7.3.1 Civil Aircraft Economic Considerations

Commercial aircraft operation is singularly driven by revenue earned from the fare‐paying passengers and cargo. In the operating sector, passenger load factor (LF_pax) is defined as the ratio of occupied seats to available seats, as follows.

Typically, for aircraft of medium sizes and larger, operational costs break even at approximately one‐third full capacity (this varies among airlines; fuel costs at 2000 level fuel price with regular fares) – that is, a load factor of about 0.33. Of course, the empty seats could be filled with deregulated reduced fares, thereby contributing to the revenue earned. Until the 1960s, passenger fares were fixed under government regulation. Since the 1970s, the airfare structure has been deregulated – an airline can determine its own airfare and vary as the market demands. With the inflation and increase in fuel price, the break‐even load factor, LF_pax, has gone up in a fluctuating manner, in some cases exceeded 50%.

It is appropriate to introduce here the definition of the dictating economic parameter, seat‐mile cost, which represents the unit of the aircraft Direct Operating Cost (DOC) that determines airfares to meet operational costs and sustain profits. DOC is the total cost of operation for the mission sector. The US dollar is the international standard for aircraft cost estimation. Operational economics are discussed in detail in Chapter 16.

7.1

The seat‐mile cost is the aircraft operating cost per passenger per nautical mile (nm) of the mission sector. The higher the denominator in Equation 7.1, the lower is the seat‐mile cost (i.e. DOC). Therefore, the longer an aircraft flies and/or the more it carries, the lower the seat‐mile cost.

To get some idea, typical Seat‐Mile Costs (cents/seat/nm), are listed here. Turboprop Seat‐Mile Costs are about 10–20% in the comparable class, currently not bigger than regional jets.

A careful market study could fine‐tune an already overcrowded marketplace for a mission profile that offers economic gains with better designs. Section 2.5.1 addresses the market study so that readers understand its importance.

7.4.1 Civil Aircraft Regression Analysis

There are definite statistical relationships between aircraft weight and geometries. The choices are made to configure aircraft not arbitrary – definite reasons are associated with the choices made. Figures 7.2--7.9 (except Figure 7.7) [2] indicate strong relational trends in geometry and weight within the class of aircraft studied evolved through design considerations. Chapters 4–6 covered relevant points on aircraft component properties and status, as summarised next.

1. Wing group. Chapter 4 is concerned with planform shape and size of the wing. It gives various types of wing, their relation between aerofoil thickness to chord ratio and wing sweep versus Mach number. Options for high‐lift devices are described in Section 4.13. This chapter gives the statistics of wing size, wing load, aspect ratio (AR) and wing span versus aircraft weight.
2. Fuselage and nacelle group. Chapter 5 is concerned with shaping and sizing of the fuselage. It gives statistics of fuselage; for example, cross‐section types, fuselage length for the passenger capacity, fineness ratio, seating arrangement and fore and aft end closure data. It also covered nacelle geometries and locations. This chapter covers the statics of engine size versus aircraft weight, which in turns gives the nacelle size.
3. Empennage group. Chapter 6 is concerned with H‐tail and V‐tail geometries and introduces the role of tail volume coefficients. This chapter gives the statistics of H‐tail and V‐tail tail volume coefficients.

7.4.2 Maximum Takeoff Mass versus Number of Passengers

Figure 7.2 describes the relationship between passenger (PAX) capacity and MTOM, which also depends on the mission range for carrying more fuel for longer ranges. In conjunction with Figure 7.1, it shows that lower‐capacity aircraft generally have lower ranges (Figure 7.2a) and higher‐capacity aircraft are intended for higher ranges (Figure 7.2b). 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 4.17. It is possible for the aircraft size to grow, provided the supporting infrastructure is sufficient.

Figure 7.2 shows an excellent regression of the statistical data. It is unlikely that this trend will be much different in the near future. Considerable scientific/technical breakthroughs will be required to move from the existing pattern to better values (the blended‐wing‐body (BWB) configuration is an example when it becomes available). Light but economically viable material, superior engine fuel economy, miniaturisation of systems architecture, microprocessor‐based aircraft and engine controls are some of the areas in which substantial weight reduction is possible.

In conjunction with Figure 7.2, it can be seen that longer‐range aircraft have higher MTOM per passenger as shown Table 7.1 Longer‐range capability requires carrying a higher proportion of fuel. Below 2500 nm, the accuracy degenerates; the weight for in‐between ranges is interpolated.

7.4.3 MTOM versus Operational Empty Mass

Figure 7.3 provides crucial information to establish the relationship between the MTOM and the OEM. Figure 7.3a shows the regression analysis of the MTOM versus the OEM for 26 turbofan aircraft, indicating a predictable OEM growth with MTOM being almost linear. At the lower end, aircraft with fewer than 70 passengers (i.e. Bizjet, utility and regional jet class) have a higher OEMF (around 0.6 – sharply decreasing). In the midrange (i.e. 70–200‐passenger class – single‐aisle, narrow‐body), the OEMF is around ≈0.55. At the higher end (i.e. more than 200 passengers – double‐aisle, wide‐body), it levels out at around ≈0.5 ± 0.05; the MTOM is about twice the OEM. The decreasing trend of the weight fraction is due to better structural efficiencies achieved with larger geometries, as well allowingmore accurate design and manufacturing methods.

Aircraft load (see Chapter 17) is 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 manoeuvrability 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 7.3).

Figure 7.3a is represented in higher resolution when it is plotted separately for the midsized aircraft, as shown in Figure 7.3b for midrange‐size aircraft. It also provides insight to the statistical relationship between the derivative aircraft of the Boeing737 and Airbus320 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 Boeing737‐100 was the baseline design, the smallest in the family. Its growth required corresponding growth in other aero‐structures 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 aero‐structures. In the shortened versions, the A319 came before the even shorter A318, maintaining the philosophy of retaining component commonalities. The variants were not the optimised size, but they were manufactured at a substantially lower cost on account of low development costs from retaining component commonality, decreasing the DOC and providing a competitive edge.

7.4.4 MTOM versus Fuel Load

Figure 7.4 shows the relationship between fuel load, M_f and MTOM for 20 turbofan aircraft, this graph also shows the fuel fraction, M_f/MTOM. It may be examined in conjunction with Figures 7.1--7.3, which show the range increase with the increase in MTOM, fuel mass, M_f and fuel fraction, M_f/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 about half the MTOM. For the same passenger capacity, there is statistical dispersion at the low end. This indicates that, for short ranges (around 1400 nm to cross‐country ranges of around 2500 nm), with a wider selection of comfort levels and choice of aerodynamic devices, the fuel content is determined by the varied market demand. At the higher end, the selection narrows, showing a linear trend. Figure 7.3 indicates that larger aircraft have better structural efficiency and Figure 7.4 indicates that they also have a higher fuel fraction for longer ranges.

7.4.5 MTOM versus Wing Area

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 14). Figure 7.5 shows the relationships between the wing planform reference area, S_W and the wing loading, W/S_W, versus the MTOM.

Wing loading, W/S_w, is defined as the ratio of the MTOM to the wing planform reference area (W/S_W, kg m⁻² or lb_m ft⁻², if expressed in terms of weight; then, the unit becomes N m⁻² or lb_f ft⁻²). 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 7.5. 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. On account of the cost and complexities involved in sophisticated high‐lift devices, the current trend is to have lower wing load to accept simpler high‐lift devices (see Figure 4.34). 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.

The growth of the wing area with an in 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 14) provides the minimum wing planform area to simultaneously satisfy 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. To keep the weight and cost down, simpler types of high‐lift devices are used in small aircraft. This results in lower wing loading 40–100 lb ft⁻² (i.e. ≈200–500 kg m⁻²), as shown in Figure 7.5a. Aircraft with a range of more than 3000 nm need more efficient high‐lift devices. It was shown previously that aircraft size increases with increases in range, resulting in wing loading increases 80–150 lb ft⁻² (i.e. ≈400–700 kg m⁻² for midrange aircraft) when better high‐lift devices are considered.

Here, the trends for variants in the family of aircraft variant design can be examined and given in Figure 7.5b,c. The Airbus 320 baseline aircraft is in the middle of the family. 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 induced drag. For large wide‐body twin‐aisle, subsonic jet aircraft (see Figure 7.8c), the picture is similar to the midrange‐sized, narrow‐body single‐aisle aircraft but with higher wing loadings to keep wing size relatively small (which counters the square‐cube law discussed in Section 4.17).

7.4.6 Wing Geometry – Area versus Loading, Span and Aspect Ratio

The main parameters associated with the wing are its area, AR, span and sweep. Figure 4.23 gives the relations of wing sweep, its aerofoil t/c versus maximum operating Mach number. Here, Figure 7.6 gives the relation between wing load, span and AR and wing area. It may be noted that wing AR requires consultation with the structural engineers to obtain the a high value. In the graph, the sample data are within 7–10. The AR shows a scattering trend. In the same wing‐span class, the AR could be increased with advanced technology but it is restricted by the increase in wing load. Current technology requires it is in a position to have an AR up to 15.

Figure 7.6 shows that the growth of the wing span is associated with the growth in wing loading.

With steady improvements in new material properties, miniaturisation 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.

Examples of some current subsonic‐transport aircraft wing planforms are given in Figure 7.7.

7.4.7 Empennage Area and Tail Volume Coefficients versus Wing Area

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

Statistics show large spread in the values of S_H and S_V with respect to wing area, S_W. Figure 7.8 shows growth for H‐tail and V‐tail surface areas with the wing area, S_W. Reference [2] is the best source to get accurate data on S_H and S_V, these are included in Figure 7.8. However, [2] does not give the tail volume coefficients C_HT and C_VT. These will have to be laboriously computed from the three‐view diagrams given for the aircraft. This can be done with reasonable accuracy if the drawings are enlarged, the drawing scale can be determined from the linear dimensions; for example, wing span, fuselage length and so on. There is a large discrepancy in the C_HT and C_VT values as shown in the band in Figure 7.8. The same pattern can be observed in [1] of a large number of aircraft data given for older aircraft. Other good sources of aircraft statistics are found in [3, 4] and the websites listed in (Websites: (i) http://richard.ferriere.free.fr/3vues/; Specifications of F‐22 (F/A‐22), www.f22fighter.com/Specs.htm; Milavia, military aircraft; www.milavia.net/aircraft/; University of Southampton, Aircraft design resources; www.soton.ac.uk/∼jps7/AircraftDesignResources/LloydJenkinson%20data; Elsevier, Civil Aircraft Design, book website, http://booksite.elsevier.com/9780340741528/authors/default.htm; VRR, aircraft containers, https://vrr‐aviation.com/products/aircraft‐containers/; Searates, air cargo ULD containers, www.searates.com/reference/uld/)

Modern aircraft with fly‐by‐wire (FBW) technology can operate with more relaxed stability margins, especially for military designs, and hence require smaller empennage areas compared with older conventional designs (see Figure 12.18). It is suggested that readers make a separate plot to generate their own aircraft statistics at a better resolution for the particular class of aircraft in which they are interested in order to obtain an appropriate average value.

Typically, for civil aircraft designs, the horizontal tail, S_H, area is about 25–30% of the wing reference area, depending on the length of fuselage. For long, large aircraft, the vertical tail, S_V, area is around ≈12–0.15% of wing reference area, S_W. For long, large aircraft, the vertical tail area is around ≈12–0.15% of wing reference area, S_W. For short aircraft, the vertical tail area is around ≈15–0.18% of wing reference area, S_W. It is expected that readjustment of the V‐tail and adjustment to S_V will be required. More detailed analyses are also expected to be carried out in the next design phase, Phase II.

In this book, the trim surfaces locations are earmarked and are not sized. Designers have to ensure that there is adequate trim authority (trim should not run‐out) at any condition. This is normally done in Phase II after the configuration is frozen.

7.4.8 Aircraft Lift to Drag Ratio (L/D) versus Wing Aspect Ratio (AR)

Equation (4.18) derives the expression for aircraft induced drag, C_Di = (C_L²/πAR), which contributes a significant part (as high as about a third) of aircraft total drag. The higher the AR, the lower the C_Di. The aim is to design aircraft to have as high as possible lift‐to‐drag ratio (L/D).

One of the effective way is to have as high as possible AR the structural design consideration allows. High AR wings are more expensive to design and manufacture, hence they are more meaningful for fare earning transport aircraft. Fuel saving as a result of having lower drag soon offsets the incremental cost of manufacture. The following values given in Table 7.2 may be used in this book, which are typical in the class of aircraft.

There are exceptions and readers may examine such designs.

7.4.9 MTOM versus Engine Power

The relationships between engine sizes and the MTOM are shown in Figure 7.9. Turbofan engine size is expressed as sea‐level static thrust (T_SLS) in the International Standard Atmosphere (ISA) day at takeoff ratings, when the engine produces maximum thrust (see Chapter 11). These graphs can be used only for preliminary sizing; formal sizing and engine matching are dealt with in Chapter 14.

Thrust loading (T/W), is defined as the ratio of total thrust (T_{SLS tot}) 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 Federal Aviation Administration (FAA) requirements in Chapters 14 and 15) with one engine inoperative. For a twin‐engine aircraft, 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 they should provide sufficient takeoff thrust to simultaneously satisfy the (i) field length specifications, (ii) initial climb requirements and (iii) initial high‐speed cruise requirements satisfying the market specifications. An increase in engine thrust with increase of 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 power rated in P per kilowatt (hp or shp), which in turn provides the thrust. Turboprops use 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 7.9a shows thrust growth with size for small aircraft. Here, thrust loading is from 0.35 to 0.45. Figure 7.9b 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 7.9c 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 0.34; the lower values are for four‐engine aircraft.

7.5.1 Flight Crew (Flight Deck) Compartment Layout

The pilot flight deck, of course, is at the front‐closure end of the fuselage to provide forward vision. The maximum accommodation is two side‐by‐side, generously spaced seats; an additional crew member for larger aircraft is seated behind the two pilots (Figure 7.10). In the past, there were two flight crews to assist two pilots; today, with improved and reliable microprocessor‐based multi‐functional display systems, two flight crews have become redundant. There could be provision for one. The windscreen size must allow adequate vision (see Figure 7.10), especially looking downward at high altitudes, during landing and during ground manoeuvres.

The pilot's seat is standardised as shown in Figure 7.11, with generous elbowroom to reduce physical stress. The seat can also be made to recline.

7.5.2 Cabin Crew and Passenger Facilities

Cabin seat arrangements are dealt with in Section 5.8. In this section, crew and passenger facilities are described. A vital fuselage design consideration is offering passenger services – the greater the passenger number, the more complex the design. This book does not cover details of interior design, a specialised state‐of‐the‐art feature that is more than the mere functionality of safety, comfort and efficient servicing. The aesthetics also offer an appealing and welcoming friendly environment to passengers. Physiological and psychological issues such as deep vein thrombosis, claustrophobia and fear of flying can be minimised through careful design of the seat‐pitch arrangement, window locations, environmental controls (i.e. pressurisation and air‐conditioning) and first‐aid facilities. Discussed herein are typical seat pitch, toilet and service‐galley arrangements in fuselage‐space management that contribute to fuselage length.

The minimum number of cabin crew is subjected to government regulations. No cabin crew is required for fewer than 19 passengers, but can be provided if an operator desires. For 19–29 passengers, at least one cabin crew is required. For 30 or more passengers, more than one cabin crew is required. The number of cabin crew increases correspondingly with the number of passengers.

7.5.3 Seat Arrangement, Pitch and Posture (95th Percentile) Facilities

Figures 7.12 and 7.13 illustrate typical passenger seating‐arrangement designs, which can be more generous depending on the facilities offered by the operator. Pitch is the distance between two seats and varies from 28 (tight) to 36 in. (good comfort). Seat and aisle width are shown in the next figure. Typically, seat widths vary from 17 (tight) to 22 in. (good comfort). Seats are designed to meet the 16‐g governmental impact regulations.

Table 7.3 gives typical seat pitch, width and aisle width (there could be a variation of dimensions from operator to operator) as currently prevalent in service. There is flexibility build in to the design to convert seating arrangements, or even remove seats, as the market demands.

Smaller aircraft with fewer passengers (i.e. up to four abreast at the lower mission range) can have a narrower aisle because there is less aisle traffic and service. For larger aircraft, the minimum aisle width should be at least 22 in. (Figure 7.13).

Recently, some operators have offered sleeping accommodation in larger aircraft for long‐range flights. This is typically accomplished by rearranging cabin space – the interior securing structure is designed with flexibility to accommodate changes.

7.5.4 Passenger Facilities

The typical layout of passenger facilities is shown in Figure 7.14 and includes toilets, service galleys, luggage compartments and wardrobes. Cabin crew are provided with folding seats, for large aircraft these are conveniently located as normal seats.

The type of service depends on the operator and ranges from almost no service for low‐cost operations to the luxury of first‐class service. Figure 7.15a illustrates a typical galley arrangement for a midrange passenger‐carrying aircraft; other types of server trolleys are also shown. Figure 7.13 shows a trolley in the aisle being pushed by cabin crew.

Galleys are located in the passenger cabin to provide convenient and rapid service. Generally, they are installed in the cabin adjacent to the forward‐ and aft‐galley service doors. Equipment in the galley units consists of the following:

• high‐speed ovens: refrigeration
• hot‐beverage containers: main storage compartments
• hot‐cup receptacles

The electrical control‐panel switches and circuit breakers for this equipment are conveniently located. Storage space, miscellaneous drawers and waste containers are also integrated into each galley unit.

Figure 7.15b shows a typical toilet arrangement for larger passenger‐carrying aircraft. For a small Bizjet class, the toilet can be minimised (or removed – not preferred) unless there is a demand for a luxury facility.

7.5.5 Doors – Emergency Exits

FAA/Civil Aeronautics Agency (CAA) has mandatory requirements on minimum number of passenger doors, their types and corresponding sizes depending on the maximum passenger capacity the fuselage intended to accommodate. This is to ensure passenger safety – certification authorities stipulate a time limit (90 s for big jets) within which all passengers must egress if an unlikely event, for example fire, occurs. The larger the passenger capacity, the larger the number is of the minimum number of doors to be installed. Not all doors are of the same size – the emergency doors are smaller. All doors are kept armed during airborne operation.

An emergency situation (say from fire hazard, ditching on water or land etc.) would require a fast exit from the cabin to safety. FAA initially imposed a 120 s egress time but from 1967 this changed to a maximum of 90 s. This was possible through the advances made in /chute technology. To obtain airworthiness certification, aircraft manufacturers have to demonstrate by conducting simulated tests that complete egress is possible within 90 s. European Aviation Safety Agency (EASA) has similar requirements.

FAR25 Section 25.783 gives the requirements for the main cabin doors and Section 25.807 for the emergency exit doors. There are several kinds of emergency exit doors given in Table 7.4 (all measurements are in inches). All are rectangular in shape with a corner radius. The sizes are the minimum and designers can make them bigger. The oversized doors need not be rectangular so long they can inscribe the minimum rectangular size.

All doors except Type III (inside step up of 20 in. and outside step down of 27 in.) and Type IV (inside step up of 29 in. and outside step down of 36 in.) are from floor level. If a Type II is located over the wing then it can have an inside step up of 10 in. and an outside step down of 17 in. Emergency doors are placed at both the sides of the aircraft and need not be diametrically opposite, but they should be uniformly distributed (no more than 60 ft away), easily accessible for even loading by passengers when it is required. Safety drill by cabin crew is an important aspect to save lives and all passengers must attend the demonstration, no matter how frequent a flier one is. From type to type there are differences.

There should be at least one easily accessible external main door. The combination of main doors and emergency doors are at the discretion of the manufacturers who will have to demonstrate evacuation within the stipulated time. Fuselage length would also decide the number of emergency doors as they should not be spaced more than 60 ft away. Table 7.5 gives the minimum number of emergency doors. It is recommended that more than minimum provision is to be made. Types A, B and C can also be used – they are deployed in larger aircraft.

There could be other types of doors, for example, a door at the tail cone, ventral doors and so on (dimensions are given in Table 7.6). Flight crew emergency exit doors are separately provided at the flight deck.

7.5.6 Cargo Container Sizes

As the fuselage diameter increases with passenger load, the under‐floorboard space can be used for cargo and baggage transportation. With operating costs becoming more competitive, the demand for cargo shipment is increasing, to the extent that variant aircraft are being designed as cargo aircraft (e.g. no windows and a lower level of cabin pressurisation). An attractive variant is the ‘combi’ design, which can convert the cabin layout according to the sector payload, in which the passenger load is smaller and the cargo load is higher. The combi layout can quickly reconfigure the cabin interior for passengers in the forward part and cargo in the rear, which facilitates passenger loading and unloading through the front door.

Cargo and baggage could be handled more efficiently by keeping items in containers (Figure 7.16) and having both destination and interior‐space management. At the destination, the entire container is unloaded quickly so the aircraft is free for quick turnaround utilisation. Container sizes are now standardised to fit in the fuselage and are internationally interchangeable.

The term unit load device (ULD) is commonly used when referring to containers, pallets and pallet nets. The purpose of the ULD is to enable individual pieces of cargo to be assembled into standardised units to ease the rapid loading and unloading of aeroplanes and to facilitate the transfer of cargo between aeroplanes with compatible handling and restraint systems.

Those containers, intended for below‐floorboard placement (designated larger container, load device: LD), need to have the base smaller than the top to accommodate fuselage curvature. Those containers have rectangular cross‐sections and are designated ‘M’. Figure 7.16 shows typical container shapes; Table 7.7 lists standard container sizes, capacities (there are minor variations in dimensions) and designations.

7.8.1 Fighter Aircraft Generations

The post‐war period of development of aircraft production was marked by a scientific and technical revolution – the early era of jet aircraft. In the late 1930s, it became clear that ordinary planes with piston engines and propellers had been developed to almost its maturity and that there was not much room for further improvement unless new technologies could be implemented. Since WWII, with the introduction of the jet age, each successive generation of fighter has demonstrated increased speed and altitude compared to its predecessor, to provide a machine better adapted than that of any competitor to the needs of all potential customers.

In a way, post‐World War II fighter aircraft designs evolved in discrete steps of technological advancement and are typically categorised in terms of generation of design evolution, currently in five (or six) generations as follows (Figure 7.20).

• ‘Zeroeth’ generation (1945–1955). The ‘Zeroeth’ generation fighter aircraft were the first military aircraft using jet engines. A few were developed during the closing days of WWII but saw very limited combat operations. These include WWII era fighters such as the Me262 and early postwar aircraft such as the F‐80 and F‐84.
• First generation (1945–1955). The first generation can be split into two broad groups: (i) WWII era fighters such as the Me 262 (also Zeroeth generation) and (ii) the mature first‐generation fighters such as the F‐86 used in the Korean War. Mikoyan's design office worked on the twin‐engine fighter MiG‐9. Yakovlev's design office brought out the single‐engine fighter Yak‐15 in October 1945; it was already on the airfield for preliminary tests and for taxiing. The MiG‐9 and Yak‐15 promised to be lighter, easier to fly, to have better flying characteristics and be more reliable than the German planes. Examples of first generation jet fighters are the F‐84, F‐86 and F‐100.
• Second generation (1955–1960). The second‐generation fighter aircraft exhibit more advanced avionics, engines and used the first guided air‐to‐air missiles. The period from 1950 until 1955 is marked by a dearth of significant interceptor prototypes except for the 1953 appearance of the MiG‐17. Second‐generation aircraft – including the MiG‐19/21 and US century series fighters – were designed during the 1950s. Although they are still found in fighter inventories worldwide, older planes probably have limited combat potential when confronting more modern fighters, since they may suffer from several disadvantages. For example, they may carry less sophisticated munitions and have less capable sensors. The early MiG‐ and Su‐series aircraft have been improved in their air‐to‐air role. The MiG‐23 Flogger B was a second‐generation fighter that had a secondary ground attack capability greater than the Fishbed or Fitter.
• Third generation (1960–1970). The third‐generation fighter aircraft exhibit yet more advanced avionics, engines and weapons. The changes in the fighter combat conception, new air‐to‐air guided missiles and the results from first‐ and second‐generation fighter operations gave rise to the third generation, such as the later versions of the F104S and MiG‐23. Third‐generation aircraft may provide somewhat more military capability, especially if they have gone through extensive modifications since they were built. Designed during the 1960s through the1970s, this generation includes the MiG‐27 series designed by former Soviet Union's Mikoyan Design Bureau. The third‐generation fighter‐bombers and tactical bombers included the Su‐24 and their derivatives, the F‐4s and A‐7s built by the USA and the European designed Mirage 3, Mirage 5, F1 and Tornado.
• Fourth generation (1970–1990). The fourth generation continued the trend towards multi‐role fighters equipped with increasingly sophisticated avionics and weapon systems. These fighters also began emphasising manoeuvrability rather than speed to succeed in air‐to‐air combat. Fourth‐generation fighters, designed during the 1960s and 1970s, include the US‐designed F‐14, F‐15, F‐16 and F/A‐18; the Soviet‐built SU‐27 and MiG‐29 and the European Mirage 2000. Fourth‐generation aircraft usually have more sophisticated avionics than their predecessors, more powerful engines and are able to operate more capable missiles.
• Fifth generation (1990–2010). The fifth‐generation fighters use advanced integrated avionics systems to provide the pilot with a complete battlespace awareness, and use of low‐observable ‘stealth’ technology. The F‐22 and F‐35 were the first fifth‐generation US fighters, with Russia following with the Mikoyan Gurevich MFI prototype and the Sukhoi PAK‐T50 and China with the J‐20.

Making the situation even more confusing, the Chinese have their own particular order of ranking. The Chinese call the Russian fifth‐generation fighter a fourth‐generation machine.

The term ‘4.5 generation’ is also sometimes seen (the Chinese call them ‘3.5 generation’). These are more recent fourth‐generation fighters, retaining the basic characteristics of fourth‐generation planes but with enhanced capabilities provided by more advanced technologies that might be seen in fifth‐generation fighters. Good examples are the F/A‐18E/F Super Hornet, Eurofighter Typhoon and Dassault Rafale. All make use of advanced avionics to improve mission capability and limited stealth characteristics to reduce visibility when compared to older fourth‐generation aircraft.

Stealth characteristics markedly enhance the capability for survivable attack of defended targets. And that remains true, even though evolving modern air defences available on the international arms markets have increasing capability against currently deployed levels of stealth. Hence, to continue to operate effectively in the face of these defences, stealth has to be supplemented with other survivability features. Nonetheless, stealth aircraft operate at much lower levels of support than conventional aircraft and even small numbers of stealth aircraft can greatly leverage the capabilities of the remainder of the bomber force and of the tactical fighter forces.

Europe and Japan are behind in applied stealth technology as evidenced by US aircraft programmes such as the F22, B2 and F117A. Most of the applied foreign low‐observable work involves basic shaping, material coating techniques and signature testing requirements. Europe has been led by France, Sweden, Germany and the UK in various types and levels of low‐observable applications. Applications on fighter aircraft have generally been at fundamental applied levels, primarily using absorbent coatings, limited structural shaping and absorbent structure. Applications seem to be limited to the areas with highest signature return rather than application to an entire airframe. European firms are also working on stealth technology applications to cruise missiles and unmanned aerodynamic vehicles.

The rapid growth of a globalised economy has deepened the degree of international cooperation and expanded the variety of methods of cooperation in the international arms production industry to such an extent that a globalisation trend has also emerged in this field. The globalisation of the national defence industry refers to the change from the traditional preference for autarky to that of a globally oriented market in terms of research, production, management and sales. At present, there are mainly three ways in which globalisation exhibits itself. The first is through the purchasing of weapons from other countries and taking part in the production of these weapons (including granting of special permits, joint cooperation and development ventures and compensation trade). An example is the joint production of F16 fighter jets by USA, Holland, Denmark and Norway. The second method is through military cooperation packages covering weapons trade, production and maintenance and joint military exercises between different countries, for example, the signing of the 10‐year military cooperation agreement package between India and Russia in 1999. The final means is through cross‐border joint ventures and joint research and development projects between nations (including international group companies, international integration and transnational amalgamation). The four‐nation joint venture for the production of the Eurofighter‐2000 by the UK, Germany, Italy and Spain is an example.

7.10.1 Military Aircraft MTOM versus Payload

Figure 7.22 shows MTOM versus payload and TTOM. There is a distinct separation on armament loading capability between air defence class fighter aircraft with armament load of around 750 kg and the multi‐role class aircraft carrying nearly twice the payload for a longer distance. Air Defence class fighter aircraft are lighter than the multi‐role class aircraft. Understandably, in both the classes the MTOM grows with increase in payload (armament).

7.10.2 Military MTOM versus OEM

Figure 7.23 gives the relation between MTOM and OEM as well as the Operational Empty Mass Fraction (ratio of OEM to MTOM). OEM grows with growth in MTOM but there is a spread in the OEM fraction arising out of differing performance capabilities and system integration. Typically, the ratio of OEM/MTOM averages around 0.42–0.52.

7.10.3 Military MTOM versus Fuel Load, M_f

Since fuel load for combat aircraft is flexible depending on usage of drop tanks and air‐to‐air refuelling, onboard fuel content is taken as the standard condition to present the statistical analysis. Figure 7.24 gives the relationship between internal fuel load, M_f and fuel fraction, M_f/MTOM versus MTOM. There is some scatter on account of diversity in requirements.

Growth in MTOM would be associated with more fuel carrying capacity to meet range, but the fuel fraction graph shows dispersion on account of the difference in role, for example, short range air defence and longer range deep penetration types. Longer‐range aircraft would be heavier due to having to carry more fuel.

7.10.4 Military MTOM versus Wing Area

The influence of wing loading shows up in the graph (Figure 7.25). Military designs could have moderate wing loadings for hard manoeuvres. Modern designs show lower wing loading – the F22 has the low order of 350 kg m⁻² compared to the Sepecat Jaguar with around 650 kg m⁻². It will be seen in the next section that the F22 also has the highest thrust loading.

7.10.5 Military MTOM versus Engine Thrust

Combat aircraft are invariably powered by jet propulsion (turboprop driven CAS aircraft are few and are excluded in the statistics). Military aircraft would require very high thrust loading, T/W (could exceed 1), for manoeuvres and short field performance. A high thrust requirement is of a short duration and is met by augmenting thrust by the use of afterburners (Chapter 10).

Section 7.10 explains the need for TTOM for combat effectiveness. Figure 7.26 presents the relationship between total T_SLS and the two types of aircraft mass, for example, MTOM and TTOM. Thrust increase is associated with aircraft mass increase. However, there is some spread in thrust loading. Typically, (total T_SLS/MTOM) averages around 0.65 but (total T_SLS/TTOM) exceeds 1. Later generations of combat aircraft have pushed thrust to weight ratio more than one would permit aircraft to accelerate in vertical climb. The F22 has the highest thrust loading as well as the lowest wing loading.

7.10.6 Military Empennage Area versus Wing Area

Military aircraft empennage configuration should be very different from civil aircraft design; many have a conventional design with an H‐ and V‐tail. On the other hand, an extreme example of the B2 appeared to not have any tail. Examining various configurations, it can be seen that there are several options for aircraft control.

Stability and control of modern combat aircraft are invariably supported by microprocessor‐based systems architecture such as FBW, when onboard computers continuously fly a slightly unstable aircraft under pilot initiated commands. This is beyond the scope of this book, so no statistical analysis could be presented here. The examples of the trainer class of aircraft would follow the conventional approach with a one H‐ and one V‐tail design (use Figure 8.5).

Military aircraft require more control authority for greater manoeuvrability and have shorter tail arms requiring larger tail areas. Typically, the horizontal tail area is about 30–40% of the wing reference area. The vertical tail area varies from 20 to 25% of the wing reference area. Supersonic aircraft have an all movable tail for control. If the V‐tail is too large then it is split into two.

7.11.1 Fuselage Group (Military)

Unlike the approach to civil aircraft configuration, military design need not start with the fuselage, but it may prove convenient to do so. Fighter aircraft fuselage does not carry any internal payload – it has the singular function to accommodate the crew (or crews) and engine (or engines) along with routing of conduits of various systems (wires, pipes, linkages for the systems), fuel tanks and encasing small arms (e.g. guns). Unlike hollow civil aircraft fuselage, it is very tightly packed, minimising the fuselage volume requirement. With the engine (or engines) buried inside, air intake is an integral part of the fuselage. The large wing root of delta (or trapezoidal wing with strake) planform offers the scope to make the wing blend with the fuselage. In that case, the configuration of the wing becomes integral to configuring the fuselage, as shown in Sections 5.2.1.4 and 5.2.1.5. A variety of well‐known combat aircraft images are shown in Figure 7.27.

The densely packed fuselage design starts with the nose cone, which has to be pointed for supersonic capability and then houses a radar that could be of around 1 m diameter. Fighter aircraft fuselage would invariably house the power plant and, therefore, there will not be any separate podded nacelle (some older bombers have pods). The necessity for area ruling makes for narrowing of the fuselage. Therefore, the fuselage would rarely have a constant cross‐section, making fuselage shape generation quite complex.

The fuselage aft ends up as an engine exhaust system and therefore will not have closure as in civil design. In case the engine dangles below the fuselage spine (Figure 7.19b – Phantom), then a pointed aft end closure follows. The fuselage belly fairing would be housing accessories, in most cases the undercarriage. Current tendencies are for the wing‐body fairing to have considerable blending for superior aerodynamic considerations, for example, to improve lift to drag ratio and fly at a higher angle of attack. In a blended fuselage it is hard to isolate it, a possible convenient choice would be where the wing root is attached.

Also, a military aircraft pilot seat has more freedom to recline to shorten carotid artery height to reduce blood starvation to the brain at high g manoeuvres that can cause blackouts.

7.11.2 Wing Group (Military)

The wing group is the most important component of the military aircraft. The wing planform shape needs to be established based on the operational requirements (e.g. hard manoeuvres, supersonic capabilities, short field performances). Unlike civil design, there is a large range of options for planform shape and fuel tankage space is restricted.

The evolution of fighter aircraft shows the dominant delta or short trapezoidal wing planform. This is for obvious reasons of having high leading‐edge sweep and low AR to negotiate high g manoeuvres that would generate a high wing root bending moment. It would restrict span growth but encourage large wing root chord of delta or trapezoid with strake planform as shown in Figure 7.28g,h. For control reasons it could have additional surfaces. The following are the configuration choices (strakes are taken as part of wing):

1. One‐surface configuration. Pure delta planform or its variation – the trailing edge of the delta like wing can be made to work like an H‐tail as an integral part of the wing (Figure 7.28a).
2. Two‐surface configuration. Delta‐like wing or trapezoidal wing with a conventional H‐tail for pitch control. In some designs, the H‐tail is replaced by a canard surface for pitch control in relaxed stability (has a destabilising effect). A two‐surface configuration has two possibilities – a tail at the back (Figure 7.28b) or a tail in front (canard – Figure 7.28c). A two‐surface configuration with strake is shown in (Figure 7.28d). Variants include the double delta (SAAB Viggen, see Figure 7.28c).
3. Three‐surface configuration. The ultimate kind is a three‐surface configuration (Figure 7.28e). It has a wing, H‐tail at the aft end and canard at the front end.

A delta wing trailing edge would have a pitch control surface integrated with it. A trapezoidal wing planform (Figure 7.20c,d) could be associated with a separate pitch control surface, typically as an H‐tail. An extreme form of the three‐surface arrangement exists (Sukhoi 37 – Figure 7.28e). Forward sweep has aerodynamic merits to bring the wing aerodynamic centre to move forward, which favours H‐tail sizing. However, aeroelastic problems could aggravate wing twist creating instability. Carefully arranged composite material has minimised the effect of twist and there are two successful designs that have been flight tested (Su 47, see Figure 7.28e, and Grumann X29).

The role of the canard in military application is quite different from that in the role of civil aircraft designs. It has been found that strakes can also provide additional vortex lift and fast responses to pitch control with a conventional tail. The choice of strake or canard is still not properly researched in the public domain. It is interesting to note that US designs have strakes while the European kinds have canards. Detailed study of aircraft control laws and FBW system architecture is required to make the choice. Until the 1990s, flaws in FBW software have caused several serious accidents.

Again, it is emphasised that this book is introductory in nature. For not having enough information on modern fighter design considerations, the author restricts military aircraft design exercises to a trainer class of aircraft. The readers will find that there is a lot to learn from this class of aircraft, enough to have a feel for military aircraft design considerations. Figure 7.29 shows the modern Advanced Jet Trainers that have variants for CAS roles. The Ae Hawk is a successful but relatively older design (still in production) that has a conventional configuration. EADS MAKO is one of the latest trainer aircraft designs capable of supersonic flight and light combat capability.

Wing attachment to fuselage varies from case to case. The leading edge can have slats and the trailing edge would invariably have flaps. Centrally mounted large air brakes to decelerate have practically eliminated the role of spoilers. Landing in shorter airfields may require deployment of a brake parachute.

Since military aircraft are expected to encounter transonic flight, aircraft cross‐sectional area distribution becomes an important consideration. A seamless smooth distribution of cross‐section (area‐rule) is explained in Section 3.13.

7.11.3 Empennage Group (Military)

Combat aircraft empennage shaping and sizing are complex procedures, primarily on account of a short tail arm and the need to fly in relaxed stability to execute fast and hard manoeuvres (Chapter 6). The B2 apparently appears to be without a tail. The F22 has a large canted V‐tail. Options for control surface configuration are shown along with wing options. The delta wing has an H‐tail integrated with it. This book adheres to the conventional configuration of H‐ and V‐tail for the trainer aircraft example. Modern designs deploy tailerons (stabilator; see Section 6.5.3) to initiate pitch and roll control by an H‐tail.

Introductory comments expressed the complexity involved in control surface design, which are primarily of the empennage group. Having short tail arms (L_HT and L_VT) the stability (stabiliser/fin) and control surfaces (elevator/rudder) will necessarily have to be large in relation to the aircraft size to make fast responses. In many designs the vertical tail is split into two (could be placed slightly inclined in a shallow Vee for stealth reasons) as can be seen in the F18. The canted V‐tail of the F22 is on account of stealth considerations. The horizontal tail is symmetrical to the vertical plane and has to be secured on the fuselage – some of the earlier designs had a T‐tail.

Military aircraft control is through FBW technology, which processes control deflection through onboard computers ensuring safety. If the pitch control demand were high requiring flying in relaxed stability, then a canard surface in front of fuselage would prove helpful. The advent of FBW has achieved yaw control without a V‐tail – it is achieved through differential use of an aileron surface that can be split to open up both in upward and downward directions, if required simultaneously. This book will not be discussing control‐configured designs (control‐configured vehicle: CCV) – these would require analysing the control laws, not dealt with here.

Delta wings can have H‐tails integrated within with the reflex built‐in at the trailing edge (Mirage 2000). The exception of an older design is shown in Figure 7.30a (YF12) with two canted vertical tails. The YF12 design paved the way for the current wing and empennage design options. An unconventional empennage exists (Figure 7.30).

The B2 in Figure 7.30 appears to be tailless. Its pitch control is at the inboard trailing edges and directional/lateral controls are carried out by controlled opening of a split aileron in both sides. Much of the future will depend on how many lifting surfaces are used. Typically, a highly manoeuvrable combat aircraft will have a large V‐tail split into two and canted for stealth reasons. Trainer aircraft favour a conventional type with an H‐ and V‐tail.

7.11.4 Intake/Nacelle Group (Military)

Unlike a simpler pod‐mounted configuration of civil aircraft, the military aircraft power plant is kept within the fuselage. Therefore, military aircraft of the combat class do not have nacelles, instead they have an (supersonic) intake that is an integral part of fuselage configuration. Supersonic intake is comparatively more complex to design. A supersonic intake has a side plate acting both as boundary layer bleeder and adjust oblique shock. Intake design requires serious considerations to cater for the full flight envelope to supply air inhalation to the engine without causing flow distortion at the engine face that can cause compressor surge and/or flame‐out problems.

Combat aircraft use the aft fuselage to house the engine. Broader classification (options) of military fighter aircraft engine intake configuration is given in Table 5.1. A large number of intake configuration types are given Figure 5.34.

Single or multi‐engines (so far mostly two) are kept side‐by‐side buried in the fuselage. Fighter aircraft intake path is longer and curvier. Older designs had a forward intake at the nose (MiG21). Aircraft with over Mach 1.8 capability have a centre body (bullet translates forward/backward to adjust bow oblique shock). These have the longest intake duct with low curvature to pass under the pilot and incur high loss, hence they are no longer pursued. Instead, chin (F8 Crusader) or side (Hawk) mounted intakes have shorter ducts and these carefully designed ducts have comparable curvature, hence less loss. A plate is kept above the fuselage boundary layer on which the intake is placed. A centre body is required for aircraft speed capability above Mach 1.8, otherwise it has a pitot intake: trapezoidal slanted side intake (F22). The B2 has over wing/fuselage intakes more suited to the BWB type of configuration.

Some older and odd type engine integration configurations with over the fuselage engine installation are shown in Figure 7.31.

In summary, intakes on fuselage have following possibilities:

1. Central forward intake (invariably circular/MiG21 – can be near circular F100). These are older designs, no longer used in practice. Considerable duct loss is associated with its long length from nose to tail and bends to pass underneath the pilot. For supersonic operation, a centre body arrangement makes diffusion through a series of oblique shocks and a normal shock. The centre intake does not have fuselage shielding at yaw and pitch attitudes.
2. Side intakes (semi‐circular, rectangular, e.g. F18 and F22 etc.) – it cuts down the internal duct length by nearly half but associates with bends, which is less for two side‐by‐side engines. It can be over or under the wing. For supersonic operation, there is a splitter plate that bleeds the fuselage boundary layer to keep it outside the intakes. It needs to be carefully sized for flying in yawed attitude.
3. Chin mounted intake (near elliptical/kidney shaped like the Falcon F16, near rectangular like Eurofighter). These are later designed aircraft. At yaw, chin intake does not have fuselage shielding as there could be for side intakes, but being close to the ground there could be ground ingestion problems, especially during war time. This concept has proven very successful and can handle high incidence flying.
4. Central over fuselage mounted intake (this is opposed to a chin mounted intake like the Predator UCAV aircraft). This configuration is not prevalent – high incidence flying may create serious flow distortion affecting engine performance. This type of configuration is gaining ground, however, on account of stealth considerations.
5. Future designs will have stealth features and F22/B2/JUCAS type nacelles will become common shapes of intake design.