Category AIRCRAF DESIGN

The overall fuselage length, L (see Figure 3.49) consists of the (1) nose cone, (2) constant cross-section midsection barrel, and (3) aft-end closure. The constant cross­section mid-fuselage length is established from the passenger seating arrangement and combined with the class arrangement (i. e., first class, business class, and econ – omy/tourist class). Section 4.7.6 provides seat dimensions for the two main classes (i. e., business and economy).

Aircraft length may not be equal to fuselage length if any other part of the air­craft extends beyond the fuselage extremities (e. g., the tail sweep may go beyond the tail cone of the fuselage; see Figure 6.8). Figure 4.14 shows the fuselage geometry relationship to the number of passengers. The fuselage width increases in incre­ments with the number of passenger-abreast seating, one seat width at a time. Because of passenger comfort, a designer selects options from the sensitivity study (i. e., drag and cost variations); the continuous line in Figure 4.14 represents a typical average value. The actual width is determined in Chapter 6.

Fuselage geometry is determined from the designed passenger capacity (see Chap­ter 6). There are two parameters to size (i. e., fuselage width [W] and fuselage length [Lf]), which determine the constant-section fuselage-barrel length. In turn, this depends on the seat pitch and width for the desired passenger comfort level. Table 4.2 lists the statistics for existing designs – a new design would be similar. The width and length of the fuselage must be determined simultaneously, bearing in mind that the maximum growth potential in the family of variants cannot be too long or too short and keeping the fineness ratio from 7 to 14 (a good value is around 10). Boeing 757-300 records the highest fineness ratio of 14.7. A seating arrange­ment with two aisles results in more than six abreast (average diameter, Dave = [H + W]/2; see Figure 4.14).

4.7.1 Fuselage Width

The first parameter to determine for the fuselage average diameter is the num­ber of abreast seating for passenger capacity. There is an overlap on choice for

the midrange capacity in the family of design; for example, an A330 with 240 to 280 passengers has seven-abreast seating whereas the same passenger capacity in a B767 has eight-abreast seating. When seating number is increased to more than six abreast, the number of aisles is increased to two to alleviate congestion in pas­senger movement. Because of the current fuselage-length limitation of 80 m, larger – capacity aircraft have a double-deck arrangement (e. g., the B747 and the A380). It would be interesting to try a two-aisle arrangement with six-abreast seating that would eliminate a middle seat. A three-aisle arrangement with ten-abreast seating would eliminate the cluster of four seats together. A BWB would have more than two aisles; there is no reason to not consider a triple-deck arrangement.

Although a circular cross-section is the most desirable relative to stress (min­imize weight) and manufacture (minimize cost), the market requirements for the below-cabin floorspace arrangement could result in a cross-section elongated to an oval or elliptical shape. The Boeing 747 with a more narrow upper-deck width is a unique oval shape in the partial length that it extends. This partial length of the upper deck helps cross-sectional area distribution (see Section 3.23) and area ruling.

Figures 4.12 and 4.13 show various options for aircraft fuselage cross-sections to accommodate different seating arrangements. All fuselage cross-sections are sym­metrical to the vertical plane. In general, aircraft with four-abreast seating and more have space below the cabin floor for baggage and cargo.

coroner space

(floor recessed)

vung partially through fuselage

Unpressurized propeller-driven aircraft operating at lower altitudes can have rectangular cross-sections to reduce manufacturing costs, as well as offer more space (e. g., Shorts 360 aircraft). A pressurized fuselage cross-section would invariably be circular or nearly circular to minimize weight from the point of hoop-stress consid­erations. A two-abreast circular cross-section would have cramped legroom; a better option is a slightly widened lower lobe (e. g., Learjet 45) to accommodate legroom. In general, with a three-passenger capacity and more, the midsection fuselage has a constant cross-section with front and aft ends tailored to suit the requirements. The wing box arrangement for smaller aircraft should pass over (e. g., high-wing DO328) (Figure 4.13) or under (e. g., Learjet 45) the fuselage.

Previous sections discussed statistical relationships of weight and geometries for a complete aircraft. Section 2.4.1 provides familiarization with typical civil aircraft and its components. The next level of information pertains to the aircraft component geometries available, as building blocks, to shape a new aircraft. There is a wide range of options available from which to choose. The choices are not arbitrary – definite reasons are associated with the choices made (see Chapter 6). This sec­tion provides pertinent information on the fuselage, wing, empennage, and nacelle groups, which are required to configure civil aircraft designs.

1. Fuselage Group. This is concerned with shaping and sizing of the fuselage, from where the civil aircraft configuration exercise begins. Related information ascertains seating arrangement, comfort level, and cabin width to accommodate passenger loading so that the longest in an aircraft family does not exceed the fineness ratio on the order of 13. The appropriate front and aft-end closure choices are then made. When the fuselage shell is established, the next task is to configure the interior for passenger and crew requirements. The flight – crew space in the forward closure (i. e., cockpit) and the pilot vision polar are then established. Inside the cabin, the crew and passenger requirements are approached simultaneously as integral requirements (e. g., seating, toilets, and galleys).

2. Wing Group. This is the most important component of the aircraft. The plan – form shape must be established and then sized for operational-field and flight – performance requirements. Options for high-lift devices are described in Sec­tion 3.12. Other smaller components (e. g., winglets) also are considered (see Section 3.21) but not all aircraft incorporate winglets.

3. Empennage Group. Choice, size, and placement result from the aircraft’s CG position and wing size. This book adheres to the conventional H-tail and V-tail configuration.

4. Nacelle Group. This topic is addressed in Chapter 10; only an outline for the shaping choice is provided herein.

These four groups of aircraft components provide the preliminary shaping of candidate aircraft configurations. After the wing-sizing and engine-matching exer­cises, the choice must be narrowed to one final configuration that offers the best compromise for the family variants to cover a wide market. The undercarriage is addressed separately in Chapter 7.

Iterations are required to position the empennage and undercarriage with respect to the wing because the CG position initially is not known. Weights are esti­mated from a provisional positioning and then the positions are fine tuned through iterations. (In a classroom exercise, one iteration is sufficient.)

Once the wing area is established along with fuselage length and matched engine size, the empennage areas (i. e., H-tail, SH, and V-tail, SV) can be estimated from the static stability requirements. Section 3.22 discusses the empennage tail-volume coefficients to determine empennage areas.

Figure 4.10 shows growth for H-tail and V-tail surface areas with the MTOM. The variants in the families do not show change in empennage areas to maintain component commonality.

Figure 4.11 substantiates Equation 3.43 in Section 3.20.1, which states that the growth of the wing span is associated with the growth in wing loading.

With steady improvements in new-material properties, miniaturization of equipment, and better fuel economy, wing span is increasing with the introduction of bigger aircraft (e. g., Airbus 380). Growth in size results in a wing root thickness large enough to encompass the fuselage depth when a BWB configuration becomes an attractive proposition for large-capacity aircraft. Although technically feasi­ble, it awaits market readiness, especially from the ground-handling perspective at airports.

The aspect ratio shows a scattering trend. In the same wing-span class, the aspect ratio could be increased with advanced technology but it is restricted by the increase in wing load. Current technology provides for an aspect ratio from 8 to 14.

The relationships between engine sizes and the MTOM are shown in Figure 4.9. Turbofan engine size is expressed as sea-level static thrust (TSLS) in the ISA day at takeoff ratings, when the engine produces maximum thrust (see Chapter 10). These graphs can be used only for preliminary sizing; formal sizing and engine matching are described in Chapter 11.

Thrust-loading (T/W), is defined as the ratio of total thrust (TSLSJot) of all engines to the weight of the aircraft. Again, a clear relationship can be established through regression analysis. Mandatory airworthiness regulations require that multiengine aircraft should be able to climb in a specified gradient (see FAA requirements in Chapter 13) with one engine inoperative. For a twin-engine aircraft,

Figure 4.10. Empennage area versus wing area

failure of an engine amounts to a 50% loss of power, whereas for a four-engine air­craft, it amounts to a 25% loss of power. Therefore, the T/W for a two-engine aircraft would be higher than for a four-engine aircraft.

The constraints for engine matching are that it should simultaneously satisfy sufficient takeoff thrust to meet the (1) field length specifications, (2) initial climb requirements, and (3) initial high-speed cruise requirements from market specifica­tions. An increase in engine thrust with aircraft mass is obvious for meeting takeoff performance. Engine matching depends on wing size, number of engines, and type of high-lift device used. Propeller-driven aircraft are rated in power P in kw (hp or shp), which in turn provides the thrust. Turboprops are rated in power loading, P/W, instead of T/W.

Smaller aircraft operate in smaller airfields and are generally configured with two engines and simpler flap types to keep costs down. Figure 4.9a shows thrust growth with size for small aircraft. Here, thrust-loading is from 0.35 to 0.45. Fig­ure 4.9b shows midrange statistics, mostly for two-engine aircraft. Midrange aircraft operate in better and longer airfields than smaller aircraft; hence, the thrust-loading range is at a lower value, between 0.3 and 0.37. Figure 4.9c shows long-range statis­tics, with some two – and four-engine aircraft – the three-engine configuration is not currently in use. Long-range aircraft with superior high-lift devices and long run­ways ensure that thrust-loading can be maintained between 0.22 and 0.33; the lower values are for four-engine aircraft. Trends in family variants in each of the three classes are also shown in Figure 4.9.

Whereas the fuselage size is determined from the specified passenger capacity, the wing must be sized to meet performance constraints through a matched engine (see Chapter 11). Figure 4.8 shows the relationships between the wing planform refer­ence area, SW, and the wing-loading versus the MTOM. These graphs are useful for obtaining a starting value (i. e., preliminary sizing) for a new aircraft design that would be refined through the sizing analysis.

Wing-loading, W/Sw, is defined as the ratio of the MTOM to the wing planform reference area. (W/SW = MTOM/wing area, kg/m2, if expressed in terms of weight; then, the unit becomes N/m2 or lb/ft2.) This is a significant sizing parameter and has an important role in aircraft design.

The influence of wing-loading is illustrated in the graphs in Figure 4.8. The ten­dency is to have lower wing-loading for smaller aircraft and higher wing-loading for larger aircraft operating at high-subsonic speed. High wing-loading requires the assistance of better high-lift devices to operate at low speed; better high-lift devices are heavier and more expensive.

The growth of the wing area with aircraft mass is necessary to sustain flight. A large wing planform area is required for better low-speed field performance, which exceeds the cruise requirement. Therefore, wing-sizing (see Chapter 11) provides the minimum wing planform area to satisfy simultaneously both the takeoff and the cruise requirements. Determination of wing-loading is a result of the wing-sizing exercise.

Smaller aircraft operate in smaller airfields and, to keep the weight and cost down, simpler types of high-lift devices are used. This results in lower wing-loading (i. e., 200 to 500 kg/m2), as shown in Figure 4.8a. Aircraft with a range of more than 3,000 nm need more efficient high-lift devices. It was shown previously that aircraft size increases with increases in range, resulting in wing-loading increases (i. e., from 400 to 700 kg/m2 for midrange aircraft) when better high-lift devices are considered.

Here, the trends for variants in the family of aircraft design can be examined. The Airbus 320 baseline aircraft is in the middle of the family. The A320 family retains the wing to maintain component commonality, which substantially reduces manufacturing cost because not many new modifications are necessary for the vari­ants. This resulted in large changes in wing-loading: The smallest in the family

MTOM (kg)

(c) Large twin-aisle aircraft

Figure 4.8. Wing area, , versus MTOM (A318) has low wing-loading with excellent field performance, and the largest in the family (A321) has high wing-loading that requires higher thrust-loading to keep field performance from degrading below the requirements. Conversely, the Boeing 737 baseline aircraft started with the smallest in the family and was forced into wing growth with increases in weight and cost; this keeps changes in wing-loading at a moderate level.

Larger aircraft have longer ranges; therefore, wing-loading is higher to keep the wing area low, thereby decreasing drag. For large twin-aisle, subsonic jet aircraft (see Figure 4.8c), the picture is similar to the midrange-sized, single-aisle aircraft but with higher wing-loadings (i. e., 500 to 900 kg/m2) to keep wing size relatively small (which counters the square-cube law discussed in Section 3.20.1). Large air­craft require advanced high-lift devices and longer runways.

Figure 4.7 shows the relationship between fuel load, Mf, and the MTOM for twenty turbofan aircraft; this graph provides the fuel fraction, Mf/MTOM. It may be

examined in conjunction with Figures 4.4, 4.5, and 4.6, which show the range increase with the MTOM increase.

Fuel mass increases with aircraft size, reflecting today’s market demand for longer ranges. The long-range aircraft fuel load, including reserves, is less than half the MTOM. For the same passenger capacity, there is statistical dispersion at the low end. This indicates that for aircraft with a wider selection of comfort levels and choice of aerodynamic devices, the fuel content is determined by the varied market demand: from short ranges of around 1,400 nm to cross-country ranges of around 2,500 nm. At the higher end, the selection narrows, showing a linear trend. Figure

4.3 indicates that larger aircraft have better structural efficiency, offering a better OEMF; Figure 4.7 indicates that they also have a higher fuel fraction for longer ranges.

Figure 4.6. OEM versus MTOM

250,000

50,000

0 100,000 200,000 300,000 400,000 500,000 600,0

MTOM (kg)

decreasing). In the midrange (i. e., 70- to 200-passenger class – single-aisle, narrow – body), the OEMF is around 0.56. At the higher end (i. e., more than 200 passen­gers – double-aisle, wide-body), it is leveling out at around 0.483; the MTOM is slightly more than twice the OEM. The decreasing trend of the weight fraction is due to better structural efficiencies achieved with larger geometries, the use of lighter material, and the more accurate design and manufacturing methods of more recent designs.

The OEM is a function of aircraft load experienced on both the ground and in the air, which depends on the MTOM. The load in the air is a result of aircraft speed-altitude capabilities, the maneuverability limit, and wind. A higher speed capability would increase the OEMF to retain structural integrity; however, the OEM would reflect the range capability for the design payload at the MTOM (see Figure 4.5). Payload and fuel load can be exchanged to reach the MTOM from the OEM.

Figure 4.6a is represented in higher resolution when it is plotted separately, as shown in Figure 4.6b for midrange-size aircraft. It also provides insight to the statis­tical relationship between the derivative aircraft of the Boeing 737 and Airbus 320 families. The approaches of the two companies are different. Boeing, which pio­neered the idea, had to learn the approach to the family concept of design. The Boeing 737-100 was the baseline design, the smallest in the family. Its growth required corresponding growth in other aerostructures yet maintaining component commonality as much as possible. Conversely, Airbus learned from the Boeing experience: Their baseline aircraft was the A320, in the middle of the family. The elongated version became the A321 by plugging in constant cross-section fuselage sections in the front and aft of the wing, while retaining all other aerostructures. In the shortened versions, the A319 came before the even shorter A318, maintain­ing the philosophy of retaining component commonalities. The variants were not the optimized size, but they were substantially less costly, decreasing the DOC and providing a competitive edge.

Figure 4.5 describes the relationship between passenger capacity and MTOM, which also depends on the mission range for carrying more fuel for longer ranges. In con­junction with Figure 4.4, it shows that lower-capacity aircraft generally have lower ranges (Figure 4.5a) and higher-capacity aircraft are intended for higher ranges (Figure 4.5b). Understandably, at lower ranges, the effect of fuel mass on MTOM is not shown as strongly as for longer ranges that require large amounts of fuel. There is no evidence of the square-cube law, as discussed in Section 3.20.1. It is possible for the aircraft size to grow, provided the supporting infrastructure is sufficient.

Figure 4.5 shows an excellent regression of the statistical data. It is unlikely that this trend will be much different in the near future. Considerable scientific break­throughs will be required to move from the existing pattern to better values. Light but economically viable material, superior engine fuel economy, and miniaturization of systems architecture are some of the areas in which substantial weight reduction is possible.

In conjunction with Figure 4.4, it can be seen that longer-range aircraft gener­ally have higher MTOM; estimates of MTOM per passenger are provided herein (Table 4.1). At the start of a conceptual study, the MTOM must be guessed – these statistics provide a reasonable estimate. Below 2,500 nm, the accuracy degenerates; the weight for in-between ranges is interpolated.

EXAMPLE: For a mission profile with 300 passengers and a 5,000-nm range, the MTOM is estimated at 750 x 300 = 225,000 kg (comparable to the Airbus 300-300).

This section examines the statistics of current aircraft geometry and weight to iden­tify aircraft sizing parameters. Regression analyses are carried out to demonstrate a pattern as proof of expectations. With available statistics, aircraft can be roughly sized to meet specifications. This is the starting point; Chapter 11 discusses formal sizing to finalizing aircraft configuration.

(a) Lower capacity (b) Higher capacity

Figure 4.5. Number of passengers versus MTOM

Definitions of various types of aircraft mass (i. e., weight) are provided in Chapter 8; some are required in this section, as follows (payload could be passengers and/or cargo):

MEM: manufacturer’s empty mass – the finished aircraft mass rolls out from

the factory line

OEM: operator’s empty mass = MEM + crew + consumable – it is now

ready for operation

MTOM: maximum takeoff mass = OEM + payload + fuel – loaded to maxi­mum design mass

MEM is the design outcome from catering to the MTOM, in which fuel load and payload are traded. The trade-off between payload (i. e., passenger) and fuel is at the operator’s discretion, who has the choice to trade between them (see Chap­ter 13). Keeping the MTOM invariant, the operator can offload some passengers to increase the fuel load to the extent that the tankage capacity would allow a farther flying distance. Conversely, fuel could be offloaded to a shorter range, allowing an increase in passengers to the extent the aircraft can accommodate. Mass per pas­senger is revised to 100 kg (220 lb) from the earlier value of 90 kg (200 lb), which includes baggage allowance; there could be additional cargo load.