Fuselage Design

CHAPTER 5

Fuselage Design

5.1 VOLUME CONSIDERATIONS 5.2 AERODYNAMIC CONSIDERATIONS 5.3 DRAG ESTIMATION 5.4 SPREADSHEET FOR FUSELAGE DESIGN 5.5 PROBLEMS

Photograph of the Aero Spacelines B-377PG "Pregnant Guppy," which was designed to transport outsized cargo for the NASA Apollo Program. The upper fuselage was more than 20 fe et in diameter to accommodate portions of the Saturn V rocket. (NASA Dryden Research Center Photo Collection.)

Following the main wing, the next logical step in the conceptual design involves the design of the fuselage. The fuselage has a number of functions that vary depending on the type and mission of the aircraft. These include accommodating crew, passengers, baggage or other payload, as weil as possibly housing internai engines. Other considerations for the fuselage design include possible fuel storage, the structure for wing attachments and accommodations for retractable landing gear.

86

Section 5.1 Volume Considerations 87

Ail of these generally aim at setting the internai volume, height, width and length of the fuselage. For example, the primary mission of the "Pregnant Guppy" pictured on the previous page, was to carry 20-foot diameter sections of the Apollo Saturn V rocket. This specifically dictated the size and shape of the fuselage, and resulted it its distinctive appearance.

Depending on the Mach number regime, the optimum shape of the fuselage, or more specifically the length-to-diameter ratio, may be determined on the basis of mini-mizing the aerodynamic drag. With subsonic aircraft, this ratio is historically far from the optimum and dictated more by function. With supersonic aircraft, however, the penalty for deviating from the optimum leaves little room for compromise.

VOLUME CONSIDERATIONS

... .1 Passenger Requirements

The size and shape of subsonic commercial aircraft are generally determined by the number of passengers, seating arrangements and cargo requirements. Table 5.1 gives typical dimensions for the passenger compartments. These are generally based on an average passenger who is assumed to weigh 180 lbs. Figure 5.1 accompanies the table and provides graphical definitions of the nomenclature used.

Seating arrangements on commercial passenger aircraft vary depending on the size and range. Examples of different aisle seating arrangements for a variety of air-craft is illustrated in Figure 5.2. Table 5.2 lists additional seating information for these aircraft.

The seating arrangements in Figure 5.2 generally correspond to "coach class." In large commercial aircraft, there is additional variability in seating arrangements based on other passenger classes. This is illustrated in Table 5.3 with a comparison between coach and first class for two typical jet transports.

As indicated in Table 5.1 , passengers are allocated from 40-60 lbs of baggage. On average, this is expected to occupy a volume of 15-25 f3 . In addition, passengers are allocated 3 f3 for onboard overhead baggage storage.

In modem aircraft, checked baggage and other cargo are carried in standard cargo containers. Examples of these are shawn in Figure 5.3.

TABLE 5.1: Passenger compartment requirements.

Long-Range Short-Range

Seat Width (in) 17-28 16-18 Seat Pitch (in) 34-40 30-32 Headroom (in) >65 Aisle Width (in) 20-28 >15 Aisle Height (in) >76 >60 Passengers/Cabin 10-36 ~50 Lavatories/Passenger 1/(10-20) 1/(40-50) Galley Volume/Passenger (f3) 1-8 0-1 Baggage/Passenger (lbs) 40-60 40

88 Chapter 5 Fuselage Design

~ --- ----------- 1 1 1 1 1 ' 1 1 1 1 1 1 1 1 1

: Seat Pitch : 1 1

1 1 1 1

FIGURE 5.1: Schematic drawing of a passenger seating arrangement defining parameters.

TABLE 5.2: Passenger aircraft seating arrangements.

Passenger No. Fuselage Aisle Seating Ex amples Diam. (in.)

4-9 64 1 + 1 Citation V 10-20 58 1 + 1 Beech 1900

94 2+1 Gulfstream II 20-50 91 2+1 Saab 340 50-75 106 2+2 DHC-8/300

75-190 130 2+3 MD-80 148 3+3 Boeing 757

190-270 198 2+3+2 Boeing 767 222 2+4+2 Airbus A300

270-360 222 2+4+2 Airbus A330 236 2+5+2 DC-10, L-1011, Boeing 777

360-450 256 3+4+3 Boeing 747

Sect ion 5.1 Volume Considerations 89

Boeing 767 AirbusA300

BB Lockheed L- 1011 McDonnell Douglas DC-10

BB BB Boeing777 Boeing 747

FIGURE 5.2: Schematic drawings of coach compartment cross-sections that are typical of different commercial passenger aircraft. Also shown are types of lower deck cargo containers used on these aircraft.


LD-2,LD- 3

H

lJ D~

1 LD-4,LD-5 H

l D~

LD-8

FIGURE 5.3: Schematic drawing of different styles of lower deck containers listed i Table 5.4


TABLE 5.3: Typical passenger accommodations for large jet transports.

B757 DC-10

Seats Total 178 292 First Class 16 (9%) 24 (8.2%) Coach 162 (91 %) 268 (91.8%) First-Class Pitch 38 in. 38 in. Coach Pitch 34 in. 34 in. First-Class Aisle 2+2 3+3 Coach Aisle 3+3 2+5+2

Lavatories First Class 1 1 Coach 3 6 First-Class Pass./Lav. 16 24 Coach Pass./Lav. 54 45

Galleys First-Class Volume 70 f3 120 f3 First-Class f3/Pass. 4.4 5.0 Coach Volume 231 f3 450 f3 Coach f3/Pass. 1.4 1.7

TABLE 5.4: Cargo container dimensions.

Type Height (in.) Width (in.) Depth (in.) Volume (f3) Gross Wt. (lbs) LD-2* 64.0 61.5 60.4 120 2700 LD-3* 64.0 79.0 60.4 156 3500 LD-4 64.0 96.0 60.4 195 5400 LD-5 64.0 125.0 60.4 279 5400 LD-8* 64.0 125.0 60.4 245 5400

* Trapezodal shape.

Table 5.4 lists the dimensions of the more widely used containers. The smaller containers are suitable for smaller commercial/transport aircraft, such as the Boeing 727. The larger ones, for example the LD-3, are commonly used on larger commercial aircraft, such as shown in Figure 5.2. Table 5.5 lists the number of LD-3 containers that can be carried on these transports.

Smaller, short-range aircraft do not use cargo containers, but rather have space only for bulk cargo with a volume that is based on 6-8 f3 per passenger.

Passenger aircraft also have requirements for the number, placement and type of emergency exits in the event of a survivable accident. These are based on the number of passenger seats installed on the aircraft. The requirements are summarized in Table 5.6, with the description of the exit types given in Table 5.7. They indicate a fairly straight-forward criteria for passenger numbers up to 179. After which, the number and type of exits is based on the arrangement that gives sufficient "seat credits." These credits need


TABLE 5.5: Large-body aircraft cargo compartment arrangements.

Number of Bulk Cargo LD-3 Containers Volume (f3)

B-747 30 1000 L-1011 16 700 DC-10 14 800 A-300 10 600

TABLE 5.6: Number and type of emergency exits required for passeng' transport aircraft by FAR 25.807.

No. Pass. Type I Type II Type ill

1-9 10-19 20-39 40- 79 80-109

110-139 140-179

2 2

1 1 1 2

2

Type IV

180-299 Add exits so that 179 plus "seat credits" ~ passenger number.

~ 300

Seat Credit

12 15 35 40 45

110

Exit Type

Single Ventral Single Tailcone

Pair Type III Pair Type II Pair Type I Pair Type A

Use pairs of Type A or Type I with the sum of "seat credits" ~ passenger number.

to correspond to the number of passenger seats in excess of 179 on the airera procedure holds for up to 299 passenger seats. Above this number, the designs u of Type A or Type I exits with the sum of the credits equal to the passenger se(

5.1.2 Crew Requirements The size of the crew compartment will vary depending on the aircraft. With lon military/commercial transport and passenger aircraft, the crew compartment sb designed to accommodate from two to four crew members. Recommendations


TABLE 5.7: Types of emergency exits for passenger transport aircraft defined by FAR 25.807.

Type Location

Type I Floor Level Type II Floor Leve]

Overwing Type ill Overwing Type IV Overwing Tailcone Aft of Pressure Hull Ventral Bottom of Fuselage Type A Floor Level

Min. Dimensions Width x Height (in.)

24 x 48 20 x 44 20 x 44 20 x 36 19 x 26 20 x 60

Equiv. Type I 42 x 72

Min. Step Height Inside:Outside (in.)

10:17 24:27 29:36 24:27

that the crew compartment have a length of approximately 150 inches for four crew members, 130 inches for three crew members and 100 inches for two crew members.

An important factor that impacts the shape of the forward section of the fuselage is the requirement that the pilot have an unobstructed forward view. A critical need in achieving this is obtaining the proper amount of over-nose angle. This is especially important for the landing phase for ali aircraft, and during the combat phase of military-fighter aircraft. The Concorde and Russian Tu-144 supersonic passenger jets have a nose section that deftects downward in order to give the necessary over-nose angle for landing. Figure 5.4 shows a photograph of the Tu-144 with the nose deftected during landing.

The over-nose angle, a 0 vernose, is defined as the angle between a horizontal line through the pilot' s eye, down to the point of the highest visual obstruction. A schematic representation is shown in Figure 5.5. The proper over-nose angle depends on the landing approach angle, Yapproach. and the landing approach velocity, Vso. ln the landing analysis in Chapter 8, the approach angle is found from

. -1 (-D) Yapproach = sm W , (5.1) where D is the drag and W is the weight at landing. The approach velocity, Vso, refers to the velocity at an elevation of 50 f, which starts the landing phase with

Vso = 1. 3 V.~ , (5.2)

and V5 is the stall velocity that will include an enhanced lift configuration of the main wing, which is covered in Chapter 9. If these quantities are known, a reasonable empirical relation for the over-nose angle is

C:Xovcmosc = Yapproach + 0.07Vso , (5.3)

where V5o has units of knots, and a and y have units of degrees. Table 5.8 gives sorne typical values for a variety of aircraft.

A somewhat Jess critica1 visual requirement for the crew compartment is the over side vision angle, a 0 verside This is the unobstructed viewing angle from a horizontalline


FIGURE 5.4: Photograph of Tu-144 with nose defl.ected to give necessary over-nose angle during landing. (NASA Dryden Research Center Photo Collection.)

~ Pilot's Eye --~----\-

Grazing Angle

SideView

Pilot's Eye

Front View

FIGURE 5.5: Illustration defining crew compartment vision parameters.


TABLE 5.8: Values of over-nose and over-side angles for different aircraft.

ovem ose ovcrside

Military Transports/Bombers 17 35 Military Fighter 11 -15 40 General Aviation 5-10 35 Commercial Transport 11 -20 35

through the pilot' s eye down to the highest visu al obstruction formed by the si de of the fuselage. Typical values for a variety of aircraft are also given in Table 5.8.

The upward vision angle is also important. Military/commercial transport and pas-senger aircraft should have an unobstructed viewing angle of at least 20 above the horizon. Military combat aircraft should have at least a 120 unobstructed upward view-ing angle.

Another factor that impacts pilot vision, which is mostly only an issue with combat and high-speed aircraft, is the transparency grazing angle. (See Figure 5.5.) This corre-sponds to the smallest angle between a line of vision of the pilot and the cockpit window or windscreen. If this angle becomes too small, the visibility through the window can become substantially reduced or distorted. The recommended minimum grazing angle is 30 .

. 1.3 Fuel Storage Requirements In long-range aircraft, a large percentage of the weight at take-off is due to the weight of the fuel. The volume required to hold this fuel can be allocated to the fuselage or wing, or most likely both. The decision on where to store the fuel depends on a number of factors. These include the location of the center of mass with respect to the center of lift, thus affecting the static stability, the vulnerability of crew and passengers in the event of an uncontrolled landing, and the vulnerability of the fuel in combat aircraft caused by enemy tire.

In the case of the static stability, the placement of the fuel is also important in defining how the location of the center of mass might change as a result of the changing fuel weight from the beginning to the end of cruise. In order to maintain static stability in the pitch direction, the center of mass must al ways be forward of the center of lift. As a result, if any fuel is stored in the fuselage, it should be located at or slightly forward of the wing attachment point. A detailed analysis of the static stability will be done later in Chapter 11.

The volume needed to accommodate the fuel is based on the maximum weight of fuel at take-off and the density of the fuel. The specifie volumes of various fuels are given in Table 5.9.

The actual volume that is available in locations in the fuselage or wing depends on the type of fuel tank used. There are generally three types: discrete, bladder and integral. The choice of these determines what percentage of the available volume is capable of holding fuel.


TABLE 5.9: Specifie volumes for different aviation fuels (f3 /lb).

oop 100F Mil-spec

AV-gas 0.0219 0.0235 0.0223 JP-4 0.0199 0.0208 0.0206 JP-5 0.0186 0.0196 0.0197 JP-8 0.0199

Discrete tank are generally only used for small general aviation aircraft. They consist of separately manufactured containers that mount in the aircraft. In the wing, these are often mounted at the inboard span portion, near the leading edge. In the fuselage, they are generally placed just behind the engine and above the pilots feet.

Bladder fuel tanks are thick rubber bags that are placed into cavities in the wing or fuselage. The advantage of bladder tanks is that they can be made to be self-sealing. This feature improves the aircraft survivability in the event of an uncontrolled landing or enemy f1re. The thickness of the rubber bladder walls reduces the available volume of the cavity. As a general rule, 77 percent of a cavity volume in the wing, and 83 percent of a cavity volume in the fuselage, is available with bladder tanks.

Integral tanks are cavities within the airframe structure that are sealed to form fuel tanks. Examples are the wing box areas forrned between wing spars and the area between bulkheads in the fuselage. Becau e integral tanks are more prone to leaking compared to the other two types, they should not be located near air inlet ducts or engines. The tire hazard of integral tanks can be reduced by filling the tank with a porous foam material. This, however, reduces the volume capacity by approximately 5 percent. As a general rule, 85 percent of the volume measured to the external skin of the wing, and 92 percent measured to the external skin of the fuselage, is available with integral tanks.

5.1.4 Internai Engines and Air lnlets Engines may be mounted internai to the fuselage. This is often the practice in combat aircraft and mali general aviation aircraft, but sometimes has been done with long-range commercial passenger aircraft, su ch as the B-727 and L-1 011. At this stage of the design, the total drag is not yet known so that the thrust requirements of the engine (size and number) are not yet determined. This will be covered in Chapter 7. As a result, the internai arrangement of the engine and air delivery system is difficult to a sess at thi point in the conceptual design. Therefore, if the propulsion system is likely to be internai to the fuselage, the volume to enclose it needs to be accounted for in the design. At this stage, the best approach i to rely on suitable comparison aircraft.

For internally mounted jet engines, the air delivery system is an integral element. The type and geometry of the inlet wil1 determine the pressure loss and uniformity of the air supplied to the engine. These in turn affect the installed thrust of the engine and fuel consomption.

The types of air inlets depend on the operating Mach number. In general, the objective of the air inlet system for turbojet and turbofan engines is to reduce the Mach


number of the air at the compressor face to between 0.4 and 0.5. In subsonic aircraft, this is accomplished using a subsonic diffuser. In supersonic aircraft, this is done through area changes at the inlet that result in the formation of one or more compressive shocks.

At this stage of the design, the objective is to size the fuselage. Therefore in addition to basing the design on suitable comparison aircraft, a first estimate can come from empirical data that indicate that the diameter of the air inlet be the same as that of the engine compressor face and that the length of the inlet be 60 percent of the engine length .

. 1.5 Wing Attachments The manner in which the main wing attaches to the fuselage is an important element in the fuselage design. For structural reasons, the wing is generally constructed as an integral unit. The portion that passes through the fuselage is referred to as the wing carry-through. The root-span portion of the wing has the largest thickness in order to withstand the large bending moment in the wing. As a result, the wing carry-through occupies a large volume where it passes through the fuselage. An illustration of the wing carry-through is shown in Figure 5.6. A photograph of the partially assembled fuselage section with the main wing attachment for the Boeing 777 is shown in Figure 5.7.

Since the details of the main wing are known from the analysis of the preceding chapter, the volume requirements for the carry-through structure can be directly applied to the design of the fuselage.

- .1.6 Landing Gear Placement In most aircraft, the fuselage needs to accommodate ali or sorne parts of the landing gear when it is retracted. Therefore, its placement and volume requirements need to be considered in the design of the fuselage.

The size and location of the landing gear will vary depending on the aircraft. Again, a good first estimate can come by examining suitable comparison aircraft. In selecting comparisons for the basis of the landing gear size and placement, it is important to select aircraft that have a comparable take-off weight. Figure 5.8 shows an illustration of sorne of different landing gear arrangements.

FIGURE 5.6: Illustration of a typical fuselage wing carry-through arrangement.


FIGURE 5.7: Photograph of a Boeing 777 fuselage wing attachment section during assem-bly. (Courtesy of the Boeing Company.)

Most commercial passenger/transport and combat aircraft use retractable tricycle-type landing gear. With these, the nose wheel is mounted on and retracts into the fuselage. The main wheels are usually mounted on the main wing. Depending on their size, they retract into the main wing, or into the main wing/fuselage junction.

Heavier aircraft use a multi-bogeyed tricycle arrangement that have multiple sets of wheels at three points. Two-wheel bogeys are typically used on aircraft with take-off weights in the range of 50,000 to 150,000 lbs. Four-wheel bogeys are generally used on aircraft with higher take-off weights between 200,000 to 400,000 lbs.

A fewer number of military long-range aircraft such as the B-52, use a quadricycle landing gear arrangement in which the wheels are on either side of and retract into the fuselage. An extreme variation of this is used on the Russian built An-225, which has a take-off weight of 1.3 million pounds and uses 14 two-wheel bogeys on either side of the fuselage.

Another variation on the quadricycle landing gear arrangement can be seen on the MD-Il. This uses a fourth set of wheels, which are located further aft, on the fuse-lage centerline. These are revealed in .the photograph of an MD-11 during take-off in Figure 5.9

The largest portion of the landing gear for which space has to be allotted in the fuselage is the landing gear wheels. Here wheels refer to the hub and tire. The size of the wheels is proportional to the percentage of the aircraft weight that they hold. Most typically, the tires on the main landing gear carry approximately 90 percent of the aircraft weight. The other 10 percent is carried by the nose gear.


Single PersonaU

Utility

Dual B-727 B-737 DC- 9 MD-80

Dual '!Win DR Trident

Tandom C-130

DualTandom B-707 DC-10 DC-8 L-1011 B-747

Dual Twin Tandom B-58

Triple SR-71

Triple Dual Tandom B-777 Tu-144

Twin Tricycle C-S A

FIGURE 5.8: Illustration of different main landing gear footprints.

At this point in the design, the size of the main landing gear wheels can be estimated using a statistical fit of historie data. This gives the diameter and width of the wheels, with units of inches, as

Main Whee1 Diameter or Width (in.)= A W!ain (5.4) with A and B as given in Table 5.10. The weight on each wheel of the main landing gear depends on the number of wheels, so that

0.9WTo Wmain = ---

Nwhccls (5.5)

After determining the size of the wheels on the main gear, the size of the nose whee1 can be assumed to be approximately 40 percent smaller. An exception is for the quadricycle arrangement, where the nose wheel is typically the same size as the main wheels. The calculated values for the diameter and width of all the wheels should be increased by 30 percent if the aircraft is intended to be operated on unpaved runways.

5.1. 7 Arma ment Placement With combat aircraft, the number and size of bombs and armament are generally decided in the initial design proposai when the mission requirements are set. At this stage, when

Fuselage Design

Documents

Transcript of Fuselage Design