Viper cabin-fuselage structural design concept with engine installation and wing structural design

r Table of Contents Section

Pa e

2.0 Description of the Design .........................................

6.0 Weight Summary ............................................... 2O 7.0 Environmental Considerations ......................................

8.0 Conclusions ...................................................

Appendix A: Spatial Requirements Specification Document .................... A.1 Appendix B: V-N Diagram ........................................... B.I C.I Appendix C: Torsional Diagram .......................................

D.1 Appendix D: Su'uctural Decomposition ...................................

Appendix E: Drawings .............................................. E.1 r List of Figures and Tables 2.2.1 Standard Longeron Mounting Bracket ...............................

2.3.1.1 Spline Interface ..............................................

3.4.1.1 Wing Load Distribution ........................................

3.4.1.2 Wing Moment Distribution ......................................

4.4.1.1 Front Spar Section Modulus Requirements ...........................

4.4.1.2 Rear Spar Section Modulus Requirements ...........................

Tabl......._e 1.0 Summary of Critical Detail Parts ....................................

32.1.1 Yield Strengths of 2024-1"3 Aluminum .............................

3.2.1.2 Allowable Stresses for Longeron Sizing .............................

3.2.1.3 Allowable Loads for Longerons ..................................

II 4.2.1 Cabin Margins of Safety .........................................

4.3.1 Allowable Shear and Buckling Stresses ..............................

4.3.2 Drive Shaft Margins of Safety .....................................

4.4.2 Skin and Rib Margins of Safety ....................................

4.4.3 Carrythrough Margins of Safety ....................................

6.0 Subassembly Weights ............................................

r 1.0 Project Summary Table 1.0: Summary of Critical Detail Parts In response to the requirements of the "Cabin- Component M.S. Load Page Fuselage Structural Design Concept with Engine Center 0.05 bending 12 Installation" (421F93ADP01-2) Statement of Work Longeron (SOW) and its Addendum, "Wing Structural Design," the following report was preparecL This report 2" Longeron 0.05 bending 12 describes the process and considerations in designing the cabin, nose, drive shaft, and wing assemblies for I" Longeron 0.05 bending 12 the "Viper" concept aircraft. Interfaces of these Center 7.16 normal 11 assemblies, as well as interfaces with the sections of the aircraft aft of the cabin, are also discussed. The Longeron results of the design process are included.

Center 8.71 buckle 11 Longeron The goal of this project is to provide a structural design which complies with FAR 23 requirements MS20470DD-4 0.82 shear 12 regarding occupant safety, emergency landing loads, Floor skin 2.25 shr bkl 12 and maneuvering loads. The design must also address the interfaces of the various systems in the cabin, MS20470DD-8 5.0 shear 12 nose, and wing, including the drive shaft, venting, I vacuum, electrical, fuel, and control systems.

Floor Ribs 4.62 buckle 12 Interfaces between the cabin assembly and the wing carrythrough and empennage assemblies were Primary Shaft 0.81 buckle 13 required, as well. In the design of the wing Secondary 2.68 buckle 13 assemblies, consistency with the existing cabin design Shaft was required.

Primary Shaft 2.13 shear 13 The major areas considered in this report arc materials and construction, loading, maintenance, Secondary 2.13 shear 13 environmental considerations, wing assembly fatigue, Shaft and weight. The first three areas are developed Front Spar 0.231 bending 13 separately for the nose, cabin, drive shaft, and wing assemblies, while the last three are discussed for the Rear Spar 0.231 bending 13 entire design.

Wing Skin 0.10 torsion 14 For each assembly, loading calculations were Wing Rib 0.19 torsion 14 performed to determine the proper sizing of major load carrying components. Table 1.0 lists the Stringers 0.062 torsion 14 resulting margins of safety for these key components, along with the types of the loads involved, and the Front 0.698 bending 15 page number upon which they are discussed.

Can, ythrough Rear 1.81 bending 15 Carrythrough 2.0 Description of the Design 0.05 shear 16 Wing 2.1 Nose Assembly Attachments The nose assembly was designed to reduce the impact load on the cabin. This was accomplished by a load of 8g's. The second part of the nose assembly designing the nose into two sections, as may be seen crumples at 17g's. The load carrying members which in drawing F93-1C-128-2. The forward section determine the crumpling are the longcrons in the nose assembly. They were designed specifically to buckle accounts for 60% of the nose length and crumples at under the given loadings. will be satisfied in the floor assembly.

The forward nose longerons are mounted to the prop The four center longerons are designed to carry the entire bulkhead and a nose frame, which acts to split the forward impact load of the aircraft, without failing in loading into two stages. The aft nose longerons are compression or by buckling. The bending loads were mounted to the nose frame and nose bulkhead. The designed to be distributed between the center longerons longerons are mounted using a bracket on either end and the floor longerons. The layout of the longerons may of each longeron, which connects to the interface be seen in drawing F93-1C-129-2. The center longerons surface (i.e. a bulkhead or frame). are actually part of the drive shaft assembly and may be seen in detail in drawing F93-1C-150-2.

An access panel was included in the nose assembly, There are a total of six major load carrying longerons, all on the port side of the nose. The access panel runs the length of the nose, and is mounted with AN526C- mounted as part of the floor structure. The longerons 6-32 screws to the prop and nose bulkheads, as well running through the base of the door frame assembly are as the nose frame. The skin is mounted on the actually split into four selxn'ate longerons in order to remaining surface area of the nose. MS20470DD-4 interface with the door frames. The longeroos are all rivets are used to fasten the skin to the longerons, the mounted at their ends to their respective interface surfaces prop and nose bulkheads, and the nose frame. These (i.e. bulkheads or door frames) by mounting brackets that are shown in detail in Figure 2.2.1.

rivets are also used to mount half inch angles to the inside of the nose skin, parallel to the longerons.

These channels are to stiffen the nose skin against aerodynamic loads from the prop.

2.2 Cabin Assembly 2.2.1 Longeron, Cabin Side, and Roof Design Before doing the actual structural design of the cabin, the volume constraints imposed by FAR 23 were considered. A Spatial Requirements Specification Document that complies with governing regulations was prepared to address the volume conswaints. This document determined the minimum volume required in the cabin for the pilots and JAARS crew seats, as may be seen in Appendix A. The cabin structure was Figure 2.2.1: Standard Longeron Mounting Bracket then ready to be designed around these volume constraints.

The door frame assemblies have been sized accurately.

In approaching the cabin structural design, it was Actual detail design of these assemblies was conducted in order to obtain the sizes. However, these details are determined that the cabin would undergo normal, beyond the requirements of the SOW, therefore, they are bending, buckling, and shear loads. These loads are due to maneuvering and emergency landing not included in this report. The resulting dimensions of conditions, as discussed further in section 3.0.1. The this design process are applied to the cabin swucmre fwst considerations made were for normal, bending, volume conslraints and layouL and buckling loads which were designed to be taken in the longerons of the cabin su'ueture. These Two-inch longerons, similar to the floor longerons are considerations were further split into examining the mounted at about the middle of the height of the cabin.

center structure longerons of the drive shaft support These are to help carry the overall loadings on the cabin, but do not carry any major loads. The same is true for box and the longerons run in the floor structure. By designing the longerons to handle the previously the half inch angles mounted between these upper two- mentioned loads, the requirements of the SOW inch longerons and the floor longerons (refer to drawing regarding occupant safety demands of FAR 23 are F93-1C-129). They are riveted to the skin to increase the met in part. The remaining part of the requirements stiffness, but are considered to carry no major load.

r Additional one-inch longerons are mounted in the may be seen in the floor cross section view in drawing F93-IC-127-2.

bottom of the floor structure to aid in carrying the bending loads on the structure. The two-inch longerons are fastened with brackets s'unilar to the The function of the floor and seat suptxm ribs is mainly to absorb the downward load of the JAARS seats on two-inch floor longerons, but the half inch angles and one-inch longerons require no bracket mounts. impact. They also aid in providing bending and torsion support, but are not designed to carry these loads directly The nose and aft firewall are supported by angles as major load paths. The floor ribs are the ribs running which run around the fuselage cross section, perpendicular to the longitudinal axis of the aircraft. They sandwiching the bulkhead and firewall between the are designed to carry a full lgg downward load from the two rings of these angles. This may be seen in seat and pilot, as the worst possible loading case. The drawing F93-I-129-2. The forward bulkhead was seat support ribs, which run parallel to the aircraft designed solely to act as a close out between the longitudinal axis, are designed for the same load. The cabin area and the control systems and nose gear longerons mounted to the bottom of the floor structure are mountings. This is to keep loose items in the cabin to run through cutouts in the floor ribs. In order to accommodate the floor skin thickness, MS20470DD-8 area (such as pencils, etc.) from rolling into the rivets are used in the ribs to mount the ribs to the floor control system and nose gear area. This bulkhead is skin. However, MS20470DD-4 rivets are used to mount mounted to its angle frame with AN526C-6-32 screws to allow removal of the bulkhead for access to the the ribs to the bottom of the floor structure, which is the 0.03 inch thick fuselage skin.

control systems and nose gear. Cutouts are made in the bulkhead to allow the longerons to pass through, The remaining floor components are the one and two-inch uninterrupted.

longerons. The longerons were discussed in detail in The roof structure and the aft door frames interface section 2.2.1 and are riveted to the bottom of the floor with the aft firewall and angle as shown in drawing structure with MS20470DD-4 rivets, as seen in drawing F93-1C-129-2. The roof structure consists of three F93-1C-127-2.

channels mounted together with the typical longeron bracket mounts. This structure is also mounted to the top of the two door frame assemblies.

2.3 Drive ShaftSystem The other major components of the cabin assembly 2.3.1 Design and Crashworthiness consist of the instrument panel, cabin light, door, The unique design of this aircraft is the mounting of the windshield, JAARS seat, drive shaft, and floor engine on top of the wing box. This called for an assemblies. The floor assembly will be explored in extensive design of a drive system to span the 108 inches further detail in the next section. The drive shaft separating the engine and the propeller. The design assembly will be discussed in detail in section 2.3. inherently required significant considerations to the size, weight, vibration, occupant safety, manufacturability, and maintenance requirements of the shaft.

2.2.2 Floor Assembly The floor assembly was designed to withstand the loads imposed by the pilots and their seats on impact, The la'imary function of the drive system is to transmit the in addition to withstanding the loads imposed on the torque produced by the engine efficiently and safely. One aircraft while maneuvering. All torque in the aircraft of the first considerations was safety during the event of is designed to be carded by the floor structure. an emergency landing where in either the shaft or engine would shear from its attachments and strike the occupants.

It was decided to place a support system directly below In approaching the layout of the floor structure, considerations were first made for the shear flow the shaft as a primary load path and to ensure sufficient throughout the cross section. The floor skin and support for the shaft in the event of a crash. The four members which assemble to form the central structure longerous were designed to take the torque loads, without buckling from the resulting shear flow. The were sized to carry bending loads through the cabin and floor skin was designed to be 0.06 inches thick, with also to provide substantial support for the 18g loads two minor angles spaced 4.6 inches apart and experienced by the shaft during emergency landings.

mounted on the floor skin between the floor ribs. This These members also define the outside edges of the drive shaft housing that tuns through the cabin area.

The housing itself provides the area to run necessary equipment from the engine to the cabin compartment.

The control systems for the elevator and rudder run along the floor directly underneath the drive shaft.

The electrical and vacuum tubes are attached to the structural member supporting the shaft. Venting for cabin heat also runs through this channel. A plastic skin is all that is required to cover the sides and top of the channel. This skin is held on by quick release fasteners so as to provide easy access to the internal parts. Allocation of space for each subsystem is Figure 2.3.1.1: Spline Interface denoted in drawing F93-1C-152-2.

The entire drive shaft assembly includes a primary and secondary shaft, three bearings, and two universal longeron structure under the drive shaft. The support joints. Bearings are located at stations 1, 37, and 103.

would then prevent the engine from rotating about the Universal joints reduce vibration and forces due to universal, thus completing assurance that there is no misalignment and are located at stations 5 and 108. possible way for the engine to enter the cabin area in crashes.

The universal connections also allow the drive system to operate free of bending loads resultant from aerodynamic, inertia or engine loading. The relation This support can not be designed at this point, as research of each universal joint with the spline gear into the mounting points on the 0-235 engine resulted in connections of the primary and secondary shaft along no existing engine mounts at the required points.

with that of the union of the engine mounting shaft However, there are bolt locations in the required area can be seen in drawing F93-1C-151-2.

which may be picked up in order to mount the support.

Coordination of efforts with Lycoming would allow the The spline connection at the rear bulkhead allows for mount to be made in this required location. Once the thermal expansion and easy removal of the drive shaft support location is confirmed from this process, it will be and may be seen in Figure 2.3.1.1. The Second spline possible to size the support accordingly. This is system is located in the nose cone just aft of the recommended for futme completion of the Viper aircraft forward bulkhead. This spline system incorporates a overall design.

shear pin connection for emergency landing safety.

The Shear pin is specially designed to fail at 8g's, 2.3.2 Bearings allowing the secondary shaft to collapsewithin the Bearing location and type plays an important function in primary shaft and therefore eliminating the chance of the drive shaft design. The first bearing is located at the either shaft shearing apart and injuring an occupant. forward bulkhead. The bearing is a thrust bearing This design is similar to that of a collapsible steering designed to hold the thrust of the propeller and distribute the load into the structure of the aircraft. The second system used in ears today. The collapsing nose assembly, in addition, absorbs the shock of impact bearing is locatedjust aft of the instrument panel and and therefore further protects the occupants. Once the inside the cabin area. The beating's location allows nose assembly has completely crumpled, the shaft acts support for the weight of the shaft. The bearing itself is as a support to hold the engine aft of the cabin me& mounted onto the structural member running in the middle of the cabin area. This bearing is also responsible for From this design, the drive shaft holds a thre_-fold holding the shaft in place during an emergency landing.

crash worthiness agenda. Therefore, the only The third bearing is located at the fuewall and holds the possibility for the engine to enter the cabin area is for end of the drive shaft. This bearing is to take out engine the engine to rotate around the universal, up into the vibration. The bearing is mounted to stiffeners running cabin area. This is to be prevented by mounting an from longerons along the faewall. These same stiffeners hold the structural member.

engine support to the front, bottom of the engine.

This support would then be connected to the center Bearings areallsealed to keep out dust and dirt. The and can be seen in Drawing F93-1C-160-2.

bearings are also all greased bearings. An oil system was considered, but was found to be costly and hard The ribs are designed in two parts. The leading edge to maintain. The grease system is adequate for the portion is riveted to the front spar and the leading edge engine rpm expected. The bearings will have to be skin panels are riveted to this portion of the rib. The wing greased as part of the I00 hour maintenance check.

box portion of the rib is riveted to the front and rear spars.

The ribs are designed to allow the stringers to pass To provide access to the bearings, access panels in through uninterrupted. Therefore, cut outs are made at the the drive shaft assembly housing are provided at the appropriate positions in the ribs (refer to Drawing F93-1C- firewali and at the base of the instrument panel.

I60-2).

Access to the forward bearing can be made through the skin of the nose cone.

.2..4.3 .Carry_ough The carrythrough Slzucnge was designed to carry the lift and drag loads on the wing for worst ease load conditions.

2.4 Wing Assembly The carrythrough assembly was also designed to include interfaces between the cabin floor slrucan-e and the 2.4.1 Spar Desisn empennage strucaa-e.

The front spar is located at the quarter chord, carrying a maximum of 91% of the moment due to lift. The In designing the carrythrough structta-es, the size rear spar lies on the 70% chord and is responsible for coasa'aints due to the size of the wing spars at the root up to 21% of the total moment.

were imperative to consider. The location of the can'ythrough structures relative to the firewall and rear Both the front and rear spar are 2024-T3 aluminum bulkhead were also of concern, as the manufacturing alloy sheets, which have been blank pressed and process at these locations determined the shape of the brake formed to their respective shapes. The principal structures. Another important factor to be considered was form of each cross section is a "c" channel which fatigue performance. Upon consideration of all of these faces to the interior of the wing.

factors, the carrythrough su'uctm_ cross sections depicted in drawing F93-1C-170-2 were designed.

On the front spar, the flange width initiates at 2 inches to accommodate the fuselage interface.

The front carrythrough structure, consisting of the front Between stations 43 and 55 this width is reduced to plate and front channel, is mounted to the firewall. The 1 inch, which provides sufficient strength. The design rear carrythrough struco.we, consisting of the rear plate and of the rear spar is slightly different with the flange rear channel, is mounted to the rear bulkhead. This may starting at 3 inches and reducing to 1 inch at station be seen in drawing P93-1C- 170-2 and will be discussed in detail in section 5.4.3. The cross sections of the structures 90. Lightening holes have been employed to reduce weight where the design section modulus exceeded allow for mounting of the wing attachment fittings within the front and rear channels.

the requirements.

2.4.2 Wing Skin and Ribs Two-inch longerons are run between the front and rear The wing skin, ribs, and stringers were designed to carrythroughs at the same height as the floor longerons in withstand the torsional loads applied during flight and the cabin structure. The two upper outboard longerons are landing conditions.

mounted to the carrythrough channels with standard brackets. The two lower inboard longerons are mounted In approaching the layout of the rib spacing, the to the flanges of the canythrough channels with standard resulting shear flow is distributed through the skin brackets. This may be seen in drawing F93-1C-170-2 and will be discussed in detail in secdon 5.4.3. One-inch panels, ribs, and sa'ingers. The skin panel and rib thicknesses vary across the span from 0.02 inches at longerons are also placed in line with the floor ribs to the tip to 0.032 inches across the center to 0.071 continue those load paths from the fuselage to the inches thick over the fuel tank. Three stringers are to empennage.

be spaced at 36.7%, 48.4%, and 60.2% of the mean aerodynamic chord to reduce the shear flow in the Due to the design of the cabin slructure, the cross section skin panels. Seven ribs are spaced across the span of the Viper was increased, thus also requiring an increase _=- in the diameter of the empennage. Therefore, the carrythroughs, it was necessary to determine the maximum empennage will need to be redesigned to account for percentages of lift experienced by each component. This this. During this redesign, it is planned that the was done for the front spar by finding the x_ position for" longerons in the empennage may be placed such that the maneuver condition. This was then divided by the distance between the front spar and rear spar to give the they coincide with the placement of the longerons in percentage of lift carried there. The result was a the carrythrough assembly. This would complete the maximum of 91% of the lift required to be carried by the load path through all longerons throughout the aircraft.

front spar. The process was done similarly for the rear spar, but the x_ for the dive condition was used, as this 2.4.4 Wing Attachments condition produces the worst case load for the rear spar.

The result was a maximum of 21% of the lift to be carried The wing fitting attachments consist of the conventual quadruple shear lug and the double shear lug. The by the rear spar.

front spar attachments consist of a top and bottom 3.1 Nose Assembly quadruple shear lug on both sides of the aircraft. The front spar lugs taper into c-channels which are riveted Since the nose assembly was designed to fail in two parts, into the carry through strncture and the front spar.

the design of the longerons were considered in two parts.

The c-channel slides between the carry through A forward load factor of 18 was used in the c_culations, structure and the firewall. The other side of the lug is fitted inside the c-channel spar. The rear spar as was discussed in section 3.0.1. The longerons were then designed to buckle at 8g's and 17g's by determining attachment consist of a single double shear lug tapering into an I-beam, which again is riveted to the the required moment of inertia of the longerons, as will be discussed in section 4.1. The forward load is assumed to carrythrough and spar. Again the I-beam is fitted inside the carry through structure and is fitted inside be carried through the longerons only, as the skin will the rear spar. The front spar lugs have 2024 buckle before longeron failure.

aluminum NAS1314 bolts and the rear spar has 2024 aluminum NAS1310 bolts.

3.2 Cabin Assembly 2.5 Structural Decomposition 3.2.1 Longeron, Cabin Side, and Roof Design Appendix D iUuswates the total strucaual In first approaching the cabin longeron design, the yield decomposition for the cabin-fuselage and wing strengths in tension, compression, and shear were determined for 2024-1"3 aluminum. They appear in Table assembly structural design concepts. This overall 3.2.1.1. These were used, assuming a margin of safety concept combines the separate assemblies which are discussed in sections 2.1 through 2.4. (MS) of 0.05, to determine maximum allowable stresses for each failure mode, as shown in Table 3.2.1.2. The general equation used was MS = (F_)/(f,,.,,) - 1.

3.0 Loads and Loading Table 3.2.1.1: Yield Strengths of 202.4-T3 Aluminum 3.0.1 Loading Constraints In considering the required loading constraints, FAR Fy.t 61 ksi part 23 was consulted. The worst case loads were Fy.c 41 ksi determined to be 4.4g's up, due to maneuvering; 2.2g's down, due to maneuvering; 18g's forward, due F. 32 ksi to emergency landing; and 4.5g's sideways, due to i i emergency landing. These loads were determined by These stresses were used to size the cabin structure examining the flight and crash load criteria in FAR A23 and FAR 23.561(b)(3), which were found to be iongerons through determining the maximum loads permitted on each longeron. The maximum allowable the highest required loadings. The flight loads caused loads determined for the longerons are discussed in detail by gusts (n3 and n4) were checked using FAR A23 and were determined to be no greater than the in sections 3.2.1.1 through 3.2.1.3 and are listed in table 3.2.1.3.

maneuvering limit load factors (nl and n2).

Before designing the front and rear spars and (which is the worst case moment arm from the cg to the Table 3.2.1.2: Allowable Stresses for Longeron Sizing 2024-T3 Al & MS=.05) applied load).

The total bending load was then designed to be distributed f_o,,.,_oo 57.95 ksi between the center longeron and the floor longerons. The f,_.,,,._ 38.95 ksi allowable load on each center longeron was first determined. This was done by finding the allowable f,a.,,_._ 30.4 ksi moment, M.,,_ _, on a center longeron from f.,.,,, = (M.,,,, l,_)(c)/fl), using c and I properties calculated for one center longeron's cross section. A corresponding Table 3.2.1.3: Allowable Loads for Longerons load, P_t**r of 579.6 pounds was found by dividing this P,o_ 28,800 Ib center struct moment by the length of the longeron. This load is the greatest load that may be applied to a center iongeron P,_ t._t 579.6 lb bending without failure. Next, P.,,_ _,q was divided by Pm,_ to detm-mine the percentage of the total bending load that .P2-]._ 131.4 lb bending one center longeron can carry. This corresponds to 16.9% 20.5 lb of the total load. Therefore, the four center longerons bending P1- t**_ carry a total of 67.6% of the total load.

3.2.1.1 Normal Loads The remaining bending load was designed to be carried by In approaching the loading for the longeron structures, six two-inch and fo_ one-inch longerons. The loads carried by these longerons were found iteratively using the three separate areas were considered. They are normal, bending, and buckling. The first load types Mathematica (Wolfram, 1988) computer program. The considered were normal loading and buckling on the two-inch longeron loads were found first This was cabin slructure. All of the normal and buckling loads accomplished by programming moment of inertia were designed to be carried by the center longerons, equations with corresponding variable cross sectional due to the large size of the center structure. The dimensions into Mathematica. This generated "c" and "I" normal load, P,_,_, was calculated to be 28,800 values for inputted cross sectional dimensions. These "c" pounds, due to an lgg impact load, n_d, using P_ = and "I" values were then imported into the second part of the program which solves the equation f,n.,, = (M2.

(nf,,u)(W_,, ,_ ._). This normal load was used to determine the compression and buckling loads on the i,q)(c)/(I)for M2-_**r The Mr i_ value was then divided by center longerons. the length of the longeron to determine the maximum allowable load on _e longerun, P:. t**c- Various cross 3.2.1.2 Bending Loads sectional dimensions were inputted into the program until The next load considered was for the bending induced an optimal trade between load capacity and structural on the cabin structure by maneuvering loads. The volume constraints was determined. The resulting value maneuvering loads act at the eg of the aircraft, which was P2-1,_ -- 131.4 pounds, corresponding to 3.8% of the is located at the firewall. Each maneuvering load total bending load. Therefore, six of these iongerons carry then results in a reaction load from the weight of the a total of 22.8% of the total bending load.

forward fuselage acting in the opposite direction of the maneuvering load. The worst case loads are due The same process used for the two-inch longerons was to an upward acceleration of 4.4g's and a sideward repeated for a smaller cross section and the one-inch acceleration of 4.5g's. The structure was designed longeron cross section was determined. The one-inch entirely for the very worst ease load of 4,5g's. The longeron was determined to carry 20.5 pounds (P_. _), resulting in a capacity of 3.5% of the total bending load resultant bending load, P_,_, was determined to be 3,429 pounds, by multiplying the weight of the for four of these longeroas. The remaining 6% of the forward fuselage (762 pounds, including all bending load is assumed to be carried by the floor, door frame, and roof slrucmres.

components forward of the aircraft cg and the JAARS seats and pilots) by the 4.5 load factor. The moment 3.2.1.3 Additional Load Paths and Interfaces carried by the total bending structure was found to be In addition to carrying part of the bending load, the two- 277,749 inch-pounds. This was found from M_,a_ = (Pb,,_)(L), where L is the length of the longerons inch longerons act as load paths to the nose and T,,,, z, of 51,360 inch-pounds.

engine/wing box stations. The nose longerons are fastened through the nose bulkhead to the upper two- The inertial torque was next found. The first step in doing inch longerons and the two-inch floor longerons. this was to find the resultant cg position for the seats, These same two-inch longerons are then fastened pilots, and instrument panel. This eg position was referenced to the center of the floor cross sectional area.

through the firewall to interface with the structure in the engine/wing box section of the aircraft. This This arm was then multiplied by the combined weight of allows for a continuous flow of bending and normal the seats, pilots, and instrument panel to find the resultant loads throughout the fuselage. inertial torque, T_, of 10,105 inch-pounds.

The loads on the interface fasteners were determined The wing and inertial torques act opposite of one another, assuming a worst case scenario in which the weight therefore the higher torque is chosen for shear flow of the aircraft components aft of the cg acts in shear determination. The shear flow, _**o resulting from the on the fasteners. The weight (840 pounds) is higher wing torque was found by solving T,,_ = 2(An_,.

distributed over the 16 interface fasteners. Therefore, ,_.)(om,,, ). The result was qn-- = 120 pounds per inch.

This shear flow will be used to determine the maximum the load on each fastener, Pma,,, was determined by multiplying the aft components' weight times the 4.5 allowable shear buckling stress in the floor cross section.

load factor and dividing by 16. The resulting value It was also used to find the fastener bearing load, resulting from the shear flow, in the floor fasteners. This was done was 236 pounds for P,.,_,**,.

by multiplying q_**, by the floor rivet spacing of two The center longeron structure is connected with the inches, resulting in Pb,. _**, t,_._ = 1,440 pounds.

two-inch floor longerons by channels, as seen in drawing F93-1C-150-2. These channels allow the 3.2.2.2 Buckling Loads The last floor loading consideration was buckling due to center structure to be connected to the longeron structure, thus creating a load interface between the the downward load of the JAAR_ seat and pilot. This two structures. Another load path is created at the was calculated assuming a worst case scenario of a full door frame assembly interface with the two inch floor 18g acceleration down on the floor and seat support ribs.

longerons and at the roof assembly. In this manner, The smallest length rib was also chosen to provide the the door frames and roof are integrated into the total highest loading criterion. The load was found by flow of loads throughout the cabin structure. multiplying the weight of the seat and pilot by the I8g load factor, resulting in Pa,_, _._ = 3,420 pounds. The 3.2.2 Floor Assembly load must be resisted in buckling by the floor and seat As previously mentioned, the floor structure was support ribs.

designed for two main purposes, torque carrying capacity and downward loading of the JAARS seats.

The loads are discussed in detail in sections 3.2.2.1 3.3 Drive Shaft Assembly and 3.2.2.2.

The primary load consideration for the drive shaft is in the 3.2.2.1 Torque transmission of power through the eight foot span Torque is applied to the fuselage in two possible sepm-ating the engine and propeller. The drive shaft itself ways. Torque is created by the difference in lift on is not responsible, nor expected, to carry any loads resultant from aerodynamic or inertial forces. Such loads the wings when turning and by the inertial side load on items of mass in the cabin. The wing torque was will be expected to be transmitted by the cabin structure the first to be determined. This was done by and the central beam assembly supporting the shaft by the assuming a difference in lift of 30% of the total lift bearings. Sizing of the primary and secondary shaft is, (70% lift on one wing versus 100% on the other, therefore, dependent on the critical strength needed to when banking) acting at the aerodynamic center of the resist shear or bueiding under maximum loading.

wing. The aerodynamic center was assumed to be at 50% of each wing, as a worst case position. This Limits set forth by FAR 23 require the drive shaft to distance from the centerline of the aircraft was used support loads without detrimental or permanent deformation under the torque produced by the maximum as the moment arm for the lift. It was multiplied by rated takeoff power and related prop speed. Noting that 30% of the total lift to determine the resultant torque, at- bearing, thermal expansion forces are placed on the frontal both shear and buckling restrictions are dependent on the maximum torque carded by the drive shaft and thrust bearing and into the nose structure.

because of the importance of pilot safety in the unique design of the Viper, the frequency of the The final sizing consideration is in the ability for the shaft propeller was "taken at a value 400 rpm slower than to remain centered within the housing from the point of the spline gear union aft. Motion in both lateral and anticipated along with the use of a factor of safety of 1.5. Through the direct relationship between power, longitudinal directions is limited to the shear strength of torque and angular velocity, the maximum torque on the central bearing mounting bolts. Sizing of the bolts to the drive shaft was computed as Torque,_ = meet the 18g crash worthiness requirement in the forward direction ensures substantial strength to resist travel from (Power,_)(FS)/(Angular Velocity._). The Avco- Lycoming 0-235 is certified at a max power output of side-to-side. Eight bolts, as seen in drawing F93-IC-151-2, 125 hp. Coupled with an expected 700 rpm frequency are responsible for carrying the 28,800 pound force in at initialization of take off, the maximum torque shear experienced in emergency deceleration. Assuming placed on the shaft assembly is 1406 foot pounds. equal distribution among bolts indicates a selection of AN5C4 bolts. Sizing of the bolts at 5/16 of an inch Iteration with size and stress levels indicated the provides a margin of safety of: MS,,ow,_ _,. = 0.59.

required wall thickness of the primary arid secolldary shaft to be 0.0625 inches. Methods provided by NACA TN 3783 (Gerard and Becker) indicated the 3.4 Wing Loads cridca/shear strength for buck/ing to be weft below that required for shear fracturing to occur. Prediction 3.4.1 Spar Design of the critical buckling stress under pure torsional The wing loading was calculated by assuming two simple load was made as follows: pressure distributions: a trapezoidal and an elliptical. The T_,,_.a=(I_*Pi_'E/(12*(1-.3):)*(t/L) 2 where K_ is a average of the two curves was then taken to obtain the factor based on the parameter _ as denoted in the final loading seen in Figure 3.4.1.1. The moment as a technical note and its figure 26. Shear stress in the function of buttock line location was then calculated using shaft was calculated as Shear stress = the estimated loading distribution. Figure 3.4.1.2 illustrates the total moment distribution over the half span (Torque_)(Radius**_)/(Polar moment of Inertia).

of the wing. Location of the front and rear spars on the Transmission of torque between the primary and wing planform established the percent of total moment secondary shaft is concentrated in the spline gear distribution carried by each structure.

union. The shear pin which is designed for failure at 8g's, ten times the expected load encountered under L_DING [Iblll I engine torque, is not used for this function. The primary purpose is to restrict crumpling of the nose assembly until the design impact level has been reached.

Secondary load considerations of the drive shaft

\

include forces produced due to vibration and

misalignment of the shaft. In order to hold such loads I

to a minimum, the drive shaft design has incorporated

°".T, , , , , , , . i : : : : : :1

a support system of three aluminum housed steel °,,.............,.,,.°,,...,,.,.,,.o,.

BUTTOCK LINE LOCATION bearing in a combination with two universal joints at each extreme of the shaft assembly.

Figure 3.4.1.1: Wing Load Distribution A variation of the operating temperature of the shaft in the range of 250 F results in a thermal expansion The front spar location at 0.25c requires that it carry a of approximately 0.25 inches over the 100 inch span.

Aft of the mid-bearing, this expansion is to be maximum of 91 percent of the total load on the wing. The absorbed by the rod extending into the aft spline gear rear spar, located at 0.70c, is required to carry at most union. In the section between the nose and mid- only 21 percent. The maximum percentage of the moment r handled by the front spar was found to be at B. the loads applied to the wing from the flaps and maneuvering speed at a high angle of attack. The ailerons are accounted for by the difference in C,. C= for limiting design criteria for the rear spar was found to the clean wing, using 652-415 airfoil, was determined in be at dive speed and low angle of attack.

the preliminary design report to be -0.0473 and C,_,p is - 0.3 using a conservative 60 ° split flap.

_nmt O'_na mt_} "10 To account for skin buckling during landing, the force of SO one main wheel upon touchdown is calculated by F=knW, where k is a correction factor, n is the load factor, and W is the aircraft gross weight. From FAR 23, I(--0.25 and n=2.67 and the force computed was 1167.5 pounds. This 3O force multiplied by the distance from the landing gear attachment point to the ground caused a counter-clockwise torque (as seen from the tip) of 32689 inch pounds. This torque is distributed through the rib to the spars and the inboard to the fuselage. The design of the skin and ribs 12 24 31; 48 §rl 72 84 g6 108 1211 132 444 156 t88 IilO lq2 BUTTOCK LINE LOCATION is determined by the curve that represents the maximum local torque.

.Figure 3.4.1.2: Wing Moment Distribution 3.4.3 Carrythrough As mentioned in section 2.4.3, the carrythrough structures Sizing of the spar cross section was determined by the were designed to carry the wing loads. The moment due secdon modules requirements needed to achieve an to lift acting on the wing at a distance of half of the semi endurance limit on cyclic loading of 107 cycles. The span from the center of the wing, was first determined.

S-n curves for 2024-T3 aluminum alloy placed the This was done at the worst case of 4.4g's, as required by maximum allowable stress at 27.5 ksi. The required FAR 23 for utility category aircraft, resulting in a moment section modules was therefore calculated as follows: of 365,541 inch-pounds. The percentages discussed in So-- OvD(maximum cyclic kf.)/f,,,.,s_ section 3.0.1 were then applied to the front and rear where f,ao,,,,,= F/1.01, providing a margin of safety of carrythroughs to determine the worst case load 0.01. The maximum cyclic load factor was taken as experienced by each. These loads were then to be used to 2.2.

determine the required section moduli of the carrythroughs to withstand the loads, as will be _ discussed in A portion of skin, averaging 3 inches in width, was section 4.4.3. A similar proc,._ was conducted for drag.

added to the cross section of the spar to obtain the The moment due to drag was found to be 21,470 inch- effective moment of inertia in bending. Lightening pounds, corresponding to a drag of 452 pounds, calculated at the dive condition.

holes employed in the spar structure allowed reduction of the overall weight while maintaining approximately 90% of the section modules. Though The only other load to be considered in the carrythrough sufficient room was allotted for stress concentrations, structure was the shear load due to lift, which must be further analysis may need to be done on this area. carried by the fasteners. This was found to be 3,848 pounds in each half of the carrythrough structure. The 3.4.2 Wing Skin and Ribs landing loads were not considered in any part of the In approaching the loading for the skin and fibs, design, as these loads are less than the 4.4g loads torsion was determined to be the sizing criteria for the experienced in flight. The snow load condition was also skin panels, fibs, and stringers. A spanwise torsional checked, as will be discussed in section 7.1.5, and found graph (shown in Appendix C) was constructed to be insignificant in comparison to flight loads, as well.

according to FAR 23A. The flight loads at dive, maneuver, flap speed, and flap speed with ailerons 3.4.4 Wing Attachments There were several conditions looked at in order to size deflected are determined by the equation T=qC=ct,,S.

the lugs for the front and rear spar attachments. One of C= and the dynamic pressure, q, are determined according to the above configurations and the speeds the loads considered is the landing loads. Two types of landing conditions were considered the direct moment of are taken from the v-n Diagram shown in Appendix r the load to bend the wing up and the torsional load of cabin components. The material selected for the longerons twisting the wing off. The other loads considered was 2024-1"3 aluminum. Sections 4.2.1.1 through 4.2.1.4 wear flight loads. The conditions how high angle of discuss the longeron and interface related structural attack and high velocity were calculated. After substantiation results. Table 4.2.1 lists the margins of calculating all these loads the highest value was found safety which were determined in this process.

to be the moment due to lift at 4.4 g's. The proportion of the loads on the spars turned out to be Table 4.2.1: Cabin Margins of Safety 91% on the front spar and 21% on the rear spar. The load found to be on the front spar was 55.8 thousand MS**,= 7.16 normal pounds and on the rear 19.4 thousand pounds. A MSb., _ 8.71 normal margin of safety of 1.02 was used to calculate the shear of the bolt and then the dimensions of the lug.

MSb==_ 0.05 bending MS==a= 0.82 shear 4.0 Structural Substantiation MSa, =_ 2.25 shr bkl 4.1 Nose Assembly 5.0 shear M_flr lint 4.1.1 Loading 4.62 buckle MS n, From the loading considerations discussed in section 3.1, a required moment of inertia, I, for the nose longerons was found. This was accomplished by 4,2.1.1 Normal Loads solving P=pi2EI/L 2 for I, where E for 2024-T3 aluminum and longeron length, L, were used. The P The fn_'t loading considered was the normal loading value used was the failure load factor times the constraint. From the constraint, it was calculated that a weight of the airerafL This was done for both the 8g required cross sectional area, A,_, of 0.739 square inches must be used to resist normal load failure. This was and 17g crumpling loads. Both loads resulted in similar I values. An extrusion cross section, whose found by solving f,n.,,., - (P===)/(A,.q) for A_. The cross properties matched the loading requirements, was then sectional area of one center longeron, A=.= _,_, was selected.

designed to be 1,433 square inches, thereby showing that the center structure alone easily supports the normal load.

The corresponding stress in the total center structure, f.,==, 4.1.2 Fastener Spacing Once the sizing was completed for the longerons, the was found to be 5,024 pounds per square inch from f.,,,= fastener spacing for each longeron was considered.

= (P=,=)/(A=_ z=_). The resulting margin of safety was The fasteners were sized to be four times the skin found to be 7.16 from MS==== (Fyc)/(f,==) - 1.

thickness of 0.03 inches, resulting in 0.125 inch fasteners. The skin thickness also determined that the The buckling of the center structure was next examined longerons be 0.09 inches thick (three times the skin from the normal loading condition. A required moment of inertia of 0.00377 in4 was determined from the loading thickness). The minimum edge distances for the fasteners were then found to be 0.25 inches (two constraint. This was found by solving P**,= = pi=E(I,.,)/L = times the fastener diameter). Minimum and for I_ using E for 2024-'1"3 and the longeron length, L maximum spacing for the fasteners were determined The moments of inertia of one center longeron were next calculated and the lower of the two values was selected to be 0.5 inch (four times the fastener diameter) and 1.0 inches (eight times the fastener diameter), for the design. The design stress for the total center respectively. These dimensions were used to lay out structure, fh,,,a_, was found from f_, = piZEfI_/(A,=,,,, the fastener patterns shown in drawing F93-1C-128-2. _.q)L 2 to be 4,223 pounds per square inch. The resulting margin of safety was 8.71 from _ = (Fyc)/(f_=) - 1.

4.2 Cabin Assembly 4_..1.2 Bending Loads 4.2.1 Longeron, Cabin Side, and Roof Design In response to the loading criteria discussed in section Bending loads were next examined, Section 3.2.1.2 3.2.1, various sizing requirements were found for the discussed the process used in determining the loads on r each longeron in detail. In this section it was torque on the fuselage, the floor structure must be designed to resist buckling under the torque. The stress mentioned that the bending load is distributed between the center, two-inch, and one-inch longerons. resulting from the shear flow was determined with the The load percentage for each and how each was floor skin thickness, t_,_ _,=, of 0.06 inches. The shear determined is discussed in the section. In this flow found in section 37.2.1 was divided by the floor skin thickness. This produced the maximum allowable shear procedure, the sizing of each longeron was completed as part of the process. As was mentioned in section buckling stress in the floor cross section, f_r _,_, of 3.2.1.2, the loadings were found using a margin of 2,000 pounds per square inch.

safety of 0.05. This margin of safety was designed The floor skin thickness was then used to determine the into the f,ao._ value which was used in determining the load carried by each longeron. Therefore, the critical shear buckling strength of the floor skin, F__ _,q,, margins of safety for the center, one-inch, and two- which was found from F,**, _ = kE(t_**,,_/b) 2. In this equation, a k value of 3.6, corresponding to simple inch longerons are all 0.05 and each longeron has a supports, and the E value for 2024-1"3 aluminum were stress of f.,.,,, by design.

used. The b value was the spacing between the longerons 4.2.1.3 Interface Fasteners in the floor, which is 4.5 inches. The resulting F_**,,,_ In response to the interface fastener shear load, value was 6,492 pounds per square inch. This yielded a P,,_,a_, determined in section 3.2.1.3, the fastener margin of safety of 2.25, found from MS_*** _ = (Fn_ strength was considered. Aluminum (2024-T3) _.)l(f_. _..) - I.

fasteners of a 0.125 inch diameter were selected to be used at the interfaces. Table 1, "Shear and Bearing The fastener bearing load, Pb,, _._ f,,,_, determined in section 3.2.2.1, was used to check the floor fastener Strengths of Aluminum Alloy Rivets," in Aerospace Systems Detail Design (Lade.sic, 1993) was used to strength. Aluminum (2024-'1"3) fasteners of a 0.25 inch diameter were selected to be used at the interfaces. Tables determine the fastener strengths. The value found from this table was an allowable shear load, 1 and 2, "Shear and Bearing Strengths of Aluminum Alloy P,now,_._, of 429 pounds on the fasteners, which is Rivets," in Aerospace Systems Detail Design (Ladesic, far greater than the P_.,n,_ value of 236 pounds. The 1993) were used to determine the fastener strengths. The value found from these tables was an allowable shear load, resulting margin of safety was 0.82, found from P.,,.,,,n**, fame, of 1,440 pounds on the fastener, considering MS_,_,_ = (P,n.,,,=_,)/(P_-_.) -1.

the 0.06 inch floor skin thickness. This produced a margin 4.2.1.4 Fastener Spacinj_ of safety of 5.0, found from MS_**, i-.,_ = (Pb,ja_ Once the sizing was completed for the longerons, the f_)/(Palo,,,tl--_Jm_r ) ° 1.

fastener spacing for each longeron was considered in the same manner as for the nose assembly. The identical process to that discussed in section 4.1.2 was 4.2.2.2 Buckling Loads completed for the cabin longerons. Identical results The buckling load on the floor produced by the seats and were determined for the cabin longerons, namely pilots on impact was determined in section 3.2.2.2 to be 0.125 inch diameter fasteners through a 0.03 inch In**,. _ee. = 3,420 pounds. The stress resulting from this skin. The skin thickness also determined that the load was found from f__, _,_. = (Pn._. _)I(A_,,-. ,.=_) to longerons be 0.09 inches thick. The minimum edge be 2,714 pounds per square inch. The critical buckling distances for the fasteners were also found to be 0.25 strength of the floor ribs was then required to be found in inches and minimum and maximum spacing for the order to compare with fn**,, b_d.- This strength was fasteners were determined to be 0.5 inch and 1.0 determined from Fn_,,_ = kE(ha_ .-Jo) 2. In this equation, a k value of 3.6, corresponding to simple inches, respectively. These dimensions were used to lay out the fastener patterns in the standard longeron supports, and the E value for 2024-T3 aluminum were mounting brackets shown in Figure 2.2.1. used. The b value was the spacing between the longerons in the floor, which is 4.5 inches. The resulting Fa._ad_ 4.2.2 Floor Assembly value was 15,264 pounds per square inch. This yielded a margin of safety of 4.62, found from MS_,_.b.a_ = 4.2.2.1 Torque fF_._/(f_._._)- 1.

In order to meet the requirements of the imposed at- w 4.2.2.3 Fastener Spacing Table 4.3.2: Drive Shaft Margins of Safety The same process discussed in section 4.1.2 was completed for the floor fastener sizing and spacing. Buckling Shear The resulting fastener diameter for the 0.06 inch floor Primary Shaft 0.81 2.13 skin was 0.25 inches. The minimum and maximum spacing were found to be 1.0 and 2.0 inches, 2.68 2.13 Secondary Shaft respectively, with a minimum edge distance of 0.5 inches. The thickness of the floor ribs was designed to be 0.09 inches to allow for a maximum of three Figures 4.4.1.1 & .2 clearly illuswate the "over-design" of the spar in its ability to meet the required section modulus.

times the 0.03 inch fuselage skin thickness, as well as the 0.06 inch floor skin thickness.

The required curve already incorporates a margin of safety of 0.01 in material strength characteristics, and yet the margin of safety between the designed spars and that 4.3 Drive Shaft Assembly required does not drop below 0.231.

The drive shaft was sized for buckling and torsion 4.5' loads. The design for the shaft does not require a 5 4, transverse load. Calculations were carried out to E determine the proper thickness of the shaft to handle C 3.5, r 3' buckling and shearing tendencies. A shaft thickness of ON 3.5' 0.0625 (1/16) inches was determined as substanfal to resist failure in either mode.

L 1.5- Table 4.3.1: Allowable Shear and Buckling Stresses r 0.';" Actual Shear 0 I I I I 1 I I I_,, Buckling 3_ 46 6_ 76 9_ _06 _21 _36 tS_ t66 _e_ Values

(Allow.) (Allow.)

BUTTOCK LINE LOCATION Primary 17.20 31.2 ksi 54.4 ksi Figure 4.4.1.1: Front Spar Section Modulus Shaft ksi Requirements Secondary 17.20 64.0 ksi 54.4 ksi Shaft ksi SEc'rloN M_ 2.5"

L-" I

Table 4.3.1 shows the allowable and calculated stresses in the two shaft systems. By comparing these numbers, it is seen that the shaft will buckle before it '1.5 actually shears in the primary section only. Using the buckling stress as the critical stress the thickness of the shafts was determined. The thickness was set 0,5 at 0.0625 inches to provide insurance of design in the primary section and then maintained throughout the 27 42 57 ")2 E? 102 1`1") `132 'H? `11_ 17"/ lg:2 secondary sections for producability reasons. Values BU3"FOCK LINE LOCATION from Table 4.3.1 indicated the following margins of Figure 4.4.1.2 Rear Spar Section Modulus safety.

Requirements Load requirements for the front and rear were taken at the 4.4 Wing Assembly worst case scenario: maneuvering speed and high angle of attack for the front spar, and dive speed, low angle of attack for the rear spar.

4.4. I Spar Design r Cutaways in the spar web reduce the overall weight F,,_t'=k_F,.,_. The margin of safety values in Table of the slrucun-e while at the same time allowing the 4.4.3.1 correspond to MS,_=F_.,_'/f -1.

moment of inertia of the cross section to remain fairly constant. The result is that the section modules Sections 4.4.2.1 through 4.4.2.3 discuss the skin, rib, and remains almost constant. Stress concentration around stringer structural substantiation. Results from this process are listed in Table 4.4.2.

the lightening holes will reduce the strength of the spar. Further su_ctural substantiation must be considered for these points. Table 4.4.2: Skin and Rib Margins of Safety MS_7o-19o 8.95 4.4.2 Wing Skin and Ribs MS_1,2-170 0.69 4.4.2.1 Torque MS_ffi.1_1,2 4.9 To determine rib spacing, the skin panels are designed not to buckle. This is accomplished by assuming a MSaffi.9_-120 2.68 rib location and obtaining an estimate of its area.

1.18 This area is used to determine the shear flow in the skin and ribs. Shear flow for both the skin and the ribs is computed from q=T/2A. To keep from buckling, the stress is calculated by dividing the shear flow by the appropriate thickness, f=q/t, and this is compared to the thin plate critical buckling strength F,.,_=KE(t/b) 2 where K is determined from Figure 5.4.6, in the Nui text page 139, using a conservative two side simply supported curve with the appropriate a/b ratio. For the ribs, a/b is shown above. For the skin panels, b is the spanwise dimension between the known rib location and the assumed rib location, and a is the chordwise dimension between the front and rear spars and the stringers placed in between.

To size the skin, an average torque value between ribs is determined from the graph. It is assumed that this torque is distributed evenly across the a/b dimensions discussed above.

4.4.2.2 Stringers To size the rib, the torque is obtained from the graph To reduce the shear flow in the skin panels, stringers are at the rib location. The torque is assumed to be positioned at chord positions discussed in 2.4.3. It is carried only within the wing box between the front assumed that the stringers act as beams fused at both ends and rear spar. Therefore, only the area within the where the load P is equal to the shear flow times the wing box is used to determine buckling. Once it is length of the panel. To size the stringer, the load is set determined that the ribs have a positive margin of equal to P._=pi2EI,_/L 2 and the required moment of inertia safety tightening holes can now be addressed. is calculated using the largest shear flow seen by the stringers multiplied by the length of that panel For ease of manufacturing, a 2 inch diameter flanged (q=19.081b/in and l--40in @LBL56-96). The resulting hole is used where possible. The holes occur in rib I_----0.0167 inches'.

locations of LBL96 outward to rib location LBL170.

The holes are to be centered between the stringers Knowing the flange requirement of 4 times the rivet which determines D/b needed for Figure 6.2.3 (Nui, diameter equaling 0.5 inches, a z-swinger was chosen from page 165). Using the hole diameter divided by the Ladesic. The NAS346-3 equal leg, extruded z-slringer has a flange length of 1.0 inch and I_----0.0124 inches*. It is rib height, D/h, and D/b a value of k_ is determined.

0.625 inches in height and has a cross sectional area of This is the value used to adjust F,._, where r moduli were 5.35 and 1.31 cubic inches in the lift and 0.1968 inches 2. This results in the smallest margin of safety about the length of the stringer to be 0.062. drag directions, respectively.

The resulting maximum stresses corresponding to the 4.4.2.3 Fastener Spacing section moduli were found by dividing each corresponding For riveting the skin to the ribs, spars, and z-stringers, two different size rivets are used to maximize spacing moment by the section modulus. The resulting stresses for and reduce drag. The diameters are determined by 4 the front catrythrough were 23,736 and 7,723 pounds per times the skin thickness in the appropriate section. square inch in the lift and drag directions, respectively.

The MS20470DD-3 rivet is used from the tip up to For the rear carrytlu_gh they were 14,348 and 16,389 and including the rib at location LBL142. The pounds per square inch. The resulting margins of safety for each carrythrough in the Lift and drag directions were maximum spacing of 8 times the diameter in this section is 0.75 inches. The largest shear flow in this found using the limit load determined previously. This section (rib LBLI42) is q=18.4 lbfm. this is resulted in the margins of safety in table 4.4.3.

compared to the o_,,, which is equal to the allowable 4.4.3.2 Fasteners shear strength of 241 pounds (Ladesic) divided by the As mentioned in section 3A.3, the fasteners are required spacing, resulting in q,a,,,--321 lb/in. The margin of to carry a shear load of 3,848 pounds distributed over each safety is computed by MS_,---q_.,/q -1 and is equal to 16.4.

half of the carrythrough structures. The maximum load required in each fastener was found from Aerospace In a similar manner, the rest of the skin uses Systems Detail Design (Ladesic, 1993) table 1 on page MS20470DD-4 rivets. The maximum spacing is 1.0 151 to be 429 pounds for a 1/8 inch 2024-T3 rivet.

inch with the largest shear flow of 79.0 lbfm Therefore, the minimum number of rivets required to carry occurring at rib location LBL56. The rivet allowable the shear load is 9 in each half of the carrythrough, or a total of 18 minimum rivets in the each structure. The shear strength is 429 pounds resulting in a o_._--429 lbhn. The margin of safety is 5/4.

minimum spacing requirements for the rivets, however, required a total of 180 rivets total in the structure the resulting margin of safety is 9.0, as seen in table 4.4.3.

4.4.3 Carryth_ugh 4A.3.1 Loading Constraints 4.4.3.3 Fatigue From the loading constraints discussed in section As mentioned previously, one of the greatest requirements 3A.3, a required section modulus for 100% lift in designing the carrythrough was fatigue performance. In capacity on a carrythrough was determined. This was order to withstand the required safe life of l0 T load cycles, done by first applying a factor of safety of 1.5 to the the _ugh smmmre cross section was determined to ultimate tensile strength. The resulting limit load was be as appears in drawing F93-1C-170-2. The process used used with a margin of safety of 0.01 to determine the in evaluating the fatigue characteristics of the maximum allowable stress. The stress was then used wing assemblies is discussed in detail in section 7.7.

in f=Mc/I to solve for the section modulus (_I/c), where the moment used was the moment due to lift of The front carrythrough was the most critical, as it was found that the calculated maximum stress, due to lift, was 365,541 inch pounds as determined in section 3.4.3.

The resulting section modulus was 9.07 cubic inches 11,869 pounds per square inch, versus a maximum in the lift direction. The same process was completed allowable of 12,000 pounds per square inch (found from for the drag moment of 21,470 inch pounds, resulting MIL-I-IDBK-5E figure 3.2.1.801)). The rear carrythrough was found to have a calculated maximum stress of 7,174 in a required section modulus of 0.533 cubic inches in the drag direction. pounds per square inch versus a maximum allowable stress of 9,000 pounds per square inch. For drag, the front and rear carrythroughs had calculated maximum Next, the computer program Mathematica was used to calculate the section moduli of the front and rear stresses of 3,861 and 7,450 pounds per square inch, carrythrough structures. The resulting moduli for the respectively. The corresponding maximum allowable front carrythrough structure were 15.4 cubic inches in stresses were 8,000 and 10,000 pounds per square inch, the lift direction and 2.78 cubic inches in the drag respectively. Therefore, since in all cases the calculated maximum stresses are less than the maximum allowable direction. For the rear carrythrough structure, the r t stresses, all components possess a minimum safe life 5.1 Nose Assembly of 10 7 load cycles.

The nose assembly was designed for simplicity of Table 4.4.3: Carrythrough Margins of Safety construction and ease of access to the nose bearing. The longerons are to he Alclad extruded channels of 2024-T3 aluminum, as previously mentioned. The Alclad coating Component Load f="_ M.S.

is to protect the longerons from corrosion. These channels (psi) are to he contour rolled to create the proper curvature required of the nose assembly. The longerons are then to Front Carry- Lift 23,736 0.698 through be mounted to the prop bulkhead and nose frame, and nose bulkhead nose frame. The mounting brackets will Rear already be fastened to the bulkheads and frame. The Carry- Lift 14,348 1.81 longerons will be fastened to the brackets. These brackets, through as well as the prop bulkhead and nose frame, are to be machined from permanent mold castings of 202A-T3 Front Carry- Drag 7,723 4.27 aluminum, which are also to be Alclad. All fasteners used through are cadmium plated MS 204070DD-4.

Rear Can'y- Drag 16,389 2.48 The skin sheets and access panel are to be made of 0.030 inch thick Alclad 2024-T3 aluminum sheets. Each skin through sheet is to have two half inch angles riveted on to it for Fasteners Lift 9.0 aerodynamic loading support. These channels are to be Alclad 2024-T3 extrusions which are contour rolled to the proper nose curvature. The skin panels, with angles in 4.4.4 Wing Attachments place, will then be riveted to the nose longerons. All The lug design was chosen because of the ease of rivets for the nose assembly are to be cadmium plated assembling and disassembling. Only three bolts need MS20470DD-4 rivets, as determined previously.

to be installed in order to put the wing on. The use of the quadruple shear lug on the front spar allows for The access panel is to be mounted using AN526C-6-32 a smaller bolt diameter. The channel inside of the screws. The screws will he fastened directly into the front spar carry through was designed to carry the longerons, bulkheads, and nose frame, which all will have holes threaded to suit the screws machined in them. This tensile and bending loads into the structure and allows for a double row of rivets to dissipate the load is to allow the panel to he firmly mounted in place, while throughout the suucture. The 1-beam design for the allowing for it to be removed with ease. The access panel rear spar is oversized for the load, but allows for the is to allow easy access to the drive shaft bearing and easy installation of the lugs into the structure. The bearing support mounted on the inside of the prop lug can be slid into the structure and riveted into bulkhead. Sufficient room is made available by the panel place with the double row of rivets on either side of opening for easy mounting, inspection, and replacement of the I-beam web. The web carries the load from the the parts.

bolt lug to the I-beam structure. The margin of safety for the lug assemblies was selected to be 0.05. 5.2 Cabin Assembly The spar attachments have many cycles applied to 5.2.1 Longeron, Cabin Side, and Roof Design them each flighL The lugs are very accessible to The cabin structure was designed for ease of fatigue failure. The shear stresses for the bolts and manufacturing and maintenance, as well as part lugs were at a mean stress of 5.3 ksi and below. This commonality between the cabin and the existing gives a maximum stress of 11.7 ksi. At this stress empennage design. The empennage design used 2024-T3 aluminum in all of its key parts such as longerons, rivets, level the part has a life of one hundred million cycles.

and the like. The same material, Alclad 2024-T3 aluminum, was used in all components of the cabin 5.0 Manufacturing and Maintenance design, except for the fn'ewall, which is made of 17-4PH stainless steel.

Thecabin is tobemanufactured by incorporating the 2024-T3 aluminum, which are shown in Figure 2.2.1.

floor, insu'ument panel, windshield and door These brackets are common to mounting all of the two- assemblies into one structure. This is done by inch longerons in the cabin and floor assemblies, and are also referred to as "standard" mounts due to this fact.

starting with the floor assembly (the floor assembly is explored in detail in section 5.2.2). The drive shaft box assembly is first mounted to the floor assembly. The roof assembly may now be consmacted. The roof frame is constructed of three channels. Two are connected The instrument panel assembly will be mounted to the with standard mounts to the firewall frame. These mounts drive shaft assembly at this point (the drive shaft and instrument panel assemblies are discussed in section are to be attached with temporary fasteners at his point, 5.3). later to be riveted to the Upper engine/wing box longerons.

The third channel is to be mounted between the two side channels. This channel is mounted with standard mounts The ftrewall, including the angle frames sandwiching at either end, each end riveted to the sides of the side the firewall is mounted to the drive shaft assembly at this point. The firewall is manufactured from 0.02 channels. This frame now forms a box, with the forward inch thick 17-4PH stainless steel sheet, in accordance channel at the front, the two side channels on the sides, with FAR 23.1191 which requires a minimum 0.015 and the firewall frame forming the back.

inch thick stainless steel ftrewall. The angle frames are each made up of four 0.75 inch by 0.75 inch by The forward bulkhead may be installed at this time. The 0.09 inch thick extruded angles. The channels are to bulkhead is formed similarly to the nose bulkhead. Its be contour rolled into their required curvatures. The angle frame is formed in the same manner as the fuewall channels are each alodined on the firewall interface frame. Yet, only one angle frame is constructed, not a two channel sandwiched frame like the firewall. This is side and riveted to the ftrewall with cadmium plated MS20470DD-4 rivets. The firewall is also to be to allow the bulkhead, which is actually an access panel, mounted to the floor longeron mounts with to be screwed into place on the angle frame. AN526C-6- 32 screws are to be used here, as were used in the nose MS20470DD-4 rivets. However, at this point in assembly, temporary fasteners are to be used to fasten access panel. Due to this construction, the bulkhead may to these mounts. This is to keep the fastener holes be easily removed for access to the control systems and open for later riveting to the carrythrough longeron the nose gear. This access is extremely easy, as it is a mounts on the other side of the firewall.

large opening, thus allowing for ease of maintenance and or inspection of these areas.

At this point, the JAARS seats are to be installed on their mounts (which are already a part of the floor Once the nose assembly and carryOtrough assembly have assembly). The seat belt (a four point harness) may been riveted into place at the longeron mounts, the skin may then be installed on the cabin. The skin sheets are to also be installed at this point. Next, the nose bulkhead be 0.03 inch thick Alclad 2024-I3 aluminum sheets. Half is mounted to the floor longeron mounts, similarly to the firewall mounting. This is to keep the fastener inch angles will be riveted to the skin panels that are mounted on the side of the cabin, similar to what was holes open for later riveting to the nose assembly longeron mounts on the other side of the nose done on the nose skin. These andes are to placed such bulkhead. The nose bulkhead is cut from 0.020 inch that they are mounted half way between the upper two- thick Alclad 2024-T3 aluminum sheet and is inch longerons and the two-inch floor longerons. The door, windshield, and cabin light assemblies may now be assembled to its angle frame in the same manner as the firewall. The door frame assemblies may now be mounted to the fuselage. Once the canythrough, wing, mounted to the floor assembly at the floor attachment and empennage sections are joined to the forward fuselage assembly, the aircraft will be painted with white enamel points.

paint to further protect the skin from corrosion. Various Next, the upper two-inch longerons are to be mounted colored enamel paints may then be used for detail graphics on the aircraft.

between the nose bulkhead and door frame and between the door frame and firewall. They may be riveted to the door frame, but temlx)rary fasteners 5.2.2 Floor Assembly must be used in the nose bulkhead, similarly to the Again, keeping in accordance with the rest of the cabin structural design, all materials in the floor assembly are floor longeron mountings. The longeron mounting brackets are to be Alclad permanent mold castings of Alclad 2024-T3 aluminum. The commonality of parts is r assembly. This is done by blind riveting MS204070DD-4 maintained once again, as the same standard mounts rivets through the bottom skin into the floor and seat for the two-inch longerons are used. The cadmium plated MS20470DD-4 rivet is also used again support rib flanges. This completes the floor assembly.

throughout the floor structure, similar to the nose, cabin, and empennage assemblies.

5.3 Drive Shaft Assembly The first components of the floor assembly to be The &ive shaft itself is made of 4130 steel alloy. It is constructed are the floor and seat support ribs. These are to be cut from 0.030 inch Alclad 2024-T3 3.25 inches in diameter and 0.0625 inches thick.

aluminum sheet. The cutout pattern includes cutouts Manufacturing of the drive shaft sections should begin for the longerons to pass through in the floor ribs, with extrusion of steel tubing of the indicated outer diameter and thickness which would then be trued to when installed, as shown in drawing F93-1C-128. The reduce vibration while spinning. Both male and female cutout patterns are then hydropressed into the required shapes. The seat support ribs are then to be riveted halves of the spline interface should be bored separately to the floor ribs with MS2IM070DD-4 rivets.

and plasma welded to the trued shaft. Spline sections are also 4130 steel alloy. Similar metals are needed to ensure strong weld connections and the hard alloy (in comparison The second major floor component is the two-inch to an aluminum alloy) is needed for proper spline channel longeron. These longerons are brake formed from 0.09 inch thick Alclad 2024-T3 aluminum and interface. Extreme care should be taken in the balancing then contour rolled into the required curvature for of each shaft assembly, since this will be the primary assembly. One-inch floor longerons are also formed somr, e of vibration in transmission of torque to the prop.

for the floor. These are extruded and contour roiled to the required assembly curvatme. A secondary drive shaft is connected to the primary shaft in an almost identical fashion as at the shaft to engine interface. The forward spline gear connection, however, Once the ribs are formed, they will be mounted to the uses a 0.46 inch diameter aluminum shear pin to aid in 0.06 inch thick floor skin with cadmium plated MS20470DD-8 rivets. Cutouts in the floor skin are engine torque transmission and to ensure resistance to crumpling of the front nose assembly until an 8 g impact provided toward the aft of the floor, through which the elevator trim and elevator control linkages are to has been reached. This assembly includes the thrust pass. The JAARS seat track is fastened through the bearing attachment which is bolted to the nose cone bulkhead.

floor skin and the seat support rib's flanges at this point with MS20509DD-6 countersunk rivets (the seat may he slid into place on the tracks later in the cabin The drive shaft uses three steel bearings with aluminum assembly, as spaces are cut in the track for seat housing and mounting brackets to cut down on weight.

removal). Half inch angles will now be fastened to The bearings and housings can be subcontracted to find a the underside of the floor skin with MS20470DD-4 light weight durable bearing. New composite materials may also provide a lighter more effective design. Bearing rivets. The angles are to be spaced 4.6 inches apart, as shown in drawing F93-IC-I28-2, to provide extra design, however, should incorporate use of split journal stiffness in the floor. The two-inch channels must bearings. This will aid in assembly of the drive shall now be riveted into place, above and below the floor skin and into the floor ribs with MS204070DD-4 The universal joint is a standard universal MS20271 which rivets, thus forming the frame of the floor assembly. can be joined to the rear spline section. The universal The top of the floor assembly is now complete. meets military specifications and should last the life of the aircraft. It has a permanent lubrication and should not require any maintenance.

The one and two-inch floor longerons will now be riveted to the bottom skin sheets, thus forming the bottom of the floor assembly. All attachments are The universal joint is connected to the engine by a spline made with cadmium plated MS204070DD.-4 rivets. gear attachment. The bearings and mounts are attached to The standard mounts must all be installed at this point the drive shaft and are inserted together through the front of the aircraft with the nose cone removed. The bearings in the ends of all of the two-inch channels, using the are then mounted to the ftrewall and center structural technique mentioned previously. The bottom floor members.

assembly may now be attached to the top of the floor Each structural member is made up of two channels. with the exception of the ribs surrounding the fuel tank.

The channels are made from 0.125 inch thick 2024 Four 2 inch diameter holes are centered between the z- T3 aluminum. Each channel is to be brake formed stringers. The holes provide a reduction in weight and with .5 inch minimum bend radii. All four channels room for maintenance and inspection. The ribs are riveted are identical in dimensions and shape and, therefore, to the front and rear spar using MS20470DD-4 rivets from allow for part commonality. LBL31 up to LBL142 and MS20470DD-3 rivets are used from LBL142 to the tip.

All electrical, vacuum, and control systems are to be installed prior to installation of the drive shaft Once the ribs are attached, the stringers can be riveted to assembly. In order to simplify assembly, installation the top and bottom of the ribs at LBL56 and the tip of the instrument panel and seat assemblies should positioned at 36.7%, 48.4% and 60.2% chord. The follow that of the primary shaft. stringers are equal leg, extruded aluminum z-stringers (NAS346-3) and are riveted using the rivets as discussed above.

For routine maintenance of the bearings, the upper shield which covers the primary shaft, as seen in drawing F93-1C-152-2, can be removed for access Finally, the skin can be flat wrapped and riveted into from within the cabin.

position. ,again, the skin used is Alclad 2024-T3 aluminum and comes in standard thicknesses and in 48 The electrical system and control systems can be inch wide sheets. All joints me overlapped and all of the accessed through the side panels of the drive shaft skin panels must be in position before riveting can begin, channel. These panels are attached with quick release fasteners to allow maintenance entry into the channel. First, the upper and lower skin panels are temporarily held The top of the channel is also removable, so the drive into position while the leading edge skin panel is shaft itself can be inspected and maintained.

positioned over them. This is done across the entire semi- span until all panels are into position.

Drive shaft removal would have to be done in a procedure similar to that of a reverse installation. From the tip inboard to the rib positioned at LBL142, a The nose cone would be removed by quick release 0.02 inch thick sheet is riveted using the MS20470DD-4 fasteners on the skin.

rivets. Two panels of 0.032 inch thickness m,e used from LBL142 inboard to LBL96 and from LBL56 inboard to LBL31 and 0.071 inch thick sheet is used over the fuel 5.4 Wing Assembly tank. All three panels are riveted using the MS20470DD- 4.

5.4.1 Spar Design The front and rear spars are both manufactured in Consideration must be given to cutouts for the fuel tank similar fashions. 2024-T3 aluminum alloy sheets of access, landing gear, and maintenance access panels.

the appropriate thickness, 0.08 and 0.06 inches, are to These concepts were not specifically addressed in this be blank pressed to the specified pattern. The assignment.

blanking operation will identify the flange edges and the lightening cutaway's. The flanges are then formed 5A.3 Carrythrough by use of a power brake. Because of the sheet size The carrythrough assembly was designed to be simply limitation, each spare is made of a 96 inch inboard constructed and utilize parts common to the existing cabin and a 67 inch outboard section. The outboard section structure. The major load carrying items are the forward is fish-mouthed at the 5 inch overlap and joined by 10 and rear channels and plates. The plates are to be cut MS21M70DD-4 rivets. from 0.1 inch thick 2024-T3 aluminum sheet. The channels are to be extruded 2024-T3 aluminum in the 5.4.2 Wing Skin and Ribs cross sections depicted in drawing F93-1C-170-2. The one and two inch longerons used in the assembly are the All materials within the wing box and skin panels are Alclad 2024-T3 aluminum. The ribs are formed in same as those used in the cabin s_mcture, brake formed similar fashion to the floor ribs discussed in 5.2.2 from 2024-T3 aluminum and contour rolled to the required using the appropriate thicknesses described in 3.4.2. curvatures. The longeron mounting brackets are also identical to those used in the cabin structure, 2024-T3 In addition, lightening holed are stamped into the ribs r permanent mold castings. assembly is now complete.

The fast items to be assembled are to mount the wing 5.4.4 Wing Attachments attachment lugs to the forward and rear channels, as The lugs are machined out of 2024 bar stock. The discussed in section 5.5. Next, the upper outboard machining of the lug will allow for a higher shear stress two-inch longerons are to be riveted to the front and allowable than a cast part. The hole for the bolt will have rear channels with the mounting brackets, as shown to be drilled very precisely to make sure the holt fits in drawing F93-1C-170-2. Mounts may also be tightly. The tight fit of the bolt is necessary to maintain installed in the lower inboard two inch longerons at the strength desired of the lug and bolt. The channels on this point. the front spar attachments are tapered at the ends to not allow a stress concentration at the end of the attachment.

The front plate and channel, with the two-inch The same is true with the I-beam on the rear spar in which the web is fish-mouthed.

longerons attached, may now be riveted to the ftrewall. The rivet locations of the floor longerons will be picked up. Therefore, the temporary fasteners There is no maintenance required for the attachments in those holes (discussed in the cabin structure except for the continuous inspection of the parts for manufacturing section) may now be removed and corrosion or fatigue displays. As shown by the fatigue replaced with MS204070DD-4 rivets which will calculations, the parts should last more than one hundred fasten the interface between the floor longerons and million cycles, but this is for normal flight. Therefore, for the front carrythrough structure through the firewall. safety reasons the lugs should be inspected.

The same will be done to install the inboard two inch longerons. Their mounts will be riveted to the lower flange of the front channel at the locations of the 6.0 Weight Summary, floor longerons, thus picking up the fastener locations there. As before, the temporary fasteners in the floor The weights of all components used in the conslruction structure will be removed and replaced with of each assembly were calculated individually and then MS204070DD-4 rivets at this point, thus completing summed up for each total assembly weight. The resulting the interface from the floor longerons through the assembly weights are shown in Table 6.0. In addition, the firewall to the front carrythrough and inboard two center of gravity for the cabin-fuselage assembly and the inch longerons. occupant groupings are included in the Spatial Requirements Specification Document (refer to Appendix A).

Next, the inboard two inch longerons will be fastened to the lower flange of the rear channel with temporary fasteners through the longeron mounts, rear channel, In the design of the cabin structure, there is no place and rear plate. Temporary fasteners will also be where weight could further be reduced, as all components installed at all rivet locations in the rear channel and were designed to be as small as possible while still plate to fasten the plate and channel together. At carrying the required loads. However, in the cases of the installation of the empennage and rear bulkhead, these drive shaft and wing assemblies, further weight reductions temporary fasteners will be removed and permanent are possible, as will be discussed below.

fasteners will fasten the interface of the rear carrythrough structure through the rear bulkhead to The high margins of safety of Table 4.3.2 indicate a small the empennage longerons. sacrifice on the designer's part in respect to the drive shaft's overall weight. The purpose of the high values, The skin is now ready to be fastened to two inch however, is twofold. The first is that priority was placed longerons to form the outer shell of the fuselage on occupant safety, and the second is that an increase in surrounding the carrythrough slructure. Once the skin drive shaft thickness is expected as further research is in is place, the one inch longerons may be riveted to done on the fatigue characteristics of the shaft.

the skin at the positions shown in drawing F93-1C- 170-2. The exact skin thickness and corresponding Due to the load distribution on the wings, the curves rivet size has not been determined by the determining the spar section moduli required the outer requirements of this report, but these will be portion of the wing to be over designed. This was determined in future efforts. The cmrydu'ough necessary to retain the desired taper in the spars. In the r front carrythrough, fatigue determined that the size of With respect to temperature, all parts are designed to the front carrythrough structure could not be reduced. operate without degradation from -40 F to 122 F. The The rear carrythrough size could be reduced as far as 2024-1"3 aluminum and stainless steel parts used in the loading and fatigue constraints are concerned. It was construction all easily withstand this tempemane regime over designed in these areas due to the desire to without material degradation. With regards to the cabin maintain similar construction as the front earrythrough environment in this temperature range, the heating and ventilation systems account for pilot comfort. No actual for ease of manufacturing and cost reduction. The size requirements for interface with the rear spar were insulation is to be included in the cabin structure, beyond also a contributor to the design selected. the skins, longerons, and interior walls (standard plastic wall coverings). However, the healing system is to keep In comparison of the preliminary weight estimation to the pilot sufficiently warm, when he is properly dressed, the total weight listed in table 6.0, the preliminary within the cold extremes. The venting system is to cool estimation was 117 pounds less. This is due to the the pilots at the upper temperature extremes. These fact that the preliminary estimation did not account concepts are the same as those currently applied in today's for the weight incurred by the drive shaft assembly. existing primary trainers.

This comparison is made to the listings for fuselage, cabin systems, and wing weight entries in the 7.2 Atmospheric Pressure preliminary report versus the values of table 6.0.

The Viper is designed to operate up to an altitude of Table 6.0: Subassembly Weights 10,000 feet, which does not require pressurization or supplemental oxygen (per FAR 91.21 l(a)). Therefore, the Nose 31.42 lbs structure as it has been developed will be capable of operation up to this flight level.

Floor 67.21 lbs 7.3 Sand and Dust Cabin 92.31 lbs Drive Shaft 109.57 Ibs External surfaces, mechanisms, hinges, and associated items have been designed to endure up to 150 microns in Can'ythrough 37.52 lbs size and in combinations of sand and dust in concentrations up to 0.041 grams per cubic foot without Front Spar 20.0 lbs degradation. This has been accounted for by having the drive shaft bearings properly fitted to disallow dust Rear Spar 17.0 lbs interference. The most critical point for this is at the nose Skin & Ribs 145.7 lbs interface with the prop. The prop bulkhead seals off the shaft area, thus keeping dust and sand from entering the Wing Attachment Fittings 50.0 lbs drive shaft area. The housing around the drive shaft in the cabin area also serves this purpose. In the wing assembly TOTAL 570.7 LBS structural designs, there are no critical moving parts, and the lugs for wing attachment are protected, as they are 7.0 Environmental Considerations housed within the wing skin.

7.0.1 Considerations 7.4 Rain, Humidity, and Salt/Fog In designing the various components of the overall All external surfaces have been coated in various manners, deign in this report, several environmental as discussed in section 5. These surface coatings prevent considerations were regarded. The areas considered corrosion due to exposure to rain and up to 100% relative were temperature, atmospheric pressure, sand and dust, rain, humidity, ice, snow, salt/fog, wind gusts, humidity at 95 F without degrading. They also deter and fatigue performance. Each is discussed in the corrosion due to exposure to a salt/fog atmosphere, as following sections. may be encountered in coastal areas. There are no cavities designed into the concept which allow for containment of water which leads to corrosion problems.

7.1 Temperature Again, the prop bulkhead prevents intrusion of water into r aluminum. The resulting maximum swess was found from the drive shaft area, including rainfall up to a 4.0 inch/hour rate with net wind velocities up to 150 the figure and must have been greater than the calculated miles per hour. All other structural interfaces and maximum stress in order to provide the minimum required safe life.

skins are consmlcted to easily withstand these conditions without intrusion of water into the structural interior.

8.0 Conclusions 7.5 Ice and Snow In conclusion, this report has been prepared to answer the The coatings mentioned in section 7.1.4 are also requirements of the 421F93ADP01-2 SOW and its responsible for withstanding ice and snow at Addendum. The requirements of both have been met, temperatures as low as -40 F without degradation of regarding occupant safety and volume constraints. A materials. At loadings of 10 inches of wet snow structural concept which withstands the demands of FAR accumulation, the cabin structure easily holds without 23 has been completed within this report. This was done buckling. Snow accumulation on the wings does not by heavily considering part commonality and ease of effect the loading on the nose gear, as the wings are manufacturing and maintenance. All of the parts used in the cabin structural design match the materials used in the placed aft of the aircraft cg.

existing empennage design. The same holds true for the The snow load was also checked for the wing wing structural design with respect to the cabin and assemblies and was found to be 1,II7 pounds by empennage su_ctural designs. Also, a number of derail multiplying the wing area by the 10 inch snow depth pans, such as rivets, brackets, and longerons are used and the density of wet snow (10 pounds per cubic repeatedly to reduce the amount of different part types foot). This load is far less than the load imposed at required.

4.4 g's. Therefore, the various wing assemblies are designed to easily withstand the snow load. Perhaps the greatest challenge of designing the Viper cabin-fuselage smlcture was insuring that the drive shaft 7.6 Wind and Gust and engine would cause no harm to occupants in a crash.

This was achieved, as has been discussed in this report.

The drive shaft and its supports have been designed to The structural design for the cabin and wing assemblies has been completed for flight conditions, account for this. Indeed, it has been shown in this repcra which include loads higher than would be experienced that the drive shaftand engine will not cause harm to at tie down. Winds of 120 miles per hour and 50 occupants upon crashing.

mile an hour gusts do not impose loads at tie down greater than the required inflight design loads. In order to meet the restraints of the project, the size of Therefore, the design is capable of withstanding these the fuselage was required to be increased, thus moving the tie down conditions without degradation. wings by two inches outboard to either side of the preliminary design fuselage. This size increase was due to lack of proper space am-ibuted to realistic structural volume requirements, as well as not providing for the 7.7 Fatigue Performance JAARS seat installation, in the preliminary design. Due As discussed in previous sections, fatigue was a major to the fuselage size increase, the empennage must now consideration in the design of the wing assemblies. also be redesigned to match the new diameter of the Viper The cabin and engine interface structural designs were fuselage. It is recommended that when this is done, the not evaluated for fatigue. However, all of the wing empennage longerons should be placed in line with the assemblies were found to display a safe life of 107 existing longerons in the cabin and canythrough designs.

load cycles, as required by the SOW. The process to This would allow continuous load paths through the determine if each component can endure a maximum longerons throughout the entire aircraft.

of 107 load cycles used the stress on the component at 1 g as the mean stress. The resulting maximum An additional required modification is that the existing stress was then found as 2.2 times the mean stress.

rudder pedals must be moved 6.3 inches aft and the yoke Next, the mean suess was checked at 107 cycles on must be moved 5.1 inches up and 2 inches aft. These changes are to allow for the placement of the JAARS seat MIL-HDBK-5E figure 3.2.1.8(h) for 2024-'I"3 2_ and proper volume safety constraints, which were not created entirely using existing technology, as many accounted for in the preliminary design. Otherwise, components were modeled similarly to actual aircraft the cabin smlcture is designed entirely around the components. The Cessna 152, in particular, was examined existing control systems. for this purpose, as its pilot operating handbook was used as a reference, as well as crash worthiness tests performed One final modification made to the preliminary design on it by NASA. Approximate dimensions for certain was in removing the vents from the doors, as these components were also taken from measurements on a vents were of insufficient size. They were also placed Cessna 152 structure at ERAU's AMT facility. These in an area that is not conducive to flow into the vents.

factors, coupled with the results discussed above, indicate The engine venting was moved to an intake below the that the design presented in this report is indeed a fuselage, as may be seen in drawing F93-1C-129-2.

believable concept.

This placement allows for much improved air flow, as well as a larger intake area.

Certain fuselage components, such as the door frames, windshield, door, and cabin light assemblies were not explored in complete detail. They were included as considerations in the design. These components may be examined in detail in the fuutre. Also, as was discussed in section 2.3.1, the engine support and mounts for mounting to the front of the 0-235 must be further examined in cooperation with Lycoming.

This will allow for final sizing of the engine support between the front of the engine and the center longeron structure under the drive shaft.

With regards to aerodynamic loads, such as the prop wash on the nose assembly, or gust loads on the cabin and wings in general, it is recommended that further studies be done. They should be in the form of wind tunnel testing to ensure proper reactions to aerodynamic loadings without buckling.

Crash and fatigue tests are also recommended to substantiate the loading calculations made for the design. This should be done in future studies to ensure a twenty year service life without component failure. Actual fatigue tests are recommended to ensure accurate results.

Regarding cost, a formal cost summary was never completed, due to lack of time. However, manufacturing, maintenance, and related costs were heavily considered in the design process. It is recommended that once the entire Viper aircraft concept has been completed, that a cost analysis then be run. This would ensure the most accurate cost estimates.

In general, the design presented in this report is a viable production design, once file recommended detail studies have been completed. The design was

Appendix A

Spatial Requirements Specification Document

A.1

--2

To: Dr. J.G. Ladesic From: F93-IC Date: October 20, 1993 (F93-!C-2R!)

Subject: Spatial Requirements Specification Document Introduction This report contains the occupant safety requirements set forth in Statement of Work 421F93ADP01-2. Under Federal Aviation Regulations (FAR) Part 23, the light aircraft certification requirements specify minimum volumetric constraints for occupant safety. This report is to fulfill the need to address cabin sizing at the conceptual design level for a utility category, conventional, single mid-engine, low wing, tricycle landing gear airplane configuration, integration of future design concepts must take into account these constraints to meet certification.

Requirements F__R 23.562 specifies emergency landing dynamic conditions. This requirement calls for a complete dynamic test for seat/restraint system to ensure compliance for crashworthiness. At present, Jungle Aviation _md Radio Service (Jg_S) has dynamically tested and certified one -of the only seats that meets requirements specified by F_ 23.562. To meet aircraft certification requirements, the J_S crew seat was selected for this project and FAR 23.562 is met by rational analysis.

In the event of a crash, the occupant must be restrained from coming in contact with any portion of the cabin and not to exceed the head injury criteria (HIC) value of i000 (refer to F_ 23.562).

Due to the limitation of performing this dynamic test at this stage of development, volume approximations are made to ensure that the occupant's head does not impact any cabin object.

A standard occupant in a normal seated position, shown in Figure 1 A, must have a minimum of 2.0 inches from the top of the head to the cabin's inner structure. Upon forward impact, the body's range of motion, sho-wm in Figure 1 B, is restrained by a four point shoulder harness allowing only the head to rotate about the base of the neck. This head rotation creates an arc from the base of the neck extending 2.0 inches above the top of the head. At the forward seat adjustment position, the windshield must be positioned at an angle clear of this region. In addition, the bottom of the _=_--_=_-._ panel must _ at a _4_ height of 2. _ _c _=_ above the knees. This is to ensure that the occupant has adequate room for cabin ingress/egress.

Due to the general fuselage shape, the minimum clearance is defined as the distance from the occupant's head to the cabin's inner structure occurring at a diagonal. This is displayed in Figure 1 C. Upon side impact, the bead's side-to-side motion again creates an arc from the base of the neck extending 2.0 inches above the top r of the head. Sufficient space _:ust be a!!e..,_ed for the curvature of the fuselage and the door structure.

The J_.RS seat used in this configuration can be adjusted 6 inches forward and aft of the nominal position. This adjustment creates the sDacial envelope that is the sum of the regions to be void of any obstruction as shown in Figure 2.

Finally, this minimal volumetric envelope needs to account for the cabin structure (refer to Figures 3 and 4). If the design configuration dose not include a canopy, a minimum of two inches must be allocated for the cabin's roof structure. This two inch consideration takes into account the required structural sizing for the external skin,, roof structure, and headliner to maintain structural integrity in event of an !8g impact. Additionally, another five inches needs to be allocated for the floor. The floor structure is a majo_ load path and must house longerons and ribs sized to meet the same 18g impact criteria. Also, a minimum of three inches must be allotted for door frame structure.

Conclusions This document develops the minimum spacial requirements for aircraft certification under F_ 23. A seat/restraint system certified by dynamic testing under F_ 23.562 must be used. To meet HIC using the J_)_RS crew seat, all dimensions of the spacial envelope, shown in Figure 4, must be the minimum dimensions used.

Realistic structural volume constraints must be considered in the cabin design, as well. It is left to the conceptual designer to provide additional room for adequate occupant comfort.

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(_ E d Q_ D U U r--I u1 ,'- .t_ 0 cp -i-' £ cs L L Q; _ "0 U N

_ o_

U m C3 .

D % IA Ill c L • .- "0

_o

U -- .., .p (- U "P 13 --

\

O.

I::: H

r

Appendix B

V-N Diagram

B.1 r O

o

o

O-I-J .r,°_

o

f....

r

Appendix C

Torsional Diagram

C.1

T

I l l !

I

o r"-I

J

i

\

I I | I I 0 I I

r Appendix D Structural Decomposition D.1 L_,_, 7 _ F_rJ_mUy 7

f- Appendix E Drawings E.1 _OLDOUT FRAME I " 'I" ÷ ÷÷÷÷÷÷÷÷÷÷÷ 4" ÷ k /,4- "!- r

FRF1NT SPAR ATTACHHE

STA

REAR SPAR

ATTACHHENT

IIII_

NOTEI ALL PARTS ALCL BREAK ALL SHAR ,m_ILDOUT FRAMe _ ÷ 4.

4- ÷ ÷ -I- ,I- + .I-

T

I0 G4 RIVET MS20430A-8-10 9 4 NUT-AIRCRAFT-SHEAR AN320-D16 8 2 NUT-AIRCRAFT-SHEAR ANS20-D12 7 2 FUSELAGE ATTACHMENT LUG 2024-T4 AL G 2 REAR SPAR ATTACHMENT LUG 2024-T4 AL 5 2 BOLT SHEAR, HEXAGON HEAD NAS 1310-B 4 4 FRONT SPAR ATTACHMENT LUG 2024-T4 AL 3 4 FUSELAGE ATTACHMENT LUG 2024-T4 AL 2 264 RIVET MS20430A-8-12 i 4 BOLT SHEA_ HEXAGON HEAD NAS1314-31 ITEM QTY DESCRIPTION ;MATERIAL OR PART # UNIVERSITY EMBRY-RIDDLE AERONAUTICAL UNL[$$ _li._nbrls£ SPEC,.1F_._ DAYTONA BEACH FLORIDA

sIzE _ATE SCALE ,DP&,,7,,,i :9{

,XX _+. ,Ol

B i2/2/93 1/5 CRAIG STEVENS

,XXX + ,001

TITLE WING/SPAR ATTACHMENTS DRAWING NO.

,-,, C) EDGES & BURS

± !/c

F93-IC-180-2

3TA I

I

i 03 l

WT ]. :.A-_-

_ I II

sl//

/

FLOOR LDNGERONS (REF

ENGINE VENT

N[]TES: 1 ALL MATERIALS ALCLAD 8084 UNLESS OTHERWISE SPECIFIEI 2 REMDVE ALL SHARP EDGES ANI

RFAR BULKHEAD (RFF)

8 1 REAR PLATE F93-IC-178 7 490 RIVET MS20470DD-4 6 I FRONT PLATE F93-IC-176 5 1 REAR CHANNEL F93-IC-175 4 1 FRONT CHANNEL F93-IC-174 3 8 2 INCH LDNGERCIN MCIUNTS F93-IC-II0 2 4 i INCH LCINGERDN F93-IC-I09 I 4 2 INCH LDNGERDN F93-IC-I08 ITEM QTY DESCRIPTION PART#

(REF)

EMBRY-RIDDLE AERONAUTICAL UNIVERSITY D_ T[]L[RANCES UNLESS []THERVISE SP[CIFIED DAYTONA BEACH FLORIDA SIZE DATE SCALE DRAWN BY

,XX _+_ ,01

B ii/93 i/5 MARCHESSEAULT

XXX +_ ,001

TITLE CARRYTHRCIUGH ASSEMBLY ] DRAWING NO, _F_EET BURRS F93-IC-170-2

TA

CL

(REF)

LI3NGERE_NS

FLEIE_R

VIF____ LEI[}KII'

STA 133 REAR BULKHEAD (REF) CL DDWN AT WT 44 DWG BY DRAWING NO, MARCHESSEAULT F93-IC-170-2 DATE SHEET 2 DF 3 11/93 -r_OLL_UT FRAJML /, \ ++++++++++++++++++++++++

WT

++++++++++++++++++++++++ _++++1++1 +++++++++ +++++++-_

WT

+++++++ +!+1+ + i++-_

3O

V]E

___&.D, OUT FRA,I_ c-_ ' + +-I--F + + + + + + + + + + + ++ + + + + + + + I

>

I + +_+J + + 4- 4-- + + -!- -I- ++++++++ ++++++++

+ +l+1++ +L÷j+

RB

r A-A

DRAWING NO, MARCHESSEAULT 3 OF 3 DWG BY SHEET F93-IC-I 70-2 NDTES_ ALL CDHPONENTS ARE ALCLAD 2024-T3 REHDVE ALL SHARP EDGES AND BURRS IIII' 15 908 RIVET MS20470DD-3 14 2060 R I VET MS2 047 ODD-4 13 6 Z STR INGER NAS346-3 12 3 ,063 IN RIB F93-IC-211 11 2 ,032 IN RIB F93-IC-21 0 1 0 1 ,02 IN RIB F93-IC-209 9 1 ,02 IN TIP RIB F93-IC-208 8 1 , 071 LEADING EDGE SKIN F93-IC-207 7 2 ,032 LEADING EDGE SKI N F93-I C-2 06 6 1 ,02 LEAD ING EDGE SKI N F93-1 C-2 05 5 2 ,071 IN SKIN F93-IC-204 4 4 ,032 IN SKIN F93-IC-203 3 2 , 02 IN SKIN F93-IC-202 2 1 REAR SPAR F93-IC-201 1 1 FRONT SPAR F93-IC-200 ITEM QTY BESCR I PT I[IN PART# EMBRY-RIDBLE AERONAUTICAL UNIVERSITY DIN_IIIN r_ UNL_ _ _CIFI_ DAYTONA BEACH FLORIDA SIZE DATE SCALE DRAWN BY

,XX + ,01

B 12/93 1/20 CARR

,XXX + ,001

TITLE LEFT WING ASSEMBLY DRAWING NO, SHEET

± i/2 °

F93-IC-160-2 1 OF 4

LBL

LBL

LBL 96

LINE FUSELAGE NBTES: TO BL56 IS ,071 IN BL27 THICKNESS Z_ SKIN TO BL96 IS ,032 IN BL56 THICKNESS z_ SKIN TO BLI89 IS ,OP IN BL96 THICKNESS z_SKIN "_t.,,,'r FRAM_ _._.

LBL

SHEET DATE BY IDRAWING NO, CARR F93-IC-160-2 I-2/93 8 DF 4 I 2,00 j

©

<

> Olo

LBL LIBL {- 3,00 I 12,00

@

Io

OiO0 LBL _- 5,06 0 0 0

©

l("

! o ol ool

LBL LBL

J_

I li I I 1,00 _ OlOOOOlOOO I< LBL LBL 189 DRAWING NO, SHEET IDATE I BY F93-IC-160-2 3 FIF 4 12/931 CARR

A

DETAIL

DETAIL A

I

0 0

0 d -0

liHI fI_.DOU_F_ _ - BRAWING NFI, BATE SHEET I BY F93-IC-160-2 12/93 CARR 4 OF

I

I I

I

i J

C I

I I \ I

!

ST#, 0 NFITE; I) ALL PARTS AL( 2) BREAK ALL El){ _.DOUI" FR ,_ _._.

4 AN4-5 4 BOLT - MACHINE, AIRCRAFT 3 4 BDLT AN8-5 - MACHINE, AIRCRAFT BULKHEAD 2 I PROP QQ-A-225/6 1 THRUST BEARING F93-IC-197-2 ITEM QTY _IATERIAL DR PART # DESRiPTIDN EMBRY-RIDDLE AERONAUTICAL UNIVERSITY UUl,.£S_ aI_RV_J_[ SPEC.IF_P DAYTONA BEACH FLORIDA SIZE DATE SCALE DRAWN BY

,XX ± ,01

B 12/3/93 i/_ CRAIG STEVENS

,XX/ __+,001

TITLE PRFIP BULKHEAD All DRAWING NO.

S & BURS SHEET F93-IC-158-2 IDFI WL "°m,_ '°_ X" I I /--BEAD UP P I WL n im WL

STA 130

WL WL

STA ;:'3

ALCL, NBTE_ i) ALL PARTS 2) BREAK ALL EDGE _ @lIIxlrI FBA_e- MC152-6 12 4 BF]LT - MACHINE, AIRCRAFT AN12-5 ii 4 BOLT - MACHINE, AIRCRAFT AN4-5 I0 I FIREWALL AMS 5604 9 1 LANDING GEAR ATTACHMENT QQ-A-225/6 8 28 RIVET MS20430A-4-6 7 2 C-CHANNEL QQ-A-200/3 6 2 C-CHANNEL QQ-A-200/3 5 18 RIVET MS20430A-4-7 4 2 C-CHANNEL QQ-A-200/3 3 i BEARING F93-IC-198-2 2 4 SEAT BELT ATTACHMENT QQ-A-250/5 1 i C-CHANNEL QQ-A-200/3 ITEM QTY DESCRIPTION MATERIAL OR PART # AERONAUTICAL UNIVERSITY EMBRY-RIDDLE DAYTONA BEAC_ FLORIDA SIZE ]DATE SCALE DRA"+/_'_BY

,XX + ,01

B 12/2/93 1/80 CRAIG STEVENS

,XXX -'--,001

TIILE BULKHEADS _i IF- r" T DRAWING NO.

_LL c & BURS F93-IC-157-2 1 OF 1 / HEATING VEN ELECTRICAL!

CONTROL SYS-

/

VACUUH HARN CROSS SECTIFIN STATION B8 • '_i_DOUT FF_c

RNESS

EMBRY-RIDI)LE AERONAUTICAL UNIVERSITY iiiiii_lsliml _

MS

DAYTONA BEACH FLORIDA

_S

,XX _+ ,01

SIZE |DATE _SCALEI/10 IDRAVN BY B___0/12/9 T,S, MCCE]RKLE

XXX 4-,001

TITLE ELECTRIC, HEATING, AND VACUUM CONCEPT DRAVING NO. ISHEET

± i/2 °

F93-IC-152 _I DF i ]- 6585 I_ JJ I |I %.

STA

/ ®

15,0

L--.5 -s N TE.SH ,R '

AND CFIRNERS SHFIL 8,75 4"1 BE RFIUNDED I]FF,

SECTION A-A

9 8

BEARING PLATE

8084-T3

AN3BO-5

8 8 NUT

BOLT AN5C4

7 8

6 80

BOLT AN6-1£

5 8

"C" BRACKET 8084-T3

4 8

END BEARING 8084-T3

3 1 JOURNAL BEARING BO84-T3

8 4 "C" CHANNEL 8084-T3

1 1 DRIVE SHAFT F93-C01-150

ITEM QNTY PART DESCRIPTION NUMBER BR MATERIAL EMBRY-RIDDLE AERONAUTICAL UNIVERSITY 9_4E.HSI'DIq "I'01.KI_M'_.ES I.INI..ESS OTI-ER',,/ISE SPEC_F_..D DAYTONA BEACH FLORIDA

IGES

DRAWN BY SIZE DATE SCALE

.D

,XX + ,01

B io\ia\D3 i/£ 0 T,S,M & G,S,

,XXX + ,001

TITLE DRIVE SHAFT\BEAM ASSEMBLY DRAWING NO,

+ i/2 °

I SHEET DF 1 F93-IC-151 dql..CmUT e,_tZ i,

STA

©

w

t_ ___

A A

STA

VIEW A-A

1/2 FULL SCALE

m l_=m -N I_=m

VIEW B-B

,,_, 1/2 FULL SCALE

STA I

BL

O

S S

6 i UNIVERSAL JOINT MSB0271

5 8 BOLT AN6-12

4 i CIR CLIP 2024-T3

3 i SHEAR P IN 2024- T3

2 1 SECFINDARY SHAFT 4130

1 1 PRIMARY SHAFT 4130

ITEM QNTY DESCRIPTIEIN PART # EMBRY-RIDDLE AERONAUTICAL UNIVERSITY DIMENSION I"DI.ER.4NCK_ UNLESS OTHERVISE SPECIFIED DAYTnNA BEACH FLORIDA DRAWN BY SIZE DATE SCALE

,XX + ,01

B

[2\3\93 1/i0 T,S, MCCBRKLE

XXX + ,001

TITLE

DRIVE SHAFT ASSEMBLY

DRAWING NO, SHEET

F93-IC-150

I DF i gD, ix_lr FIU_ NFITE: ALL CFIMPFINENTS ARE ALCLAD 2024-T3 UNLESS FITHERISE NOTED ALL SHARP EDGES AND BURRS REMOVED - _l_mouTr_ _>'>'>'>'>'>'>'>' 20 I NOSE ASSEMBLY F93-IC-123 19 30 TRUSSHEAD SCREW AN526C-6-32 1B 7 0.03 INCH SKIN F93-1C-I06 1500 RIVET MS20470DB-4 SEAT BELT MOUNTS F93-IC-122 JAARS SEAT ASSEMBLY F93-IC-121 15 2 DOOR FRAME ASSEMBLY F93-IC-120 14 2 F93-IC-119 13 16 .75 INCH ANGLE F93-1C-IIB FORWARD BULKHEAD F93-1C-117 NOSE BULKHEAD ii F93-1C-116 17-4PH STAINLESS STEEL FIREWALL I0 F93-1C-115 WINDSHEILD ASSEMBLY DOOR ASSEMBLY F93-IC-I14 B CABIN LIGHT ASSEMBLY F93-IC-I13 INSTRUMENT PANEL ASSEMBLY F93-1C-III 2 INCH LONGERON MOUNTS F93-IC-II0 4 7 2 INCH LONGERONS F93-IC-I08 3 4 112 INCH ANGLE F93-IC-I05 2 I DRIVE SHAFT BOX ASSEMBLY F93-IC-130 I i FLOOR ASSEMBLY F931-C-127 DESCRIPTION PART# ITEM QTY EMBRY-RIDDLE AERONAUTICAL UNIVERSITY DIMENSION TOLERANCES UNLESS aTHERVISE SPECIFIE9 I]AYTONA BEACH FLORIDA DRAWN BY DATE SCALE SIZE

,XX + ,01

CARR B 10/93 I/PO

XXX i ,001

TITLE CAB I N ASSEMBLY DRAWING NO, SHEET

_+ 1/2 °

F93-IC-129-2 OF 2

J

SECTION

SCALE

_F93-IC-116 (REF)

(

F93-IC-121

_++- _- ¢r_ ....

++ +4- ++ _ ++ ++ + + + ++,,_ + J +

F93-IC-I 1 0 (REF)

NI]TE:

VENT INTAKE

/_ ENG I NE

C-1 1 0 (REF)

EF)

++ ++ ++ ++

(REF)

F93-1C-1 19

SHEET DATE ]3Y I]]RAWING NO, 2 OF 2

12/93 CARRI F93- 1C- 129-2

NFITE_ ALL CDMPONENTS ARE ALCLAD 2 024-T3 ALL SHARP EDGES AND BURRS REMOVED

L

_JLDOUT FF_._

RBL I

I16

MS2047 ODD-4 9 RIVET AN526C-6-32 19 TRUSSHEAD SCREW F93-IC-I 07 NOSE ACCESS PANEL 7 1 F93-IC-I 06 0, 03 SK I N 6 3 F93-IC-I 05 5 8 i/2 IN ANGLE F93-IC-I 04 4 16 NOSE LDNGERDN MDUNT S F93-IC-I 03 NOSE LDNGERDN B 8 F93-IC-I 02 2 1 NOSE FRAME WT F93-IC-I Ol PROPELLER BULKHEAD 1 I PART# BESCRIPTIrlN ITEM QTY !

EMBRY-RIDDLE AERONAUTICAL UNIVERSITY 9T, J4EHS_]N TrlLIERANCE$ UNLESS D'I'HERWZSE SPECLC'IE9 BAYTrINA BEACH --LDRIDA ]g_ELU._ DATE SCALE DRAWN BY SIZE

,XX + ,01

10/93 1/20 CARR B

XXX + ,001

TITLE NOSE ASSEMBLY S-H-EET DRAWING NO, 1 DF 1 F93- IC-128-2 I F93-IC-I 18 (REF) NOTE: 2024-T3 ALL COMPONENTS ARE ALCLAB UNLESS OTHERWISE NOTED REMDVEB ALL SHARP EBGES AN]] BURRS UPPER FLOOR SKIN IS ,06 TH ICKNESS 93-IC-I 16 (REF) 9 8 1/2 INCH ANGLE F93-IC-I 05 8 200 RIVET MS2 0470DB-8 7 31 O0 RIVET MS20470BB-4 6 2 O, 06 INCH FLOOR SKI N F93- 1C- 126 5 8 SEAT SUPPORT RIB F93-IC-125 4 4 FLOOR RIB F93-IC-124 3 36 2 INCH L[]NGERON MOUNTS F93-IC-I 1 0 2 4 1 INCH LFINGER[]N F93- 1C- 1 09 I 18 2 INCH LDNGERON F93-IC-I 08 ITEM QTY DESCR IPT InN PART# EMBRY-RIBBLE AERONAUTICAL UNIVERSITY I_JNENSION UNLE_ OTHERVISE _ECI;rIE_ DAYTONA BEACH FLORIDA SIZE DATE SCALE DRAWN BY

,XX + ,01

1 0193 1/20 CARe B

;,XXX + ,001

TITLE FLOOR ASSEMBLY SHEET DRAWING NO,

_+ 1/2 ° i o? 1

F93-10-127-2 gDJ_ouT FP,_ L NOTE: ALL PARTS ALCLAD FINISHED BREAK ALL SHARP EBGES AND BL

<!)

1 C-CHANNEL NO, 14 2024 AL 2 C-CHANNEL NO, 10 2024 AL Q 2 ANGLES EQUAL LEG NAS341-14 IG RIVET ROUND HEAD MS20430A-6-8

®

61 RIVET ROUND HEAD MS20430A-2-4

®

NAS341-4 (_) 3 ANGLES EQUAL LEG INSTRUMENT PANEL FACE QQ-A-255

(2)

MATERIAL DR PART # QTY DESCRIPTION ITEM EMBRY-RIDDLE AERONAUTICAL UNIVERSITY _NSlDN TOLERANCES UNLESS nTHERVISE _ECIFIEO DAYTONA BEACH FLORIDA _CIMAL SIZE DATE SCALE DRAWN B_ >

,XX + ,01

B 10110/93 IIi0 R, CRAIG STEVENS

,XXX + ,001

TITLE INSTRUMENT PANEL STRUCTURE DRAWING NO, SHEET

+_ I/2 °

F93-IC-III IOF i