicles, for structural weight reduction and associated performance enhancement. Unlike the conventional, skin-stringer-frame construction for a cylindrical fuselage, the box-type pressurized fuselage panels in the HWB undergo significant deformation of the outer aerodynamic surfaces, which must be minimized without significant structural weight penalty. Simple beam and orthotropic plate theory is first considered for sizing, analytical verification, and possible equivalent-plate analysis with appropriate simplification. By designing advanced composite stiffened-shell configurations, significant weight reduction may be possible compared with the sandwich and ribbed-shell structural concepts that have been studied previously. The study involves independent analysis of the advanced composite structural concepts that are presently being developed by The Boeing Company for pressurized HWB flight vehicles. High-fidelity parametric finite-element models of test coupons, panels, and multibay fuselage sections, were developed for conducting design studies and identifying critical areas of potential failure. Interim results are discussed to assess the overall weight/strength advantages.

Mukhopadhyay, Vivek↗

NASA NTRS2013

Hybrid Wing-Body Pressurized Fuselage and Bulkhead, Design and Optimization

The structural weight reduction of a pressurized Hybrid Wing-Body (HWB) fuselage is a serious challenge. Hence, research and development are presently being continued at NASA under the Environmentally Responsible Aviation (ERA) and Subsonic Fixed Wing (SFW) projects in collaboration with the Boeing Company, Huntington Beach and Air Force Research Laboratory (AFRL). In this paper, a structural analysis of the HWB fuselage and bulkhead panels is presented, with the objectives of design improvement and structural weight reduction. First, orthotropic plate theories for sizing, and equivalent plate analysis with appropriate simplification are considered. Then parametric finite-element analysis of a fuselage section and bulkhead are conducted using advanced stitched composite structural concepts, which are presently being developed at Boeing for pressurized HWB flight vehicles. With this advanced stiffened-shell design, structural weights are computed and compared to the thick sandwich, vaulted-ribbed-shell, and multi-bubble stiffened-shell structural concepts that had been studied previously. The analytical and numerical results are discussed to assess the overall weight/strength advantages.

Mukhopadhyay, Vivek↗

NASA NTRS2015

Development of the PRSEUS Multi-Bay Pressure Box for a Hybrid Wing Body Vehicle

NASA has created the Environmentally Responsible Aviation Project to explore and document the feasibility, benefits, and technical risk of advanced vehicle configurations and enabling technologies that will reduce the impact of aviation on the environment. A critical aspect of this pursuit is the development of a lighter, more robust airframe that will enable the introduction of unconventional aircraft configurations that have higher lift-to-drag ratios, reduced drag, and lower community noise. Although such novel configurations like the Hybrid Wing Body (HWB) offer better aerodynamic performance as compared to traditional tube-and-wing aircraft, their blended wing shapes also pose significant new design challenges. Developing an improved structural concept that is capable of meeting the structural weight fraction allocated for these non-circular pressurized cabins is the primary obstacle in implementing large lifting-body designs. To address this challenge, researchers at NASA and The Boeing Company are working together to advance new structural concepts like the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS), which is an integrally stiffened panel design that is stitched together and designed to maintain residual load-carrying capabilities under a variety of damage scenarios. The large-scale multi-bay fuselage test article described in this paper is the final specimen in a building-block test program that was conceived to demonstrate the feasibility of meeting the structural weight goals established for the HWB pressure cabin.

Jegley, Dawn C.↗

NASA NTRS2015

Preliminary Weight Savings Estimate for a Commercial Transport Wing Using Rod-Stiffened Stitched Composite Technology

A structural concept called pultruded rod stitched efficient unitized structure (PRSEUS) was developed by the Boeing Company to address the complex structural design aspects associated with a pressurized hybrid wing body (HWB) aircraft configuration. While PRSEUS was an enabling technology for the pressurized HWB structure, limited investigation of PRSEUS for other aircraft structures, such as circular fuselages and wings, has been done. Therefore, a study was undertaken to investigate the potential weight savings afforded by using the PRSEUS concept for a commercial transport wing. The study applied PRSEUS to the Advanced Subsonic Technology (AST) Program composite semi-span test article, which was sized using three load cases. The initial PRSEUS design was developed by matching cross-sectional stiffnesses for each stringer/skin combination within the wing covers, then the design was modified to ensure that the PRSEUS design satisfied the design criteria. It was found that the PRSEUS wing design exhibited weight savings over the blade-stiffened composite AST Program wing of nearly 9%, and a weight savings of 49% and 29% for the lower and upper covers, respectively, compared to an equivalent metallic wing.

Lovejoy, Andrew E.↗

NASA NTRS1984

Postbuckling behavior of graphite-epoxy panels

Structurally efficient fuselage panels are often designed to allow buckling to occur at applied loads below ultimate. Interest in applying graphite-epoxy materials to fuselage primary structure led to several studies of the post-buckling behavior of graphite-epoxy structural components. Studies of the postbuckling behavior of flat and curved, unstiffened and stiffened graphite-epoxy panels loaded in compression and shear were summarized. The response and failure characteristics of specimens studied experimentally were described, and analytical and experimental results were compared. The specimens tested in the studies described were fabricated from commercially available 0.005-inch-thick unidirectional graphite-fiber tapes preimpregnated with 350 F cure thermosetting epoxy resins.

Starnes, J. H., Jr.↗

NASA NTRS2022

Development of Advanced Manufacturing Approaches for Single-Piece Launch Vehicle Structures

Advanced, near-net shape manufacturing methods have the potential to enable production of structures with fewer welds and reduced machining requirements. Two such methods that are suitable for the manufacturing of single-piece, stiffened barrel-shaped structures are presented. The first solution deployed existing manufacturing technology to produce a thick-walled barrel with integrally machined stiffeners. In this study, an 8-ft. diameter barrel was produced and subjected to a buckling test. Data on the manufacturing and testing of this barrel are provided and compared against the traditional multi-piece weld construction approach. The single-piece barrel resulted in a 28% greater load carrying capacity than the welded barrel. A second solution utilized a novel flow-forming technique to produce the barrel and stiffeners in one process without the need for welding or machining. This is an emerging manufacturing method known as the Integrally Stiffened Cylinder (ISC) process. This innovative process is being evaluated for launch vehicle and commercial aircraft manufacturing. The one-piece, stiffened ISC barrels, which have been successfully fabricated at 10 ft. in diameter, offer a direct replacement for conventional multi-piece, welded or riveted structures. A cost-benefit analysis for launch vehicle cryogenic propellant tanks estimated that the ISC process offers up to a 50% reduction in manufacturing costs and a 10% reduction in mass.

integrally stiffened cylinder (ISC)↗

NASA NTRS1975

A fuselage/tank structure study for actively cooled hypersonic cruise vehicles: Structural analysis

The effects of fuselage cross-section (circular and elliptical) and structural arrangement (integral and nonintegral tanks) on the performance of actively cooled hypersonic cruise vehicles was evaluated. It was found that integrally machined stiffening of the tank walls, while providing the most weight-efficient use of materials, results in higher production costs. Fatigue and fracture mechanics appeared to have little effect on the weight of the three study aircraft. The need for thermal strain relief through insulation is discussed. Aircraft size and magnitude of the internal pressure are seen to be significant factors in tank design.

Baker, A. H.↗

NASA NTRS1986

Experimental study using nearfield acoustic holography of sound transmission through fuselage sidewall structures

The reduction of cabin noise in lightweight, propeller-driven aircraft is an especially difficult problem in noise control. Nearfield Acoustic Holography (NAH) was used to determine the mode of vibration and acoustic intensity for panels which differed in: construction (number of stiffening ribs, size of stifening ribs, construction material, and panel surface curvature); boundary support condition (free edge condition or clamped edge condition); and mode of excitation (structural-borne forces or airborne forces). The different samples of aircraft panels are described and the measurement of the natural response frequencies was discussed under various boundary support and excitation conditions. The results of the NAH measurements are presented.

Maynard, J. D.↗