Category AIRCRAF DESIGN

This extended section of the book can be found on the Web site www. cambridge .org/Kundu and discusses important considerations involving typical military air­craft thrust reversers (TR) and exhaust nozzles. Associated figures include the following.

Figure 10.28. Military aircraft nozzle adjustment scheme (from [19])

(a) Mechanism for nozzle adjustment

(b) Individual petal movement

Figure 10.29. Supersonic nozzle area adjustment and thrust vectoring

10.10 Propeller

Aircraft flying at speeds less than Mach 0.5 are propeller-driven, larger aircraft are powered by gas turbines, and smaller aircraft are powered by piston engines. More advanced turboprops have pushed the flight speed to more than Mach 0.7 (e. g., the Airbus A400). This book discusses conventional types of propellers that operate at a flight speed of less than or equal to Mach 0.5. After introducing the basics of propeller theory, this section concentrates on the engineering aspects of what is required by aircraft designers. References [16], [18], and [22] may be consulted

Figure 10.30. Aircraft propeller

for more details. It is recommended that certified propellers manufactured by well – known companies are used.

Propellers are twisted, wing-like blades that rotate in a plane normal to the aircraft (i. e., the flight path). The thrust generated by the propeller is the lift compo­nent produced by the propeller blades in the flight direction. It acts as a propulsive force and is not meant to lift weight unless the thrust line is vectored. It has aerofoil sections that vary from being thickest at the root to thinnest at the tip chord (Fig­ure 10.30). In rotation, the tip experiences the highest tangential velocity.

A propeller can have from two blades to as many as seven or eight blades. Smaller aircraft have two or three blades, whereas larger aircraft can have from four to seven or eight blades. Propeller types are shown in Figure 10.31 with associated geometries and symbols used in analysis (see Section 10.10.1). The three important angles are the blade pitch angle, в; the angle subtended by the relative velocity, <p; and the angle of attack, a = (в – <p). Also shown in the figure is the effect of both coarse and fine propeller pitch, p. When a propeller is placed in front of an aircraft, it is called a tractor (Figure 10.21a); when it is placed aft, it is called a pusher (see Figure 3.47). The majority of propellers are the tractor type.

Blade pitch should match the aircraft speed, V, to keep the blade angle of attack a producing the best lift. To cope with aircraft speed changes, it is benefi­cial for the blade to rotate (i. e., varying the pitch) about its axis through the hub to maintain a favorable a at all speeds. This is called a variable-pitch propeller. For pitch variation, the propeller typically is kept at a constant rpm with the assis­tance of a governor, which is then called a constant-speed propeller. Almost all air­craft flying at higher speeds have a constant-speed, variable-pitch propeller (when

Figure 10.31. Multibladed aircraft propellers operated manually, it is в-controlled). Smaller, low-speed aircraft have a fixed pitch, which runs best at one combination of aircraft speed and propeller rpm. If the fixed pitch is intended for cruise, then at takeoff (i. e., low aircraft speed and high pro­peller revolution), the propeller is less efficient. Typically, aircraft designers pre­fer a fixed-pitch propeller matched to the climb – a condition between cruise and takeoff – to minimize the difference between the two extremes. Obviously, for high­speed performance, the propeller should match the high-speed cruise condition. Fig­ure 10.32 shows the benefits of a constant-speed, variable-pitch propeller over the speed range.

The в-control can extend to the reversing of propeller pitch. A full reverse thrust acts as all the benefits of a TR described in Section 10.9. The pitch can be controlled to a fine pitch to produce zero thrust when an aircraft is static. This could assist an aircraft to the washout speed, especially on approach to landing.

When an engine fails (i. e., the system senses insufficient power), the pilot or the automatic sensing device elects to feather the propeller (see Figure 10.30). Feather­ing is changing в to 75 to 85 deg (maximum course) when the propeller slows down to zero rpm – producing a net drag and thrust (i. e., part of the propeller has thrust and the remainder has drag) of zero.

Figure 10.32. Comparison of a fixed – pitch and a constant-speed, variable – pitch propeller

Windmilling of the propeller is when the engine has no power and is free to rotate, driven by the relative air speed of the propeller when the aircraft is in flight. The в angle is in a fine position.

For coursework on the Bizjet, the following factors are taken from the turbofan class and type:

Uninstalled Tsls = 15.6 kn (3,500 lb)

Bare engine length = 1.547 m (5.07 ft)

Maximum diameter = 1.074 m (3.52 ft)

Engine airmass flow at takeoff = 66.2 kg/s (146 lb/s)

Engine airmass flow at maximum cruise rating « 20 kg/s (44 lb/s)

BPR = 4

Fan-face diameter = 0.716 m (2.35 ft)

This results in the following nacelle dimensions (factors are from industrial experience).

Intake Geometry (see Section 10.8.1)

• Highlight diameter = 0.9 x 0.716 m = 0.644 m (2.11 ft)

• Throat diameter (use the contraction ratio 1.12) = 0.575 m (1.89 ft), ATh = 0.26 m2 (2.8 ft2)

• Check air velocity at the throat for the maximum cruise condition at Mach 0.74

(716 ft/s) and at 41,000-ft altitude (pTO = 17,874 N/m2, TTO = 216.65 K, and =

0.284 kg/m3)

• Therefore, VTh = 20/(0.284 x 0.26) = 270 m/s (689 ft/s); i. e., Mach 0.922 is a preferred number

• This could result in low diffuser length: Ldiff = 0.65 x Dfan = 0.65 x 0.716 = 0.465 m (1.5 ft)

Lip Section (Crown Cut)

• Use (b + d) = 0.18 x Ldiff = 0.18 x 0.465 = 0.0837 m (0.275 ft)

• Use b = 1.4 x d; this results in b = 0.05 m (0.164 ft) and d = 0.0337 m (0.11 ft)

• Use a lower-lip fineness ratio of a/b = 2; this gives a = 0.1m (0.328 ft)

• Use an upper-lip fineness ratio of c/d = 4; this gives c = 0.134 m (0.44 ft)

• The intake length from the highlight (low-speed aircraft) = 0.465 + 0.1 = 0.565 m (1.85 ft)

• The Lfb is about 1 m from the highlight, which is 1.4 times Dfan = 1.4 x 0.716 = 1m(3.28ft)

• The crown-cut radius at LFB is 1.15 times Rfan = 1.1 x 0.716/2 = 0.41 m (1.34 ft) Lip Section (Keel Cut)

This book keeps the internal contour of the intake circular to fit with the rotating circular fan face; even the constrained Boeing 737 with a flat keel section must be circular at the fan face. Therefore, the internal contour of the keel cut is the mirror image of the crown cut about the centerline. However, the keel-cut radius at LFB is

1.4 times Rfan = 1.4 x 0.716/2 = 0.5m(1.64ft).

The intake internal contour can be finalized by taking the inflexion point at about mid Ldiff, maintaining maximum в within the range (i. e., 8 to 9 deg).

The average diameter at the maximum circumference = 0.41 + (0.5 – 0.41)/2 = 0.91m (3 ft).

Nozzle Geometry

Use a nozzle length = 0.75 x fan-face diameter = 0.75 x 0.716 m = 0.537 m (1.76ft). Once the control points for the geometry are established, the contour can be generated in CAD using splined curves (i. e., smooth fairing when drawn manually).

The total nacelle length = intake length + engine length + nozzle length = 0.565 + 1.547 + 0.537 = 2.65 m(« 8.7 ft), which is close to what was established in Chapter 6 from the statistical data.

The nacelle fineness ratio = 2.62/1.074 = 2.44 (i. e., within the range).

The author cautions that the empirical method presented from his industrial experience is coarse but nevertheless provides a representative geometry for the coursework exercise. This physical model serves as a starting point for further aero­dynamic refinement to a more slimline shape through CFD and testing. To obtain the factors used in the example, significant nacelle geometric data are required for better substantiation. (Readers must study many designs to get a sense of the factors used here.) Each industry has its own approach based on past designs (which form the basis of statistical data) to generate a nacelle geometry. In the industry, nacelle design is an involved process that includes the points addressed herein.

Civil aircraft nozzles are conical in shape, on which the TR is integrated. Small turbofan aircraft may not need a TR but regional jet (RJ) aircraft and larger use a TR. Inclusion of a TR may slightly elongate the nozzle, but this is not discussed in this book.

In general, the nozzle exit area is sized as a perfectly expanded nozzle (pe = pTO) at LRC condition; at higher engine ratings, it is pe > pTO. The exit nozzle of a long- duct turbofan does not run choked at cruise ratings. At takeoff ratings, the back pressure is high at a lower altitude; therefore, a long-duct turbofan does not need to run choked (i. e., the low-pressure secondary flow mixes in the exhaust duct). An exhaust nozzle runs in a favorable pressure gradient; therefore, its shaping results

in a relatively simpler establishment of geometrical dimensions. However, it is not simple engineering at elevated temperatures and for suppressing noise levels.

The nozzle exit plane is at the end of the engine. Its length from the turbine exit plane is 0.8 to 1.5 times the fan-face diameter. The nozzle-exit-area diameter is roughly half to three fourths of the intake-throat diameter in this study.

TRs are not required by the regulatory authorities (i. e., FAA and CAA). The com­ponents are expensive, heavy, and only applied on the ground, yet their impact on an aircraft’s operation is significant due to additional safety through better control and reduced time for stopping, especially during aborted takeoffs and other related emergencies. Most airlines want their aircraft to have TRs even with the increased DOC.

Aircraft designers must ensure that TR efflux is well controlled – there should be no adverse impingement of on aircraft surface or reingestion in the engine. Fig­ure 10.26 shows a typical satisfactory TR efflux pattern.

In general, there are two types of TRs: (1) operating on both the fan and core flow, and (2) operating on the fan flow only. The choice depends on the BPR, nacelle location, and customer specifications.

The first type, which operates on the total flow (i. e., both fan and core), is shown at the top of Figure 10.27. There are two types: (1) the sliding-port aft-door type, in

Figure 10.27. Types of thrust reversers

which the doors slide to the aft end as they open up to deflect the exhaust flow; and (2) the fixed-pivot type, in which the doors rotate to a position that deflects the exhaust flow.

The second type of TR operates on the fan flow only. There are two types: (1) the petal-cowl type, shown in the middle of Figure 10.27; and (2) the cascade-cowl type, shown at the bottom of Figure 10.27. There are two cascade types: the con­ventional type and the natural-blockage type. The Bombardier CRJ700/900 aircraft uses a petal cowl TR of the natural-blockage type. The external cowls translate back, blocking the fan flow when it escapes through the fixed cascades that reverse the flow. This design is attractive with a low parts count, scalability, easier maintenance, and a relatively higher retarding energy. The petal-cowl type operating on the fan flow is suitable for short-duct nacelles, as shown in the figure. The petal doors open on a hinge to block the secondary flow of the fan when it deflects to develop reverse thrust.

TRs are applied below 150 kts and are retracted at around 50 kts (to avoid re­ingestion of engine efflux), when the wheel brakes become effective. The choice of the TR type depends on a designer’s compromise with the available technology.

This extended section of the book can be found on the Web site www. cambridge .org/Kundu and discusses the important consideration for typical military aircraft intake design involving supersonic intakes. The associated figure is Figure 10.25.

Figure 10.25. Types of ideal supersonic intake demand conditions [21]

10.9 Exhaust Nozzle and Thrust Reverser

The thrust reverser (TR) is part of an exhaust nozzle and both are addressed in this section; an empirical sizing method for a nozzle is discussed but not the size and

design of the TR, which is a separate technology. Before explaining exhaust nozzles, it is helpful to understand TRs.

The role of a TR is to retard aircraft speed by applying thrust in the forward direction (i. e., in a reversed application). The rapid retardation by the TR applica­tion reduces the landing-field length. In a civil aircraft application, the TR is applied only on the ground. Because of its severity, certification rules require to either design for deployment in flight (e. g., Concorde) or prevent in-flight deployment. However, the latter is the more common approach. A TR reduces the wheel-brake load so there is less wear and fewer heat hazards. A TR is effective on slippery run­ways (e. g., ice and water) when braking is less effective. A typical benefit of having sufficient stopping distance at landing on an icy runway with TR application is that it reduces the field length by less than half. A midsized jet-transport aircraft stops at about 4,000 ft with a TR or at about 12,000 ft without it. Without a TR, the energy that was depleted to stop the aircraft is absorbed by the wheel brake and aerody­namic drag. Application of the TR also provides additional intake momentum drag (at full throttle), contributing to energy depletion. A TR is useful for an aircraft to go in reverse (e. g., a C17) on the ground for parking, alignments, and so forth – most aircraft with a TR do not use reverse but rather a specialized vehicle that pushes it.

A TR is integrated on the nacelle and it is the responsibility of an aircraft man­ufacturer to design it or it may be subcontracted to specialist organizations devoted to TR design. The next section introduces the TR in detail so that coursework can proceed on the nacelle without undertaking the detailed design.

This section describes an empirical approach for developing the intake contour of a podded nacelle that is sufficient for the conceptual design stage. This would

Figure 10.24. Schematic diagram of nacelle forebody section (crown cut)

generally be followed by proper refinement using extensive CFD analysis and wind – tunnel testing for substantiation. The nacelle external contour is influenced by interaction with the aircraft flow field. The simplified procedure is suitable for this coursework.

Figure 10.24 provides definitions of the design parameters required for the design of a pod-mounted subsonic civil aircraft intake. The nacelle section is sim­ilar to the aerofoil shape. The throat area is the minimum area of the intake duct and acts as a diffuser. The associated nomenclature follows (for the radius, replace D with R and the subscripts remain unchanged):

D1 — highlight diameter; the forwardmost point of the nacelle. If the keel

cut is not in the vertical plane with the crown cut, then its projection at the vertical line can be used.

Dth — throat diameter; the minimum cross-sectional area of the intake

geometry

D Tip — Dfan — the tip of the fan (supplied by the engine ^nanufacturer)

DHub — rotor-hub diameter (supplied by the engine manufacturer)

Dmax — maximum external nacelle diameter Ldiff — diffuser length, from throat to fan face

bps — nacelle forebody length; the distance from the highlight to the maxi­mum diameter, DMAX

a — semi-major axis of the internal lip

b — semi-minor axis of the internal lip

c — semi-major axis of the external lip

d — semi-minor axis of the external lip

в — internal contour wall angle (below 10 deg; better at 6 deg)

Associated areas are as follows (radius R is half of diameter D in the nomenclature):

Ai — highlighted area — n (Rl)2 ATH — throat area — n(RTH)2 ATO — free-stream cross-sectional area

To size the intake, the first parameter considered is establishing the throat area. The proper method is to obtain the maximum airmass-flow demand at takeoff and the maximum-cruise demand. If the takeoff demand requires a much larger size, then blow-in doors (which close automatically when demand drops – mostly appli­cable to military designs) are provided. Using ma as the intake airmass at the maxi­mum cruise gives:

Al = ma / (рю Цю)

The throat area is sized from the lip contraction ratio (LCR) = A1/ATH (typ­ically, from 1.05 to 1.20). LCR = 1.0 represents a sharp lip and 1.2 represents a well-rounded lip.

The highlighted diameter Di is typically 0.9 to 0.95 times the fan-face diameter. Keep the Ldiffuser = 0.6 at 1 time Dfan and LFB = 1 to 2 times Dfan (it must conform with the lip contour).

The next task is to establish the lip contour before developing a suitable aero­foil section for the intake cowl. As for the wing aerofoil, NACA developed nacelle forebody aerofoil contours. NACA 1 is a good design guideline for the external con­tour (i. e., the upper lip is nearly elliptical). In general, the lower lip (i. e., elliptical) contour is developed by the engine manufacturer and matches the upper lip.

In Figure 10.24, the nacelle lip is in the shape of a quarter-ellipse with semi­major axis a and semiminor axis b. The parameters that define the inlet-lip internal – contour geometry are (1) the LCR R1/RTH (i. e., Ai/ATH), and (2) the fineness ratio (a/b).

At the crown cut:

internal-lip fineness ratio, (a/b) = from 2 to 5 (typically 1.5 to 3.0)

external-lip fineness ratio, (c/d) = from 3 to 6 (typically 3 to 5)

At the crown cut, the lip-thickness ratio of (b + d)/Ldiff is around 15 to 20% (the lip-thickness ratio is not like the aerofoil t/c ratio because the cowl length extends beyond the fan face when the ratio decreases substantially). Typically, b is 1.5 to 2 times d.

At the keel cut, if it houses accessories, the thickness ratio is (b + d)/LdiffUser by about 20 to 30%.

The side cuts of the nacelle result from the merging of the crown cut and the keel cut. If ground clearance is a problem, the accessories are distributed around the keel and the contours are merged accordingly. This book keeps the design simple by using crown-cut geometry all around, with the understanding of actual problems.

The throat Mach number and the airmass-flow demand at maximum cruise determine the DTH. The throat Mach number must be maximized to the point to maintain the fan-face Mach number below 0.5 at the maximum cruise condition. At ATO < A1, there is precompression when associated spillage generates additive drag (see Figure 9.7). Then, long Ldiffusion is not required for internal diffusion because external diffusion has partially achieved it. At ATO = A1, there is no additive drag, but it would need longer Ldiffusi0n for internal diffusion. Figure 9.6 indicates that additive drag decreases as the MFR increases. At cruise (i. e., MFR above Mach 0.6), additive drag is minor.

At the throat, if the Mach number is high (e. g., reaches the local sonic speed), there is more loss and a longer diffuser is needed to decrease air velocity to around Mach 0.5 at the fan face. It is best to keep the average Mach number at the throat just below 1.0. Care must be taken that at yaw and/or high angles of attack, the fan-face flow distortion is minimized.

Finally, to make a proper divergent part of the subsonic intake acting as the diffuser, the internal contour shows an inflection point (at around 0.5 to 0.75 Ldiff). At that point, to avoid separation, the maximum wall angle в should not exceed 8 or 9 deg.

Typically, the nacelle length, LN, is 1.5 to 1.8 times the bare engine length, LE. The maximum diameter is positioned around 0.25 to 0.40 of the nacelle length (LFB) from the front end. The nacelle external cross-section is not purely circular but rather has a “pregnant-belly” shape at the keel cut to house engine accessories. Use the maximum radius at the crown cut as 1.1 to 1.4 times the engine fan-face radius and at the keel cut as 1.2 to 1.6 times the engine fan-face radius. The side cuts are faired between the two.

For the worked-out Bizjet example, use the following values (see Section 10.10.3):

Given fan-face diameter, Dfan = 0.716 m (2.35 ft)

R1 = 0.9 X Rfan

At maximum cruise, MFR = 1 (ATO = AH)

At cruise, MFR = 0.7 (ATO < AH)

LCR = 1.12

Dmax = 1.5 x 0.716 = 1.074 m (3.52 ft)

Lower-lip fineness ratio (a/b) = 3

Upper-lip fineness ratio (c/d) = 5

Ldiff = °.65 Dfan

(b + d)/Ldiff = 0.18

Lfb = 1.4 times Dfan

Engine manufacturers supply the data for bare engines. A bare engine may come with an exhaust duct as a nozzle that fits within the nacelle exhaust.

Engine mass-flow demand varies significantly, as shown in Figure 10.23. To size the intake area, the reference cross-section of the incoming airmass-stream tube is taken at the maximum cruise condition, as shown in Figure 10.23a, when it has a cross-sectional area almost equal to that of the highlighted area (i. e., ATO = A1; see Section 10.8.1 for nomenclature). The ratio of mass flow rates (MFR) relative to the reference condition (i. e., airmass flow at maximum cruise) is a measure of the spread the intake would encounter. At the maximum cruise condition, MFR = 1 as a result of ATO = A1. At the typical cruise condition, the intake-airmass demand is lower (MFR < 1, as shown in Figure 10.23b). At the maximum takeoff rating (MFR > 1), the intake airmass-flow demand is high; the streamlined patterns are shown in Figure 10.23c. Variations in the intake airmass-flow demand are significant.

If the takeoff airmass flow demand is high enough, then a blow-in door – which closes automatically when demand drops off – can be provided. Figure 10.23d shows a typical flow pattern at incidence at high demand, when an automatic blow-in door may be necessary. At idle, the engine continues running with little thrust generation (MFR ^ 1). At an inoperative condition, the rotor continues windmilling to mini­mize drag. If a rotor seizes due to mechanical failure, there is a considerable drag increase.

Currently, the engine (i. e., fan) face should not exceed Mach 0.5 to avoid degra­dation due to compressibility effects. At a fan-face Mach number above 0.5, the relative velocity at the fan-tip region approaches sonic speed due to the high blade – rotational speed.

The purpose of the intake is to provide engine airmass-flow demand as smoothly as possible – there should be no flow distortion at the compressor face due to sepa­ration and/or flow asymmetry. The nacelle intake-lip cross-section is designed using logic similar to the aerofoil LE cross-section – that is, the flow should not separate within the flight envelope.

Combat aircraft have engines that are integral to the fuselage, mostly buried inside; however, in cases with two engines, they can bulge out to the sides. Therefore, pods are not featured unless having more than two engines on a large aircraft is required.

Figure 10.21. Typical flight parame­ters for a fuselage-mounted turboprop installation

Figure 10.22. Engine installed in a com­bat aircraft

Figure 10.22 shows a turbofan installed on a supersonic combat aircraft. In this case, it is buried inside the fuselage with a long intake duct. The external contour of the engine housing is integral to the fuselage mould lines. The internal contours of the intake and exit nozzle are the responsibility of aircraft designers in consultation with engine manufacturers.

Early designs had the intake at the front of the aircraft: the pitot type for subson- ics fighters (e. g., the Sabrejet F86) and with a movable center body (i. e., bullet trans­lates forward and backward) for supersonic fighters (e. g., the MIG 21). The long intake duct snaking inside the fuselage below the pilot’s seat incurs high losses. The side-intake superceded the nose-intake designs. Possible choices for side intakes are described in Section 4.19 – primarily, they are either side-mounted or chin-mounted. The intake is placed on a plate above the fuselage boundary layer. A center body is required for aircraft-speed capability above Mach 1.8; otherwise, it can be a pitot intake, and boundary layer plates can act as the center body.

Web Figure 10.25 shows the various flow regimes associated with supersonic intake. To install and integrate an engine in a military aircraft, designers are faced with the same considerations as for a civil aircraft design, but the technology is more complex. Designers must make justifiable choices based on the following:

• Design the engine intake and its internal contour and compute the intake losses plus those from supersonic shock waves. Multiengines are side by side.

• Design the engine exit nozzle and its internal contour and compute the noz­zle losses. Military aircraft nozzle design is complex and addressed in Section 10.10.4.

• Suppression of exhaust temperature for a stealth aircraft incurs additional losses at the intake and the nozzle.

• Compute the compressor air-bleed for the ECS (i. e., cabin air-conditioning and pressurization, de-icing and anti-icing, and other purposes). The extent of the air-bleed is less than in a civil aircraft because there is no large cabin environ­ment to control.

• Compute the power off-takes from the engine shaft to drive the electric gener­ator and accessories (e. g., pumps).

• Substantiate to the certifying agencies that the thrust available from the engine – after deducting the off-take losses – is sufficient for the full flight envelope.

Military aircraft have excess thrust (with or without AB) to accommodate hot and high-altitude conditions and to operate from short airfields; they can climb at a steeper angle than civil aircraft.

Figure 10.23. Airflow demand at various conditions (civil and military aircraft intakes)

This section is a basic description of the subject and is intended only for coursework. The discussion highlights the technical challenges but exacting details are beyond the scope of this book. Turboprop nacelle design is subjected to the same consider­ations as the turbofan design. A turboprop nacelle is also a multifunctional system consisting of (1) an inlet, (2) an exhaust nozzle, and (3) a noise-suppression system. Thrust-reversing can be achieved by sufficiently changing the propeller pitch angle. There are two primary types of turboprop nacelles, as shown in Figure 10.19. The scoop intake can be above or below (as a chin) the propeller spinner. It is interesting that several turboprop nacelles have integrated the undercarriage mount with stor­age space in the same nacelle housing, as shown in Figure 10.19a. The other type has an annular intake, as shown in Figure 10.19b. Installation losses are on the same order as those discussed for a turbofan installation.

A turboprop’s nacelle position is dictated by the propeller diameter. The key geometric parameters for a wing-mounted turboprop installation are shown in Figure 10.20.

Figure 10.20. Typical parameters for a wing – mounted turboprop installation

Typically, there can be one fourth of the propeller-diameter gap between the fuselage and the propeller tip and between other propeller tips if there are four engines. The overhang should be as far forward as the design can accommodate (like the turbofan overhang) to reduce interference drag – at least a quarter to nearly one wing-chord length is sufficient. For a high-wing aircraft, the turboprop nacelle is generally underslung, especially if it also houses the undercarriage (e. g., the Bombardier Q400). For a low-wing aircraft, the nacelle is generally over the wing to give the propeller ground clearance. The propeller slipstream assists lift and has a strong effect on static stability; flap deployment aggravates the stability changes. Depending on the extent of wing incidence relative to the fuselage, there is some angle between the wing-chord line and the thrust line – typically, from 2 to 5 deg.

A fuselage-mounted, propeller-driven system is shown in Figure 10.21. The angle between the thrust line and the wing-chord line is the same as a wing-mounted, propeller-driven nacelle. Sometimes the propeller axis has about a 1-deg downward inclination relative to the fuselage axis. These parameters assist longitudinal sta­bility. An inclination of 1 or 2 deg in the yaw direction can counter the propeller slipstream. Otherwise, the V-tail can be inclined to counter the effect.

A piston engine nacelle on the wing follows the same logic. Older designs had a more closely coupled installation.

Nacelle design and engine integration typically are the responsibility of aircraft designers. A nacelle is a multifunctional system consisting of (1) an inlet; (2) an exhaust nozzle; (3) a thrust reverser, if required; and (4) a noise-suppression system. The goal of nacelle design is to minimize associated drag and noise and to provide a smooth airflow to the engine in all flight conditions. Therefore, the aerodynamic shaping – slimline as much as possible – is very important for aerodynamicists. Typ­ical nacelle positions in current practice are shown in Figures 4.31 through 4.33.

Except for the Concorde, all civil aircraft currently are subsonic with a maxi­mum speed of less than Mach 0.98. All subsonic aircraft use some form of a pod – mounted nacelle such that the design has become generic. Readers should note that designs with the engine buried in the wing (e. g., the Comet) are no longer practiced. Recently, with the advent of very small turbofans competing with propeller-driven engines, in some smaller jet aircraft the engine can be integrated with the fuselage instead of using pod mounts. The approach of this book continues with the domi­nant pod-mounted nacelles. Figure 10.15 shows a turbofan installed in a civil aircraft nacelle pod. An over-wing nacelle like that of the VFW614 is a possibility that has yet to be explored properly (Honda has reintroduced a jet aircraft). An under-wing nacelle is the current best practice; however, for smaller aircraft, ground-clearance issues force the nacelle to be fuselage-mounted.

There are two types of podded nacelles. Figure 10.16a shows a long-duct nacelle in which both the primary and secondary flows mix within the nacelle. The mixing increases the thrust and reduces the noise level compared to a short – duct nacelle, possibly compensating the weight gain through fuel – and cost-savings. Figure 10.16b shows a short-duct nacelle in which the bypassed cold flow does not
mix with the hot-core flow. Advantages include weight, interference drag, and cost reduction by decreasing the length of the outside nacelle casing. The length of a short-duct nacelle can vary. The length depends on a designer’s assessment; the shortest is about half of the nacelle length. Although larger nacelles can benefit from having short ducts, designers may decide on a smaller nacelle with a short duct if the engine-noise level is low.

Three typical positions of the nacelle relative to the wing are shown in Figure 10.17 (see [3] for details). The top wing represents a B747, the middle wing represents an A300, and the bottom wing represents a DC10. All nacelles are hung over well ahead of the wing to keep interference drag low, almost at zero. There is no quick answer for the degree of incidence, which is design-specific and varies for the type of installation. It depends on aerodynamic consideration, the engine position relative to the wing (e. g., how much inboard on the wing and the flexure of the wing during flight). Post-conceptual design studies using CFD and wind-tunnel and flight tests fine-tune the nacelle geometry and its positional geometry to the production standard. Readers should note the typical gap between the nacelle and the wing.

Aircraft designers must make their best compromises in positioning the engine on the wing. In the coursework, Table 10.4 may be used to position wing-mounted nacelles. The most-inboard engine should be kept at least 30 deg from the nose- wheel spray angle, as shown in Figure 10.18 (the B747 is somewhat widely spaced).

Fuselage-mounted nacelle contours are similar in design but the positioning rel­ative to the fuselage requires special consideration. A gap of at least one half of the nacelle diameter can be left between the fuselage and the nacelle. The vertical position can be close to the fuselage centerline or high up on the fuselage (see Fig­ures 4.31 and 4.33). For the coursework exercise, consider the following points for positioning the nacelle on the fuselage:

• Stay clear of the wing wake.

• Keep the exhaust flow from interfering with the empennage.

Figure 10.17. Typical position of the nacelle relative to the wing

Table 10.4. Wing-mounted nacelle position

2- engine 0.3 to 0.32 of semiwing span from the aircraft centerline

3- engine Same as 2-engine; the third engine is at the aircraft centerline

4- engine Inboard at 0.29 to 0.32 and outboard at 0.62 to 0.66 of the semi-wing span

• Keep the thrust line close to the aircraft CG to comply with the first two points.

• Keep the engine sufficiently forward to satisfy the CG position relative to the aircraft.

In the past, both the internal and external contours of a nacelle were designed by the aircraft manufacturer. Although there was no strict requirement, it gradually became convenient to develop the internal contour in consultation with or even entirely by the engine manufacturer. The external contour of a nacelle is developed by aircraft designers who match it with the lines of the internal contour. The contour of the nacelle cross-section is like that of an aerofoil except that it is not uniform all around – it may be perceived as a wrapped wing around the engine. The crown-cut section is thinner than the keel-cut section, as shown in Figure 10.16. The keel-cut section is thicker in order to house accessories and its fuller lip contour helps avoid separation at a high angle of attack. In principle, it is preferable to have circular cross-sectional areas for the intake throat area, but it may not always be possible – for example, for ground clearance. The Boeing 737 has a flat keel line in order to gain some ground clearance. In this book, the intake areas are considered to be circular.

In principle, the external contour lines of a good nacelle design are not neces­sarily symmetrical to the vertical plane. However, to keep costs down by maintain­ing commonality, many nacelles are designed to be symmetrical with the vertical plane. This allows manufacturing jigs to produce interchangeable nacelles between the port and starboard sides and to be able to minimize the essential difference at the finishing end. Efforts for the nacelle aerodynamic design (i. e., external mould-line shaping and internal contouring) have progressed to a point of diminishing returns and are approaching a generic shape.

Engine designers provide aircraft designers with the engine performance – currently, using a computer program amenable to input of the various off-takes. Air­craft designers must substantiate for the certifying agencies that the thrust available

Figure 10.18. Inboard nacelle position

Retracted lit iron

(a) Scoop Intake Figure 10.19. Typical wing-mounted turboprop installation

from the engine after deducting the losses is sufficient for the full flight envelope as specified. In hot and high-altitude conditions, it becomes critical at takeoff if the runway is not sufficiently long and/or there is an obstruction to clear. In that case, an aircraft may take off with lighter weight. Airworthiness requirements require that an aircraft maintain a minimum gradient (see Chapter 11) at takeoff with its critical engine inoperative (customer requirements may demand more than the minimum).

Following are the obligations of designers when installing an engine and inte­grating it with an aircraft:

• Generate the external and internal contours of the nacelle. Multiengines are either wing-mounted (i. e., larger aircraft) or fuselage-mounted (i. e., smaller air­craft).

• Compute the compressor air-bleed for the ECS (e. g., cabin air-conditioning and pressurization, de-icing and anti-icing, and other purposes).

• Compute power off-takes from the engine shaft to drive the electric generator, accessories, and so forth.

• Substantiate for the certifying agencies that the thrust available from the installed engine is sufficient for the full flight envelope.

Current developments involve laminar flow control over the external surface of the intake duct and technologies for noise and emission reduction.