A revolutionary approach to composite construction and flight management systems for small, general aviation airplanes

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A REVOLUTIONARY APPROACH TO COMPOSITE CONSTRUCTION AND FLIGHT MANAGEMENT SYSTEMS FOR SMALL, GENERAL AVIATION AIRPLANES University of Kansas Department of Aerospace Engineering Lawrence, Kansas Professor Jan Roskam Ed Wenninger, Teaching Assistant Abstract As a result, many mechanical fasteners are required , which drive up the weight and cost while also introducing The design studies for two composite general aviation delamination problems. Sad examples of the "state-of- airplanes are presented. The main consideration for both the-art" are: Beech Starship I, Boeing-Bell V-22, McDD of the designs was to avoid the typical "metal AV-8B, and the Boeing A-6 re-wing program, all of which replacement" philosophy that has hindered the outweigh aluminum equivalents.

widespread use of composites in general aviation aircraft.

The first design is for a low wing aircraft based on the The project for the Advanced Design Program at the Smith Aircraft Corporation GT-3 Global Trainer. The University of Kansas will be to develop methods in which second aircraft is a composite version of the Cessna 152.

conventional mechanical fasteners (bolts, rivets, screws, The project was conducted as a graduate level design etc.) can be eliminated in the construction of all- class under the auspices of the KU/NASA/USRA composite aircraft. These techniques will then be applied to two different aircraft. The two aircraft chosen were the Advanced Design Program in aeronautics. This paper will present the results obtained from the Fall semester of Smith Aircraft Corporation GT-3 Global Trainer and the Cessna 152. These two aircraft were chosen because 1991 and the Spring semester of 1992.

information was readily available to the design teams, and they represent what can be considered to be typical Nomenclature configurations for low and high wing aircraft. The class produced scaled production drawings and models that CRT Cathode Ray Tube show how the manufacturing process will work.

GPS Global Positioning System HUD Heads Up Display The second area of study was in the area of flight KU University of Kansas management and flight control systems. This subject was LCD Liquid Crystal Display investigated only during the Fall 1991 semester. Most RTM Resin Transfer Molding existing general aviation airplanes use mechanical flight USRA Universities Space Research Association controls. The handling qualities of these airplanes are often compromised by the friction and hinge moment feedback associated with such flight controls. In addition, Introduction many of these airplanes have undesirable Dutch roll and spiral mode characteristics. This increases pilot workload For the 1991-1992 academic year, the Advanced Design in conditions of turbulence and poor visibility. To remedy Program at the University of Kansas concentrated on two these problems, a de-coupled flight control system was main subjects. The first is in the area of composite investigated. Such a system has been shown to be very construction. The second is in the area of improving easy to fly. The results of the study included functional flight management and control systems.

diagrams and drawings describing such a system. In addition, a complete list of component weights, Most existing composite aircraft structures have been geometries, power consumption, and cost data was designed by using the "metal-replacement" philosophy. generated.

ProeeediA_ of the 8th Summer Co4¢re_¢ 426 NASA/USRA Ad_at_ed l)es4_tt Program system has proven easy to fly and is a promising solution Another problem with existing general aviation to increasing safety in general aviation.

airplanes is that pilots are required to be familiar with all navigation systems on board as well as with all FAA rules The system described above requires the use of a fly-by- with regard to air traffic control. This has made the wire flight control system. Two main considerations of current pilot environment extremely user-unfriendly. To such a system are the actuation method and the computer relieve these problems, a very user-friendly flight hardware that are required.

management system was developed. This system should be able to allow a low-time pilot to fly safely in the air Actuation Method traffic control system without the need for extensive training. This type of system was investigated in the 1990 For system redundancy and to allow for smaller, less academic year at the University of Kansas, and this study was a continuation of that work. powerful (and presumably less expensive) actuators, multiple servo tabs are used for each control surface. The selected values are as follows: Advanced Flight Management/Control Systems • Aileron 6 • Elevator 6 The purpose of this section is to present the main • Rudder 4 results from the advanced flight management and control study. This study was conducted only during the Fall 1991 The forces for each actuator were calculated and an semester.

extensive search was made to find a suitable actuator.

The Nash DL 1020 linear actuator was chosen. For parts Advanced Flight Control System commonality, the same actuator is used for all control surfaces. The installation of the actuator into an aileron The Advanced Primary Flight Control System (APFCS) is shown in Figure 1. The installation of the actuator is is a decoupled flight control system. Decoupled flight similar for the other control surfaces. 1 controls force the response of the airplane to be a function of only one input variable. This system is very different from conventional flight control systems which often require some combination of two or more pilot _ 9.1 5 ,.._ -25 DEG.

inputs to achieve a constant response. For example, to climb at a constant rate requires that the pilot pull back on the stick (or wheel) and add thrust through the throttle. To perform a steady level turn requires that the pilot pull the stick to the side to bank the airplane, pull back on the stick to maintain altitude, and add thrust through the throttle to maintain a constant airspeed. The purpose of the decoupled flight control system is to reduce pilot workload by eliminating the coupling of ___ HINGE LINE) control inputs necessary to produce steady-state responses from the airplane. The three motion variables that are controlled by the pilot through the APFCS are:

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• vertical speed • airspeed _FLEC=i'ION • heading rate _ 0.60 The APFCS couples the appropriate direct control signals and performs iterations until the response of the Fig. 1 Installation of aileron actuators airplane matches the signal input given by the pilot. This IN INCHES 42'7 ALL DIMENSIONS

200 --

ALE: I:60

AILERON TABS

150 -

100 -

Z '''' SIGNAL BUSES

J

50 o-

ELEVATOR TABS MOTION SENSORS

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Z d (D FLIGHT t-- CONTROL D COMPUTER nn SIGNAL CONVERTER/ AMPLIFIER ACTUATORS

I I I I. I I

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50 100 150 200 250

FUSELAGE STATION, F.S. (IN)

Fig. 2 Flight control system general layout (top view) 428 Proceedings of _e _ Su_r ¢oqfert_e NASA/USRA AdvamcedDesign Program Availability turned off would give sufficient accuracy for The other required equipment and the associated costs are given in Table 1. The installation of these systems is Category II landings.

shown in Figure 2.

To effectively inform the pilot, it was decided to use a Heads Up Display (HUD). This will allow the pilot to Table 1 Total system costs for the APFCS continually look outside the aircraft instead of having to monitor instruments inside the cockpit. This will give the pilot greater time to see and avoid other aircraft, thus Component Cost (US $1991) Actuators $ 3,200 increasing safety. An LCD HUD with a display size of 24 Rate transducers 18,255 x 6 inches was chosen because it is lighter and requires less power than a conventional CRT HUD.

Vertical gyroscopes 15,540 Computers 30,000 Batteries 316 To insure a safe airplane, designers conducted a failure Total $ 67,311 analysis to determine the minimum number of components required for redundancy. An acceptable failure rate was assumed to be 1 in 106 flight hours for non-flight-crucial systems and 1 in 109 for flight-crucial Advanced Flight Management System systems. The failure analysis was conducted for two different scenarios. The first was called the not-too- The main objective of the flight management study was to determine the feasibility of a very user-friendly system distant future system and the other was a more technologically demanding system. The main difference developed at the University of Kansas during the 1990-91 academic year. 2 The system is designed to allow an between the two systems is that the not-too-distant future inexperienced pilot to fly anywhere in virtually any system uses existing components and the futuristic system weather. To do this requires Category II landing uses much more integration. The listing of the required minimums. It was determined that GPS with Selective components for the not-too-distant future system is given in Table 2.

Table 2 Required components for the advanced flight management system Number needed Volume (in 3) Weight (Ibs) Power (watts) Retail price (91 Component $ us) for redundancy 103.6 272 8.5 23,572 Nav. computer/ memory/data base MFD 7.7 22,500 40.5 206 1558 TCAS II 127,533 2.74 4.2 192 6465 Airdata computer 23.1 83.0 962 30,000 Flight computer HUD 24.0 200 1200 16,000 GPS 1.6 5.0 2,610 5.0 88 FCI 14,905 40 120 TAS indicator 0.94 41 220 Altimeter 1.1 228.0 9,245 496,752 Totals including 1,199 backups From Table 2 it can be seen that this system requires a large amount of power, volume, weight, and cost.

Considering the nature of the airplane (a light general F Clamping Force aviation trainer), such a system is not feasible using [ I Infusion Pump existing technology. A reduction of the weight and cost by 50% was determined to be the upper bound of the advantage that can be obtained by using the futuristic system. This results in a system that will weigh on the order of 100 pounds and cost in the neighborhood of $250,000, still too expensive for a light trainer. However, Dry F'dx_rPreform such a system could be used in larger aircraft such as corporate or commercial transports.

Fig. 3 Fundamentals of Resin Transfer Molding Composite Structure Design The main advantage of RTM is that the resulting part has a controlled surface on all or most surfaces. This will The purpose of this section is to present the results of the composite construction and manufacturing study.

significantly reduce the amount of refitting required when This study was conducted during both the Fall and Spring all of the components of the airplane are joined together.

semesters. The main objective of the Fall semester was to Another advantage is that the materials are cheaper than find ways in which all mechanical fasteners could be conventional pre-preg materials. This is because the resin eliminated from the structure. For a representative is injected into the fibers by the partmaker instead of by aircraft, the Smith Aircraft Corporation GT-3 Global the company selling the service to the manufacturer. No Trainer was used. The main objective of the Spring freezers are required to store the materials, and the part semester was to try to incorporate these ideas into a is in near net shape after being released from the mold, design, and to compare the resulting structure with an further reducing costs.

aluminum design. The airplane chosen for the Spring semester was the Cessna 152.

The main disadvantage of this process is that twice the usual number of molds is required. This would make the Composite Manufacturing Technique process difficult for a start-up company to use due to the large initial capital investment. Finally, the technology is The importance of concurrent engineering has been not yet perfected. Despite these disadvantages, it was felt increasingly evident in recent years. This is even more the that the advantages far outweigh the disadvantages and case with composite structures. If the designer does not that in a few years the technology will be ideal for making consider manufacturing from the start, it is quite possible composite parts.

that the resulting product will be both overweight and over cost. For this reason, an extensive search of the Wing various manufacturing methods available was made. The method that seemed to have the most promise was Resin The GT-3 wing is designed to emphasize the elimination of mechanical fasteners. At the locations of Transfer Molding or RTM, a process in which dry fibers are placed in a double-sided mold. The resin is injected mechanical fasteners, the composite needs to be built up into the dry fiber at a constant rate so that all of the fibers because an interruption of the composite fibers weakens are exposed to the resin. The process is shown its structural integrity. This buildup around the fasteners schematically in Figure 3.

increases the weight of the composite, which is unacceptable. Another design driver in the wing design is ease of removal and replacement for the purposes of repairability and maintainability. The wing designs were conceptualized with these factors in mind: Proctedtn_ o/the _h Sum.wr Coqfere_¢

430 NASA/USRA Advanced Desigx Program

• repairability

• slide-on wing

• key-way joint

The assembly of the wing onto the stub consists primarily • conventional pin joint of sliding the wing on, attaching the landing gear, and applying the adhesive to hold the wing on. The stub will The slide-on wing concept will be used for the GT-3 not be symmetrical; thus, there should not be any trainer. The slide-on wing consists of a "stub" type fixture problems with mounting the wing upside down. The extending from the fuselage. The stub is integral to the actual application of the adhesive is to be investigated fuselage/carry-through structure. The stub is designed to further. Synergism is achieved when the stub is used both act as an inner layer of skin attached to the inboard for mounting the wing and for wing strength. The stub is portion of the wing. However, the wing will be assembled and then slid on this stub and attached with adhesive. an integral part of the structure of the inboard portion of This adhesive bond will then act as an interlaminar bond the wing. The landing gear mounting presents another advantage to using the stub. Because the stub extends to allowing the stub to act as a layer of skin. The stub will extend to buttock line 68 to allow for attachment of the the landing gear attachment, the stub can be used synergistically as part of the landing gear attachment.

fixed landing gear to the stub structure. The stub will be Some of the actual structural strength required for the shaped as the outer skin of the inboard wing to allow for a landing gear attachment and the inboard portion of it tight fit as the wing is slid over the stub. In the chordwise already exist in the stub.

direction, the stub will extend aft to approximately 0.70 chord where it will be rounded to an oval-type shape Some of the disadvantages of using the slide-on stub (Figure 4).

joint include: • difficulty of wing removal • tolerances Adhesives must be used to attach the wing because of the assumption that the stub will act as part of the wing skin. Thus the bond between the wing and the stub must be viewed as an interlaminar bond. This also assumes that the tolerances between the stub and the wing skin are very small (a similar metal joint requires approximately 0.0006-0.0012 inches). 6 This exact tolerance could present an accuracy problem during manufacturing.

Another concept that was developed was called the key- FIJSi[k AG_ sIl_ way joint. This joint allows the wing to slide on parallel to eL 17 "-"I_10 "7°_ the x-direction of the aircraft. This concept is shown in 81. 7s J . CNQRO Figure 5.

Fig. 4 Slide-on wing concept The advantages of using the slide-on stub joint to attach the wing include: • ease of assembly • joint/structure synergism • landing gear mounting synergism University of Kansas 431 EXTENSION OF WING SHEAR WEB MALE JOINT ATTACHES-7 TO THE WING / / _ _ HESE ARE ATTACHED JOINT MOVES IN THIS / DIRECTION / _,._. NO T E T H AT 0 T T T H H E E F M U E S T E A LLAGE FEMALE JOINT ATTACHES TO THE FUSELAGE BUSHING HAS A THICKNESS OF .062,5 IN Fig. 5 Key-way joint concept Fig. 6 Configuration of the pin joint The advantage to this joint is that it is not required to Structural Layout for the GT-3 Wing. A primary design take as much load as the slide-on wing does. This is goal of this design is to eliminate the use of mechanical because all of the bending loads are taken out by the fasteners. To accomplish this goal, the decision was made unique shape of the joint. Some adhesive will still be to develop a design that would distribute the loads and required to prevent the wing from sliding off. The main stresses more evenly throughout the wing as opposed to disadvantages are its very complex shape and, like the channeling each load into a specific structural member.

slide-on wing, the extremely narrow tolerances required The ultimate manifestation of this concept is the to prevent any movement. A model was built using monocoque wing. The pure monocoque wing, with no fiberglass and epoxy resin to gain further insight into the internal ribs, spars, or stiffeners, represents a limiting merits of the joint. During the course of many assemblies structure which designers can approach in an attempt to and disassemblies, the joint became worn and became obtain thin, hollow wings with low fabrication and more and more loose-fitting. Clearly this would not be assembly costs. Since the skin is the only structural allowable for an actual installation, so a remedy to this element, all loads on the wing will be distributed problem must be found.

throughout the skin. This concept is not feasible using conventional metal fabrication because of the high weight The final wing-to-body joint that was investigated was a that would be required to provide the necessary structural conventional pin joint. While the pin violated the stiffness. Even using high-modulus graphite composites, principle of no mechanical fasteners, it was required for the concept is impractical. For virtually any material, ribs the composite wing design for the Cessna 152. This is are required to hold the aerodynamic contour of the wing because the 152 uses a strutted high wing. By using a and to prevent the wing from flattening out, which would strut, Cessna was able to eliminate the bending moments result in structural instability. A rib is also required to at the root, and thus very little carry-through structure distribute the landing gear loads into the skin. Spanwise was required. To ensure that the bending moment stiffeners are desirable to reduce the panel width of the remained zero, it was necessary to use a conventional pin skin in compression, thus raising the buckling strength of the skin.

joint. The configurations are shown in Figure 6.

Proceedhcgsof the 8th Summer CoNfertmc¢ NASA/USRA Advanced De$1gmProgram Thestructural itemthatcanbeeliminated is thespar. Several fuselage construction concepts were investigated before deciding on a construction technique:

Thewebof a sparconcentrates theshear created by the

winglift intoa fewfinitepointsalongthechordof the

one-piece construction-wing and body

wing. The sparscan be eliminated alongwith the

concentrated loadsassociated with them,allowing the one-piece fuselage

two-piece fuselage

leading andtrailingedges ofthewingtoserve a structural

function. • front/back

• side/side • top/bottom

StructuralLayoutfor the 152 Wing. Due to the

configuration of the Cessna 152, a no-spar wing as A top/bottom concept was chosen for the construction previously discussed is not possible. This is due to the of the GT-3 and the composite 152 fuselage. It has large cutout required for the doors. There simply is not several advantages over the other ideas. A manufacturer enough room to distribute the loads. For this reason the can lay up the bottom half of the airplane at room composite wing for the 152 uses conventional shear webs temperature or in an autoclave and then install most or all placed at the same locations as the standard 152. These shear webs channel the forces into bulkheads in the of the systems without having to crawl inside the fuselage.

The idea is to put the bottom half on "sawhorses" and fuselage on both sides of the door. The composite wing have excellent access all around the fuselage, saving differs from the conventional wing in that the upper skin between the shear webs acts as the spar cap. Figure 7 is equipment installation man-hours. The top half can be set over the entire assembly to see if all the systems and an exploded view of the wing showing the shear webs and equipment fit inside. Then, the top can be lifted off and the required ribs.

installation can continue, or the two halves can be bonded together. The two-piece fuselage will have pieces that will be easier to manufacture and work with than a one-piece fuselage.

A complex curve or a stair-step may be required for the .n_ SUef*CE _, joint along thc aft end of the fuselage, which could increase the complexity of the manufacturing process.

Current examples of the top/bottom construction include: • Smith GT-3 Trainer • Wheeler Express f LO_m • Fitzgerald Cozair Wheeler actually purchased and built a Giasair before they designed the Express and decided against the left and right half concept. An additional benefit of this Fig. 7 Exploded view of composite 152 wing concept is that small, non-load bearing structures could be taped in to run the flight controls. Figure 8 shows how the top/bottom construction technique is implemented on Fuselage the composite 152.

The purpose of this section is to present the concept chosen for construction of the fuselage of the Smith GT-3 Global Trainer and the composite 152.

Un_ersO of nransas 433 some of the attachment mechanisms available. A residual clip joint is one in which one piece must "snap" into place.

That part can be removed by collapsing the joint with a special tool.

Hot-melt adhesive is suggested as the attachment I mechanism for the stub slide-on type joint. The wing is attached to the fuselage by sliding it onto a stub that is part of the fuselage. By using an adhesive that melts at a temperature below the cure temperature of the wing and fuselage, but above the maximum operating temperature of the airplane, the wing can be removed without damaging other airplane components. The stub is itself the carry-through structure.

Carry-through Structure. The design of the carry- through structure uncovered several problems with the design of the Smith GT-3. Currently, the Smith GT-3 uses two spars to carry the wing loads. The leading spar is located at 0.45 chord, and the trailing spar is located at 0.70 chord. The leading spar carries most of the load, and Fig. 8 Demonstration of top/bottom construction the trailing spar simply acts as a mount for the trailing edge devices. The wing leading spar runs through the cockpit directly below and behind the pilot's back. In the The empennage could be designed so that it fits inside case of a crash which broke the spar, the spar would drive grooves in the bottom half of the fuselage. Then, the top up through the back of the seat, severely damaging the half could fit over a section of the empennage, "locking" it pilot's spine. With the loads concentrated on one spar, in place. This is also shown in Figure 8. the likelihood of the spar's breaking is increased. This design was considered unacceptable.

The components of the wing-body joint are the wing joint and the fuselage carry-through structure. The wing- Since the pilot seat location and the aerodynamic shape body joint attaches the wing to the fuselage and also of the airplane were not items which the group was transfers the lifting loads from the wing to the fuselage allowed to alter, the carry-through structure must be structure. located behind 0.45 chord. Since the structure will still be located behind and beneath the pilot's back, the design The design criteria for the wing-body joint follow: driver for the carry-through structure was crashworthiness. The two main methods used to achieve • even distribution of loads this objective are: • no mechanical fasteners * distribution of the loads • secure attachment of wing to fuselage • repairability and replaceability • controlled failure design Joint Concepts. Several joint concepts were developed By distributing the loads over a larger area, the during the preliminary design phase of this task. The likelihood of the structure's breaking is diminished.

three most promising concepts were the stub slide-on Additionally, the carry-through structure was designed so joint, key-way joint, and one-piece wing. that, in the case of a crash, the wing would fail before the carry-through structure. Since the structure is designed in An attachment mechanism must be determined for each this manner, the fracture location is moved away from the of these joint types. Residual clips and adhesives are pilot and passengers.

Proceedings of Ore 80t Summer CoNference 434 NASA/USRA Advanced Design Program drilled in the composite. Thus, the full strength of the composite can be expected. A cross-section of this joint is The maximum loading placed on the carry-through structure occur in the one-wheel landing cased. The loads shown in Figure 9.

are: bending 834,000 in-lbs shear 22,400 lb torsion 113,000 in-lbs UIX_" F rm'rm Conceptual Design. A tube-type design was chosen for the carry-through structure. Use of a box or tube carry- Front Sl_u" Stud H OI'IZO_ _ TU through structure rather than a two-spar structure was shown to save weight while maintaining the required load- carrying capability. Additionally, the tube structure lends

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itself more easily to the use of the stub slide-on joint and the no-spar wing concepts.

The actual shape of the structure will follow the internal C_moe:l Frame Front V_w contour of the wing airfoil shape. The structure will be rounded at the leading and trailing edges. The shape will Fig. 9 Clamped joint concept be approximately an ellipse. Once the structure penetrates the fuselage skin, the leading edge will curve back from the leading edge to 0.45 chord to fit around the Landing Gear pilot seats. This cutout significantly reduces the torsional strength of the structure, but has little effect on the shear Composite landing gears have been used for many years or bending strengths. To recover some of the torsional in general aviation aircraft, so design of the gear legs is strength lost, stiffeners will be added along the 0.45 chord not all that difficult. However, one of the main problems location in the carry-through structure. Additional with landing gear design is finding a way which will stiffeners will be added for support of the wing structure, the fuel tanks and the control runs. In addition to the introduce the fairly large point loads into the structure.

stiffeners, the fuselage structure will add torsional For composite design, distributed loads are much easier to accommodate. One possible solution to this problem is stiffness to the carry-through structure.

the concept shown in Figure 10.

By over-designing the strength of the structure and In this concept, the cross at the top of the landing gear adding stiffeners to further improve the strength, the strut is designed to take out all the landing gear loads.

designers have chosen not to take full advantage of the This also would eliminate the need for a drag brace, weight savings possible over a conventionally designed reducing drag.

tube structure. However, the pilot's safcty in the event of a crash is greatly improvcd.

Empennage Though the empennage was discussed previously, the details of its attachment will be discussed in this section.

The empennage is clamped in place by both the upper and lower fuselage skins and by two bulkheads located where the front and rear spars of the horizontal tail intersect the fuselage. At these points a clamped joint is used. The clamps used for this joint do not require holes Fig. 11 Engine mount concept Conclusions By using novel techniques of composite construction, designers may avoid the problem of bolts and screws in

Fig.10 Landing gear attachment concept

composite structures. Table 3 makes a comparison of the structural weights for a composite and a conventional aluminum 152.

Engine Mount Table 3 Comparison of composite and conventional The engine mount for the composite 152 also posed the structural weights problem of how to introduce a point load into a composite structure. A "bathtub" type fitting was Component Composite Aluminum developed (Figure 11) at five different locations around Wing 206 216 the firewall. The engine mount uses the same metal Fuselage 138 231 structure that the standard 152 uses.

Empennage 28 31 Landing Gear 80 96 Table 3 shows that the differences in weight, with the exception of the fuselage, are generally not significant.

This indicates that, considering the assumptions required for preliminary design of the composite aircraft, there is no significant advantage from a weight standpoint to either material.

References 1o Jones and Winger. Advanced Primary Flight Control System for the Smith Aircraft Corporation GT-3.

Proceedings of the 801 Summer Coqference 436 NASA/USRA Advanced Design Program Lawrence, Kansas: The University of Kansas, December 1991.

, Clatterbuck, Newton, and Widup. Advanced Flight Management System Study for the Smith Aircraft Corporation GT-3. Lawrence, Kansas: The University of Kansas, December 1991.

° Hoffman, Rodkey, and Roper. Advanced Guidance and Display Study for the APT. Lawrence, Kansas: The University of Kansas, Fall 1990.

4o .Krause, Miller, Mouch, Schmitz, and Houdeshell.

Wing Design and Manufacturing Study for the Smith Aircraft GT-3. Lawrence, Kansas: The University of Kansas, December 1991.

5° Krause and Wenninger. Composite Wing Design for the Cessna 152. Lawrence, Kansas: The University of Kansas, April 1992.

, Hixson, Thacker, and Ellrott. Fuselage Design for the Smith Aircraft GT-3. Lawrence, Kansas: The University of Kansas, December 1991.

7.

Thacker, M. Composite Fuselage Design for the Cessna 152. Lawrencc, Kansas: The University of Kansas, April 1992.

° Jones, V. Composite Empennage Design for the Cessna 152. Lawrence, Kansas: The University of Kansas, April 1992.

9, Schmitz, K. Landing Gear Study for the Composite Version of the Cessna 152. Lawrence, Kansas: The University of Kansas, March 1992.

10. Headrick, J. Engine Attachment Design for a Composite Version of the Cessna 152. Lawrence, Kansas: The University of Kansas, March 1992.

11. Headrick and Schmitz. Manufacturing and Assembly for the Redesign of a Metal Cessna 152. Lawrence, Kansas: The University of Kansas, April 1992.