Category AIRCRAF DESIGN

Following are additional considerations that could enhance aircraft performance but are not addressed here. At this design stage, none of the additional surfaces described needs to be considered except the dorsal fin. All add to aircraft weight.

1. Winglets. It took some time to establish the merits of having winglets that can reduce or induce drag – some manufacturers claim a reduction as high as 5% of induced drag (i. e., approximately 1.5% in total drag reduction), which is sub­stantial. Currently, almost all large-aircraft designs incorporate winglets. Lear – jet has been using them for some time and they have become a symbol of its design.

2. Dorsal Fin. A dorsal fin ahead of the V-tail could work like strakes on a wing, and they are incorporated in many aircraft – at least to a small degree. They prevent the loss of directional stability.

3. Ventral Fin. This is sometimes installed at the tail end as an additional surface to the V-tail. The ventral fin also serves as a skidding structure to protect the fuselage from damage at excessive early rotation, which causes tail-dragging.

4. Delta Fins. These fins come in pairs at the aft end of the lower fuselage. Not all designs have delta fins; they are used if an aircraft shows poor stability and/or control problems. Aircraft with a flat, rear-loading, raised fuselage upsweep demonstrate these problems and delta fins are deployed to resolve them. A good design should avoid incorporating delta fins; however, on some designs, drag reduction can be achieved with their installation.

Several external-surface perturbations on aircraft add to parasitic drag, including antennas, inspection-hatch covers, vent pipes, and lightning dischargers. Engine and system intake and exhaust ducts and vents also increase drag.

It is suggested that readers determine whether there are any innovative require­ments that should be incorporated in the conceptual design. Trends should be inves­tigated continually for ideas to improve on aircraft design.

It is interesting to observe how the aircraft is gradually taking shape – it is still based on a designer’s past experience but soon will be formally sized to a satisfying rational configuration to offer the best characteristics for the design.

A preliminary three-view diagram of the civil aircraft can now be drawn (Figure 6.14). It will be revised after the remaining aircraft components are positioned and a revised CG location is established. The next iteration is after air­craft sizing in Chapter 11.

At this stage, all aircraft components are ready to be assembled using the building-block concept to generate a preliminary aircraft configuration, as shown in Figure 6.14. The three variants (see Figure 6.8) maintain the same wing, empennage,

and nacelle (some internal structures are lightened or reinforced without affecting manufacturing jigs and tools).

The configuration is similar to the Learjet 45 but it is not the same; there are con­siderable differences in configuration, component weights, and performance. Read­ers may compare the two using the Jane’s All the World’s Aircraft Manual.

Chapter 11 sizes the aircraft to its final dimensions and finalizes the configura­tion based on the aircraft and component mass worked out in Chapter 8. Following is a summary of the worked-out civil aircraft preliminary details (from statistics):

Market Specifications

Payload: 10 Passengers + Baggage: 1,100 lb HSC Mach: 0.74 Initial Climb Rate: 16 miles/s Takeoff Field Length: 1,000 m

Baseline Aircraft Mass (from statistics)

MTOM: 9,500 kg («21,000 lb)

Fuel Mass: 1,200 kg («2,650 lb)

Baseline External Dimensions

Fuselage (determined from passenger capacity)

Length: 15.24 m (50 ft)

Overall Width: 173 cm (68.11 in)

Average Diameter: 175.5 cm (70 in)

Wing (aerofoil 65-410)

Planform (Reference) Area: 30 m2 Root Chord, CR: 2.87 m (9.4 ft)

MAC: 2.132 (7 ft)

Dihedral: 3 deg, Twist: 1 deg (washout)

V-Tail (Aerofoil 64-010)

Planform (Reference) Area: 4.4 m2 (47.34 ft2) Root Chord, CR: 2.57 m (8.43 ft)

MAC: 2.16 (7.1 ft) t/c: 10%

Rudder: 0.75 m2 (8 ft2)

H-Tail (T-tail, aerofoil 64-210 – installed with negative camber)

Planform (Reference) Area: 5.88 m2 (63.3 ft2) Span: 5 m (16.4 ft) AR = 4.42 Root Chord, CR: 1.54 m (5.04 ft) Tip Chord, CT: 0.77 m (2.52 ft)

MAC: 1.19 m (3.9 ft) Taper Ratio, X: 0.5 Л1 = 15 deg

Dihedral: 5 deg Elevator: 1.21 m2 (13 ft2) t/c: 10%

Nacelle

Length: 2.62 m (8.6 ft) Maximum Diameter: 1.074 m (3.52 ft)

Bare Engine (each)

Takeoff Static Thrust at ISA Sea Level: 3,800 lb (17,235 N) per engine with BPR = 5

Engine Dry Weight: 379 kg (836 lb)

Fan Diameter: 0.716 m (28.2 in)

Length: 1.547 m (60.9 in)

Short Variant (all component dimensions except the fuselage length are invariant) Fuselage: Length: 13.47 m (44.2 ft) (see Figure 6.8).

Long Variant (all component dimensions except the fuselage are invariant) Fuselage: Length: 16.37 m (53.7 ft) (see Figure 6.8).

Chapter 7 provides details of the undercarriage (i. e., landing gear) design. There is little difference between civil and military aircraft design layouts in undercarriage positioning.

Undercarriage positioning is CG-dependent. At this design stage, the CG posi­tion is not established because aircraft component weights are not known. It is now evident that an iterative process is necessary. From experience, the undercarriage may be positioned after estimating the CG position and rotational tail clearances. Ensure that the aircraft does not tip in any direction for all possible weight distri­butions. (Tipping occurs in some homebuilt designs – especially the canards – when the pilot steps out of the aircraft.) This book addresses only the tricycle type – that is, a forward nose wheel followed by two main wheels behind the aftmost CG. The undercarriage position shown in Figure 6.13 is approximately 60% of the MAC. Readers should use the three views.

This section provides an example for configuring the nacelle based on an engine bought from an engine manufacturer. (Figure 4.9 gives the relationship between MTOM and engine thrust. Chapter 10 gives more details of engine dimensions).

Figure 6.13. Statistics in the aircraft class: the uninstalled thrust of a turbofan

coarse; however, Figure 6.13 provides similar information in finer detail confined to the aircraft class. The author recommends that readers produce graphs in higher resolution for the aircraft class under consid­eration. Unlike aircraft in general, the external dimensions of variant engines in a family do not change – the thrust variation is accomplished through internal changes of the engine (see Chapter 10). The same nacelle geometry can be used in all vari­ants. For major variations, the engine size changes slightly, with minimal changes affecting the nacelle mould lines.

From the statistics in Figure 6.13, for a MTOM of 9,500 kg, a typical uninstalled engine thrust for this aircraft class indicates that TSLs/engine = 3,800 lb ± 25% for the derivative variants for the aircraft family to be offered. This may be considered a smaller engine. For better fuel economy, a larger BPR is desirable. Not many engines are available in this class. It is important that a proven, reliable engine from a reputable manufacturer be chosen; of interest are the following:

Honeywell (originally Garrett) TFE731 turbofan-series class.

Pratt and Whitney (Canada) PW 530 series class (not many variants available)

(In the small engine class, Williams is coming up but is still below the required size.)

The Rolls Royce Viper and the Turbomeca Larzac have a low BPR and are suited to a military application. This leaves the Honeywell TFE731-20 turbofan class as practically the only choice. It has a fan diameter of 0.716 m (28.2 inches), a bare engine length of 1.547 m (60.9 inches), and a dry weight of 379 kg (836 lb). At this stage, a generic long-duct nacelle pod to house is used (see Figure 6.13).

Using the relationship given in Equation 6.6, the maximum nacelle diameter =

1.5 x 0.716 = 1.074 m (5.52 ft).

Using the relation given in Equation 6.7, the nacelle length = 1.5 x 0.716 + 1.547 = 2.62 m (8.6 ft).

The nacelle fineness ratio = 2.62/1.074 = 2.44.

Being a small aircraft, the engines are aft-fuselage-mounted, one at each side. At this stage, a horizontal plate may represent the pylons that support the nacelles. The pylon length = 2.44 m (8 ft) with a thickness of 25 cm (9.8 in) and having a symmetrical cross-section aerofoil-like structure for ease of manufacture. Inlet and exhaust areas are established in Chapter 10.

The nacelle pod size depends on the choice of engine. At this design stage, a sta­tistical value of uninstalled TSLS per engine is considered to determine the size of an engine. A formal engine sizing and matching is accomplished in Chapter 11. For better fuel economy, a large bypass ratio is desired. Dialogue with engine manufac­turers (that can offer the class of engines) continues with “rubberized” engines (i. e., engines scalable and finely tuned to match the aircraft performance requirements for all variants). There are not many engine manufacturers from which to choose.

Numerous engine accessories (see Chapter 10) are part of the engine power plant. They are located externally around the casing of the engine (i. e., turbofan or turboprop). In general, these accessories are located below the engine; some are distributed at the sides (if the engine is underwing-mounted with less ground clear­ance). Therefore, the nacelle pods are not purely axi-symmetric and show faired bulges where the accessories are located.

Long-duct nacelles, chosen for the example, appear to be producing a higher thrust to offset the weight increase of the nacelle, while also addressing environ­mental issues of substantial noise reduction. Also, long-duct designs could prove more suitable to certain types of thrust reverser designs. This book only considers long-duct design but it does not restrict the choice of short-duct nacelles.

For this example, the maximum nacelle diameter « <1.5 x engine-face diameter

(6.6)

In general, the intake length in front of the engine face « <1.0 x engine-face diameter, and the exhaust jet-pipe length aft of the last stage turbine disc « <1.5 x engine-face diameter.

The total nacelle length « (engine length) + (k x engine-face diameter) (6.7)

where 1.5 < k < 2.5. For smaller engines, the value of k is lower.

For long-duct nacelles, the fineness ratio (i. e., length/maximum diameter) is between 2 and 3.

Pylons are the supporting structures (i. e., cross-section streamlined to the aero­foil shape) of the nacelle attaching to the aircraft and carrying all the linkages for engine operation. Aft-fuselage-mounted pylons are generally horizontal but can be inclined if the nacelle inlet must be raised. For wing-mounted nacelles, the pylon is invariably vertical. The depth of the pylon is about half of the engine-face diame­ter; the pylon length depends on the engine position. For an aft-fuselage-mounted installation, the pylon is nearly as long as the nacelle. For a wing-mounted installa­tion, the nacelle is positioned ahead of the wing LE to minimize wing interference. In general, the t/c ratio of the pylon is between 8 and 10%.

The nacelle size is determined from the matched-engine dimensions. Using the considerations listed in Section 6.3.4, the following stepwise approach is suggested. The engine-thrust level indicates engine size (Figure 6.13). It is best to obtain the engine size from the manufacturer as a bought-out item.

Step 1: Configure the podded nacelle size.

The maximum engine diameter determines the maximum nacelle diameter. The ratio of the maximum nacelle diameter to the maxi­mum engine diameter is given statistically in Chapter 10. Similarly, the length of the nacelle is established from the engine length. The keel cut is typically thicker than the crown cut to house accessories. In this book, the nacelle is symmetrical to the vertical plane but it is not a requirement.

Step 2: Position the nacelle relative to the fuselage.

The nacelle position depends on the aircraft size, wing position, and stability considerations (see Section 4.10). Subsequently, CFD analy­sis and wind-tunnel testing will fine-tune the nacelle size, shape, and position.

Step 3: Use pylons to attach the nacelle to the aircraft.

A worked-out example follows in the next section.

Continuing with the fuselage and wing design example carried out in the previous sections, this section presents a worked-out example of empennage design. The aircraft specification used so far to configure the fuselage and wing is sufficient for empennage design. Figure 6.10b provides empennage statistics of the current Bizjet aircraft class. The empennage area size depends on tail arm length, which is not compared in the graphs. A coursework example would have a slightly smaller tail area than shown in Figure 6.10b for having a relatively larger tail arm (the high sweep of the V-tail is added to the tail arm – shown is an example of a designer’s choice for weight reduction). It is the tail volume coefficients that decide the tail areas.

To maintain component commonality, the empennage is the same for all three variants. The baseline-designed empennage area is made sufficient for smaller air­craft; larger aircraft have a longer tail arm to enhance the empennage effectiveness. So far, the civil aircraft design exercise provided the following data:

• Estimated aircraft weight = 9,500 kg (at this stage, not required for empennage sizing)

• Wing reference area = 30 m2 (low-wing design is popular and therefore chosen)

• Wing MAC = 2.2 m (computed from Equation 3.21)

• Fuselage length = 50 ft (aircraft length is different – see Figure 6.3)

To minimize the fuselage length, a T-tail configuration is chosen. The V-tail design arrangement is determined first to accommodate the position of the T-tail on top. Figure 6.12 illustrates the tail-arm lengths used to compute empennage areas.

Section 12.5 provides statistics for the V-tail volume coefficient, Cvt, within the range 0.05 < Cvt < 0.12. In the example, Cvt = 0.07 is appropriate for the smaller aircraft variant. The V-tail quarter-chord sweepback is 15 deg in line with the wing sweep, to increase the tail arm LVT = 7.16 m (23.5 ft) measured from the aircraft CG to the V-tail MAC. In general, SVT/SW ^ 0.12 to 0.2. The symmetrical aerofoil section is the NACA64-010. The V-tail height (semispan) = 7 ft (2.14 m) and the taper ratio = 0.6 to bear the load of a T-tail.

Equation 3.31 gives the V-tail reference area SVT = (CVT)(SW x wing span)/LVT. The V-tail is positioned on the fuselage end in consultation with structural engi­neers. Then, SVT = (0.07 x 30 x 15)/7.16 = 4.4 m2 (47.34 ft2). This would result in sensible geometric details of the V-tail, as follows:

• Note: Area, Sv = 12 (Cr + Ct) x b or 4.4 = 0.5 x 1.6 Cr x 2.14

• Root Chord = 8.43 ft (2.57 m)

• Tip Chord = 5.05 ft (1.54 m)

• Aspect Ratio = 2.08

• MAC = (2 x [(8.43 + 5.05) – (8.43 x 5.05)/(8.43 + 5.05] = 6.8 ft (2.07 m)

• The V-tail area must be shared by the rudder and the fin. Typically, the rudder encompasses 15 to 20% of the V-tail area – in this case, it is 17%. This gives a rudder area of 0.75 m2 (8 ft2).

To check the Cvt for the smaller variant, it should be more than 0.06. With one seat pitch plug removed from the aft fuselage, LVT_shon = 7.16 – 0.813 = 6.347 m (20.823 ft). This gives CyT_short = (4.4 x 6.347)/(30 x 15) = 0.062 (sufficient for the shorter variant).

Section 12.5 provides the statistics of the H-tail volume coefficient, CHT, within the range 0.5 < CHT < 1.2. In this example, CHT = 0.7 is appropriate for the smaller aircraft variant. The H-tail is placed as a T-tail (dominant for smaller aircraft to increase the tail arm). The H-tail sweepback is 15 deg, in line with the wing sweep, and slightly more to increase the tail arm LVT = 7.62 m (25 ft) measured from the aircraft CG to the H-tail MAC. In general, SHT/SW ^ 0.2 to 0.25. The aerofoil section is the NACA64-210 and the installation is inverted. The H-tail span equals 16.7 ft (5.1 m) and the taper ratio equals 0.5. Equation 3.30 gives the H-tail reference area, Sht = (Cht)(Sw x MAC)/Lht.

The H-tail is positioned to give SHT = (0.7 x 30 x 2.132)/7.62 = 5.88 m2 (63.3 ft2), which is about 20% of the wing area. This area must be shared by the elevator and the stabilizer. Typically, the elevator uses 18 to 25% of the H-tail area; in this case, it is 20%, which results in an elevator area of 1.21 m2 (13 ft2).

This would result in sensible geometric details of the H-tail, as follows:

• Note: Area, SH = % (CR + CT) x b or 5.88 = 0.5 x 1.5 CR x 5.1

• Root Chord = 5.04 ft (1.54 m)

• Tip Chord = 2.52 ft (0.77 m)

• Aspect Ratio = 4.42

• MAC = (3) x [(5.04 + 2.52) – (5.04 x 2.52)/(5.04 + 2.52)] = 3.9 ft (1.19 m)

To check the CHT for the smaller variant, it should be more than 0.6. With one seat pitch plug removed from the aft fuselage, LHT_shoii = 7.62 – 0.813 = 6.807 m (22.33 ft). This gives CuT_shon = (6.063 x 6.807)/(30 x 2.2) = 0.625 (sufficient for the shorter variant).

Typically, for civil aircraft, the H-tail planform area is from one fifth to one fourth of the wing planform size. Figure 12.11 shows a cluster of H-tail designs with a tail volume coefficient of 0.7. As in wing design, the H-tail can have a sweep and a dihedral (a twist is not required). Sweeping of the H-tail would effectively increase the tail arm LHt, which is an important consideration when sizing the H-tail. For a T-tail configuration, the tail arm further increases.

6.6.1 Vertical Tail

Typically, for civil aircraft, the V-tail planform area is about 12 to 20% of the wing reference area. For propeller-driven aircraft, the V-tail could be kept slightly skewed (less than 1 deg) to offset a swirled-slipstream effect and gyroscopic torque of rotating engines and propellers. The V-tail design is critical to takeoff – espe­cially in tackling yawed ground speed resulting from a crosswind and/or asymmet­ric power of a multiengine aircraft. A large V-tail can cause snaking of the flight path at low speed, which can be resolved easily by introducing a “yaw-damper” (a matter of aircraft control analysis). At cruise, a relatively large V-tail is not a major concern.

From the statistics given in Figure 12.11, it can be seen that there is a cluster of V-tail designs with a tail volume coefficient of 0.07. For the T-tail configuration, the tail volume coefficient could be reduced to 0.06 because the T-tail acts as an endplate at the tip of the V-tail. As in wing design, the V-tail can have a sweep, but the dihedral and anhedral angles and the twist are meaningless because the V-tail needs to be symmetric about the fuselage centerline. Sweeping of the V-tail would effectively increase the tail arm LVT, an important dimension in sizing the V-tail. It is important to ensure that the V-tail, especially the rudder, is not shielded by the H-tail to retain effectiveness, especially during spin recovery. With a T-tail, there is no shielding.

The empennage design has considerable similarity to the wing design. Sec­tion 4.9 describes various types of empennage; here, only the conventional design with an H-tail and a V-tail are considered. Following is a stepwise approach to empennage design:

Step1: Decide the aerofoil section.

In general, the V-tail aerofoil section is symmetrical but the H-tail has an inverted section with some (negative) camber. The t/c ratio of the empennage is close to the wing-aerofoil considerations. A compro­mise is selected based on the aircraft design Mach number and the wing sweep chosen.

Step 2: Establish the H-tail and V-tail reference areas.

Initially, during the conceptual study, the H-tail and V-tail reference areas are established from the statistical data of the tail volume coeffi­cients (see Section 12.5). The positions of the H-tail and V-tail relative to the fuselage and the wing are decided by considering the aerody­namic, stability, control, and structural considerations.

Step 3: Establish the empennage aspect ratio, sweep, taper ratio, and dihedral. The empennage planform is generally but not restricted to a trape­zoidal shape. A strake-like surface could be extended to serve the same aerodynamic gains as for the wing. The choices for the empen­nage aspect ratio, wing sweep, and taper ratio are interlinked and fol­low the same approach as for the wing design. The empennage aspect ratio is considerably lower than that of the wing. All these param­eters are decided from stability considerations and eventually fine – tuned through CFD analysis and wind-tunnel testing, with the hope that flight-test results will not require further tweaking.

Step 4: Establish the control surfaces.

Initially, the control areas and dimensions of the elevator and the fin are earmarked from statistics and semi-empirical data. At this stage of study, the control surfaces can be postponed until more details are available to accurately size the control areas. In this book, the control surfaces are not sized. Subsequently, in the next design phase, when the finalized aircraft geometry is available, the empennage dimensions are established by formal stability analysis. A worked-out example fol­lows in the next section.

The function of the empennage is to provide a force/moment for stability and con­trol. The fuselage length, wing reference area (SV), and tail arms LHT and LVT are the main parameters governing the empennage size. Semi-empirical relations given in the definition of tail volume coefficient (see Section 12.5) provide the statistical empennage size required (see Figure 12.11).

The H-tail is placed as a T-tail on a swept-back V-tail that would provide an increased tail arm, LHT and LVT, which would save weight by not having a longer fuselage. Smaller aircraft would benefit from a T-tail; however, to support the T-tail load, the V-tail must be made stronger with a small increase in its weight. Care must be taken to ensure that the T-tail does not enter the wing wake at a high angle of attack. This can be achieved by positioning it high above the wing wake at near stall or having a larger H-tail and/or an all-moving H-tail acting as an elevator. (Earlier aircraft encountered these problems; in a deep stall, there was insufficient elevator power in the low-energy wing wake for the aircraft to recover in the pitch plane before crashing.)

Selection of the empennage aerofoil and planform follows the same logic as for the wing design. V-tail designs have symmetrical aerofoil sections. The H-tail camber is influenced by the aircraft’s CG position. In general, negative camber is used to counter a nose-down moment of the wing. H-tail and V-tail designs are discussed separately in the following subsections. The current design tendency indi­cates a little higher tail volume coefficient as compared to the historical design trend (see Figure 12.11).

Continuing with the fuselage-design example outlined in Section 6.4, following are specifications required for wing design:

Maximum Cruise Speed: Mach 0.74 (HSC)

Initial Cruise Altitude: Above 40,000 ft (ceiling more than 50,000 ft)

Takeoff Field Length: 1,000 m at sea level (balanced field length)

Landing Distance from 50 ft: 1,000 m at maximum landing weight, as high as

0.95 MTOM at sea level

Initial Rate of Climb: 16 m/s

Unlike the fuselage, the approach to wing design starts with past statistics and is properly sized in Chapter 11. Following the considerations listed in Sections 6.3.2 and 6.5.2, a wing design could progress in a stepwise approach as suggested herein.

A worked-out example follows. Figure 6.10 is specifically for the aircraft class under consideration. Aircraft in the graphs are the Century, Cessna CJ2, Cessna Excel, Cessna 650, Lear 60, Cessna 750, and Challenger.

Step1: Decide the aerofoil section.

This is one of the most important aspects of aircraft design. Aircraft performance depends considerably on the type of aerofoil adopted. Today, most designers in the major aircraft industry design their own aerofoil and keep the profile “commercial in confidence.” There are also many industries that use the established NACA-type aerofoil. This book uses the established aerofoil section available in the public domain. Aerodynamicists prefer the aerofoil to be as thin as possible, whereas structural engineers prefer it to be as thick as possible. A com­promise is reached based on the aircraft design Mach number and the chosen wing sweep.

Step 2: Establish the wing reference area.

Initially, the wing reference area must be estimated from previous statistics. First, estimate the aircraft MTOW from the payload-range capability (see Figure 4.5). Next, estimate the wing reference area, SW, from the MTOW (see Figure 6.10a); this gives the wing-loading. Both the SW and the MTOW are accurately sized in Chapter 11. Position the wing relative to the fuselage, considering the aerodynamic and struc­tural features.

Step 3: Establish the aspect ratio, wing sweep, taper ratio, dihedral, and twist (see Section 3.16).

The wing planform is generally of but not restricted to a trapezoidal shape – it can be modified with a glove and/or a yehudi. The choices for the wing-aspect ratio, wing sweep, and taper ratio are interlinked to keep the compressibility drag increase within twenty drag counts at the high-speed design specification (see Section 3.18). The aspect ratio should be the highest that the structural integrity will permit for the aerofoil t/c ratio and the wing root chord based on the taper

ratio. At this stage, wing twist is empirically determined to improve stalling effects. The wing dihedral is decided from stability consider­ations. All these parameters are eventually fine tuned through CFD analysis and wind-tunnel testing, with the hope that flight-test results will not require further tweaking.

Step 4: Establish the control surfaces (e. g., aileron and spoilers).

Initially, these are approximated by reference to statistics and semi­empirical data; the sizing could be postponed until more details are available. In this book, the control surfaces are not sized.

Step 5: Establish the high-lift devices (e. g., flap and slats).

The first task is to decide the type of high-lift device required to meet the maximum CL to satisfy the specified field performance re­quirements (i. e., takeoff and landing). Once established, the area and other geometrical parameters are initially approximated by reference to the statistics and semi-empirical data. The sizing can be postponed until more details are available. In this book, the high-lift devices are not sized.

To maintain component commonality, the wing should be the same for all three variants; obviously, it would be slightly larger for the smaller variant and slightly smaller for the larger variant. How this is determined satisfactorily is addressed in Chapter 11.

Maintaining the established design trends, the planform shape of the example is taken as trapezoidal and assembled as a low wing to the aircraft. The aerofoil for the aircraft is a NACA 65-410 (i. e., 10% t/c ratio; see Appendix C).

At this stage, the wing reference area and aircraft weight are not known. This is when the statistics of previous designs prove useful to initiate the starting point. Unfortunately, Figure 4.8 is very coarse; however, Figure 6.10 provides similar information in finer detail confined to the aircraft class. The author recommends that readers produce similar graphs in better resolution for the aircraft class under consideration.

Figure 6.10a indicates a MTOM of approximately 9,000 to 10,000 kg, corre­sponding to 10 passengers. An average value of 9,500 kg (21,000 lb) is used for the example. The corresponding wing area is «30 m2 (322.9 ft2) of trapezoidal wing planform, which gives a wing-loading of 316.67 kg/m2 (65 lb/ft2). These are pre­liminary values and are formally sized in Chapter 11. However, the aspect ratio is reduced to 7.5 to keep the OEW light (it will be iterated). A taper ratio of 0.4 is used, which reduces the wing span. With a relatively low LRC Mach of 0.65, the compressibility effect is low and a quarter-chord sweep angle of 14 deg (see Figure 3.36) would keep the wave drag to zero.

The wing span is worked out as b = V(AR x SW) = V225 = 15 m (49.2 ft). The wing root and tip chord (CR and CT) can now be worked out from the taper ratio of 0.4:

CT/CR = 0.4 and SW = 30 = b x (CT + CR)/2, solving the equations CR = 2.86m (9.38 ft) and Ct = 1.143 m (3.75 ft).

Using Equation 3.21, the wing MAC = | x [2.87 + 1.148 – (2.87 x 1.148)/ (2.87 + 1.148)] = 2.132 m (7 ft). Figure 6.11 gives the wing plan form geometry.

It is interesting that most typical values of taper, twist, and dihedral are derived from statistics and are about the same for the aircraft class. From the statistics, a twist of -2 deg (i. e., washout) and a dihedral of 3 deg are typical for the class. Even­tually, CFD and wind-tunnel testing will fine-tune the values. A wing-loading of 316.67 kg/m2 (3,106.5 N/m2) is a moderate value that would provide good field per­formances. A single-slotted Fowler flap without a LE slat would be sufficient, saving considerably on costs.

Control areas are provisional and are sized in Phase 2. Initially, a company’s statistical data of previous experience serve as a good guideline. Aileron, flaps, and spoilers are placed behind the wing rear spar, which typically runs straight (or piece­wise straight) at about 60 to 66% of the chord. With a simple trapezoidal wing plan – form, the rear spar runs straight, which keeps manufacturing costs low and the oper­ation simpler; therefore, it has a lower maintenance cost. With a third of the wing span exposed, the aileron area per side is about 1 m2 (10.764 ft2). Similarly, the flap area is 2.2 m2 (23.68 ft2) per side. Subsequent performance analysis would ascertain whether these assumptions satisfy field-performance specifications. If not, further iterations with improved flap design are carried out.

From the test data, the following maximum lift coefficients are given:

Flap deflection – deg 0 8 20 40

CLmax 1.5 1.7 1.9 2.1

For a small aircraft with limited ground clearance, the engines would be mounted on the rear fuselage. At this stage, the wing is placed just behind the middle of the fuse­lage. The wing location is subsequently fine-tuned when the CG and undercarriage positions are known. A smaller aircraft wing could be manufactured in one piece and placed under the fuselage floorboards, minimizing a “pregnant-looking” fairing (Figure 3.35 shows a generous fairing to smooth the hump; however, the example in this book has more streamlined fairing).

Positioning of the wing relative to the fuselage is an iterative process dictated by the location of the aircraft CG at a desired position, expressed in terms of percent­age of the wing MAC. The aircraft CG is kept close to the quarter-chord position of the wing MAC. Unfortunately, at this stage of design, the aircraft weight and CG are not accurately known.

A designer’s expertise is the way to estimate the wing position relative to the fuselage as a starting point. Experienced designers minimize the number of itera­tions that could occur with “wing-chasing,” explained in Section 4.11. The CG posi­tion varies with aircraft loading, fuel status, and military aircraft armament carried. Positioning of the wing should be such that the aircraft stability margin is not jeop­ardized by extremes of the operational CG position.

For newcomers to aircraft design, this offers an interesting exercise: Very quickly, a “feel” for locating the wing can be developed. A starting position for wing placement relative to the fuselage is approximately at the middle of the fuse­lage (somewhat farther behind for aft-mounted engines).