Category AIRCRAF DESIGN

Engine thrust = 17,230 N (3,800 lb) per engine with a BPR < 4.

Mnac+pylon = 6.2 x thrust (kN) per nacelle = 6.2 x 17.23 = 106 kg per nacelle Two nacelles = 212 kg.

8.11.2 Undercarriage Group Mass

MTOM = 9,500 kg low-wing mount; MU/C_wing = 0.04 x 9,500 = 380 kg

8.11.3 Miscellaneous Group Mass

Fortunately, there are no miscellaneous structures in the examples considered herein.

8.11.4 Power Plant Group Mass

This is determined from statistics until it is sized in Chapter 10. A typical engine is of the class Allison TFE731-20 turbofan with thrust per engine = 15,570 N to 17,230 N (3,500 to 3,800 lbs).

If a manufacturer’s dry weight is available, it is better to use it rather than semi-empirical relations. In this case, the manufacturer’s dry weight is in the public domain. MdrYeng = 379 kg per engine. This gives:

(T/ Wdry_e„gine) = 17,230/(379 x 9.81) = 4.63

The total power plant group mass can be expressed semi-empirically as

Meng per engine = 1.5 x MdrYeng per engine = 1.5 x 379 = 568.5 kg.

For two engines, Meng = 1,137 kg.

Consider a 10% composite secondary structure; that is:

kmat = (0.9 + 0.9 x 0.1) = 0.99

It has no slat, making ksl. = 1, and without a winglet, kwl = 1. For the spoiler, ksl = 1.001, and for a wing-mounted undercarriage, kuc = 1.002.

SW = 30 m2, nult = 4.125, b = 15 m, AR = 6.75, X = 0.375, fuel in wing, MWR = 1,140kg, Л = 14°, t/c = 0.105, and Vd = 380knots = 703.76kmph = 195.5m/s. The load factor n = 3.8.

Equation 8.21 becomes:

Mw = 0.0215 x 0.99 x 1.002 x 1.001 x (9,500 x 4.125)0 48 x 29078

x 6.75 x (1 + 0.375)0 4 x (1 – 1,140/9,500)a4/(Cos14 x 0.105a4)

Mw = 0.0213 x 1.003 x 160.2 x (13.8 x 6.75 x 1.136) x 0.88a4/(0.97 x 0.406) Mw = 3.42 x 105.8 x 0.95/(0.3977) = 351.8/0.3977 = 864 kg

8.11.1 Empennage Group Mass

For a H-tail, a conventional split tail has a kco„f = 1.0. Consider a 20% composite secondary structure; that is:

kmat = (0.8 = 0.9 x 0.2) = 0.98

MTOM = 9,500 kg, nuit = 4.125, Sht = 5.5 m2 (exposed), AR = 3.5, X = 0.3,

Л = 16°, t/c = 0.105, and VD = 380 knots = 703.76 kmph = 195.5 m/s.

Mht = 0.02 x 0.98 x (9,500 x 4.125)0484 x 5.50 78 x 3.5 x (1 + 0.3)°’4/(Cos16 x 0.10504)

= 0.0196 x 167.1 x 3.8 x 3.5 x 1.11/(0.961 x 0.406) = 47.7/0.39 = 124kg

For a V-tail (T-tail), kconf = 1.1. Consider a 20% composite in a secondary struc­ture; that is:

kmat = (0.8 + 0.9 x 0.2) = 0.98

MTOM = 9,200kg, nuit = 4.125, Svt = 3.5m2(exposed), AR = 2.0, X = 0.5,

Л = 20°, t/c = 0.105, and VD = 380 knots = 703.76 kmph = 195.5 m/s Mvt = 0.0215 x 0.98 x 1.1 x (9,500 x 4.125)0484 x 3.5078 x 2 x (1 + 0.5)04/ (Cos20 x 0.10504)

= 0.02318 x 167.1 x 2.66 x 2 x 1.176/(0.94 x 0.406)

= 23.877/0.382 = 63 kg

Figure 8.5 provides the graphical solution. It reads i/Wf = 0.02 (top line corre­sponding to span, b = 49.2 ft). Therefore, wing weight i = 22,000 x 3.8 x 0.021 = 1,756 lb = 800 kg, a difference of about 6%.

Consider a 5% weight reduction due to composite usage in nonload-bearing struc­tures (e. g., floorboards). Use Equation 8.15:

L = 15.24 m, Dave = 1.75 m, VD = 380KEAS = 703.76 kmph = 195.5 m/s and Cfus = 0.04, ke = 1.04, kp = 1.09, kuc = 1.06, and kmat = (0.9 + 0.9 x 0.1) = 0.99

fuselage

horizonta tail

Then:

MFcivil = cfus X ke X kp X kuc X kdoor X; (MTOM X nult) X (2 X L X Dave X

For civil aircraft, (MTOM x nult)x = 1; therefore:

M-Fcivii = 0.04 X 1.04 X 1.09 X 1.06 x 1 x 1 x (2 x 15.24 x 1.75 x 195.50’5)1’5 = 0.048 x (2 x 15.24 x 1.75 x 195.50’5)1’5 = 0.048 x (53.35 x 13.98)15 = 0.048 x (745.95)15 = 0.048 x 20,373.3 = 978 kg

There is a 5% reduction of mass due to the use of composites:

MFcivil = °.95 MFcivil-all metal ^ 930 kg

This is checked using Torenbeek’s method (Equation 8.13); refer to Chap­ters 6 and 9 for dimensions:

WFcivil = 0.021 Kf {VdLht/ (W + D)}05 (f s-gross _area )

Kf = 1.08, L = 50 ft, W = 5.68 ft, D = 5.83 ft, Lht = 25 ft, Vd = 380 KEAS,

Sfus „gross – area = 687 ft

Therefore,

WFcivil = 0.021 X 1.08 X 1.07 X {380 x 25/(5.68 + 5.83)}0 5 x (687)12

= 0.0243 x (825.4)0 5 x 2,537.2 = 0.0243 x 28.73 x 2,537.2 = 1,770 lb (805 kg)

The higher value of the two (i. e., 930 kg) is retained, which gives a safer approach initially.

The fuel load is mission-specific. For civil aircraft, required fuel is what is needed to meet the design range (i. e., market specification) plus mandatory reserve fuel. It can be determined by the proper performance estimation described in Chapter 13. At this design stage, statistical data are the only means to estimate fuel load, which is then revised in Chapter 13.

The payload and fuel mass are traded for off-design ranges; that is, a higher payload (if accommodated) for less range and vice versa.

8.11 Worked-Out Example – Civil Aircraft

The semi-empirical relations described in Section 8.10 are now applied to obtain an example of the configuration worked out in Chapters 6 and 7 in the preliminary configuration layout. This chapter more accurately estimates component and air­craft mass along with the CG locations (Figure 8.4). Therefore, the preliminary con­figuration needs to be refined through an iterative process with more accurate data. The iteration process may require the repositioning of aircraft components (see Sec­tion 8.13). The aircraft configuration is finalized in Chapter 11. From Chapter 6, the following specifications are obtained for the baseline-aircraft preliminary configura­tion. They are required to estimate aircraft component mass, as shown here:

MTOM = 9,500 kg (refined in this exercise)

Two turbofans (i. e., Honeywell TFE731), each having TSLS = 17,235 N (3,800 lbs), BPR < 4 and dry weight of 379 kg (836 lbs)

The results from this section are compared with the graphical solutions in Figure 8.3 and in Table 8.5.

A civil aircraft payload is basically the number of passengers at 90 kg per person plus the cargo load. The specification for the total payload capacity is derived from the operator’s requirements. The payload for cargo aircraft must be specified from market requirements.

This group includes the seats, galleys, furnishings, toilets, oxygen system, and paint (see Section 8.6.1). At the conceptual design stage, they are grouped together to obtain the furnishing group.

MfUr = 0.07 to 0.08 x MTOW for large aircraft > 100passengers (8.49) MfUr = 0.06 to 0.07 x MTOW for smaller transport aircraft of

< 100 passengers (8.50)

MfUr = 0.02 to 0.025 x MTOW for unpressurized aircraft (8.51)

8.10.9 Contingency and Miscellaneous – Civil Aircraft

A good designer plans for contingencies; that is:

Mcont = (0.01 to 0.025) x MTOW (8.52)

Miscellaneous items should also be provided for; that is:

Mmisc = 0to1% of MTOW (8.53)

8.10.10 Crew – Civil Aircraft

A civil aircraft crew consists of a flight crew and a cabin crew. Except for very small aircraft, the minimum flight crew is two, with an average of 90 kg per crew member. The minimum number of cabin crew depends on the number of passengers. Opera­tors may employ more than the minimum number, which is listed in Table 8.4.

Carefully examine which structural parts are omitted (e. g., delta fin). Use mass per unit area for a comparable structure (i. e., a lifting surface or a body of revolution; see Section 8.4). If any item does not fit into the standard groups listed herein, then it is included in this group. Typically, this is expressed as:

Mmisc — 0 to 1% of the MTOM

8.10.2 Power Plant Group – Civil Aircraft

The power plant group consists of the components listed in this section. At the con­ceptual design stage, they are grouped together to obtain the power plant group mass. It is better to use the engine manufacturer’s weight data available in the public domain. However, given here are the semi-empirical relations to obtain the engine weight.

Turbofans

(1) Equipped dry-engine mass (ME)

(2) Thrust-reverser mass (MTR), if any – mostly installed on bigger engines

(3) Engine control system mass (MEc)

(4) Fuel system mass (MFS)

(5) Engine oil system mass (MOi)

Turboprops

(1) Equipped dry-engine mass (ME) – includes reduction gear mass to drive pro-

peller

(2) Propeller (MPR)

(3) Engine control system mass (MEc)

(4) Fuel system mass (MFS)

(5) Engine oil system mass (MOi)

Piston Engines

(1) Equipped dry-engine mass (ME) – includes reduction gear, if any

(2) Propeller mass (MPR)

(3) Engine control system mass (MEc)

(4) Fuel system mass (MFS)

(5) Engine oil system mass (MOi)

In addition, there could be a separate auxiliary power unit (APU) – generally in bigger aircraft – to supply electrical power driven by a gas turbine.

Engine manufacturers supply the equipped dry-engine mass (e. g., fuel pump and generator) and the engine thrust-to-weight ratio (T/Mdry. engine’, thrust is mea­sured in Newtons) as a measure of dry-engine weight in terms of rated thrust (TSLS). Typically, T/Mdryengine varies between 4 and 8 (special-purpose engines can be more than 8). For turboprop engines, the mass is expressed as (SHP/Mdryengine); for piston engines, it is (HP/Mdry. engine).

The remainder of the systems including the thrust reverser (for some turbofans), oil system, engine controls, and fuel system are listed here. The total power plant group mass can be expressed semi-empirically (because of the similarity in design, the relationship is fairly accurate). The power plant group mass depends on the size of the engine expressed by the following equations:

Turbofan

Civil aircraft power plant (with no thrust reverser):

MENG-tf = 1-4 x MdryENG per engine Civil aircraft power plant (with thrust reverser):

MENGjtf = 1-5 x MdryENG per engine

Turboprop

Civil aircraft power plant:

MENG-tp = ktp X MdryENG per engine

where 1-4 < ktp < 1-5-

Piston Engine

Civil aircraft power plant:

MenGp = kp X MdryENG per engine

where 1-4 < kp < 1-5-

APU (if any)

8.10.8 Systems Group – Civil Aircraft

The systems group includes flight controls, hydraulics and pneumatics, electrical, instrumentation, avionics, and environmental controls (see Section 8.6.1). At the conceptual design stage, these are grouped together to obtain the power plant group.

Msys = 0.1to0.11 x MTOWfor large aircraft > 100passengers MSys = 0.11 to 0.12 x MTOW for smaller transport aircraft of < 100 passengers

Msys = 0.05 to 0.07 x MTOW for unpressurized aircraft

Chapter 7 describes undercarriages and their types in detail. Undercarriage size depends on an aircraft’s MTOM. Mass estimation is based on a generalized approach of the undercarriage classes that demonstrate strong statistical relations, as discussed herein.

Tricycle Type (Retractable) – Wing-Mounted (Nose and Main Gear Estimated Together)

For a low-wing-mounted undercarriage:

Tricycle Type (Retractable) – Fuselage-Mounted (Nose and Main Gear Estimated Together)

These are typically high-wing aircraft. A fuselage-mounted undercarriage usually has shorter struts.

Muc_fus — 0.04 X MTOM

For a fixed undercarriage, the mass is 10 to 15% lighter; for a tail-dragger, it is 20 to 25% lighter.

H – and V-tails also are lifting surfaces and use semi-empirical equations similar to those used for the wing. The empennage does not have an engine or undercarriage installation. It may carry fuel, but in this book, fuel is not stored in the empennage. The drivers are the same as those in the wing group mass.

Equation 8.20 is modified to suit the empennage mass estimation. Both the H – tail and V-tail plane mass estimations have a similar form but they differ in the values of constants used.

MEMPcivil = 0-0213 x (MTOM x nult)0-48 x Sw0-78 x AR x (1 + k)0-4/(CosA x t/c0-4)

(8.23)

If nonmetals are used, then mass changes by the factor of usage. For example, x% mass is nonmetal that is y% lighter, the component mass would be as follows:

MEMPcivil nonmetal — MEMPcivil – x/y x MEMPcivil + x x MEMPcivil (8.24)

In a simpler form, if there is reduction in mass due to lighter material, then the mass is reduced by that factor. If there is a 10% mass saving, then:

MEMcivil_nonmetal — 0-9 MEMcivil „all metal

Writing the modified equations in terms of this book’s nomenclature, Equa­tion 8.23 is changed to the empennage for an H-tail and a V-tail as follows. For all H-tail movement, use kconf = 1.05; otherwise, 1.0.

Mht = 0-02 x kconf x (MTOM x nult)0 48 x Sw0-78 x AR

x (1 + k)0-4/(CosA x t/c0-4) (8.26)

For V-tail configurations, use kconf = 1.1 for a T-tail, 1.05 for a midtail, and 1.0 for a low tail.

MVT = 0.0215 x kconf x (MTOM x nult)0-48 x Sw0-78 x AR x (1 + k)0-4/(CosA x t/c0-4)

8.10.1 Nacelle Group – Civil Aircraft

The nacelle group can be classified distinctly as a pod that is mounted and interfaced with pylons on the wing or fuselage, or it can be combined. The nacelle size depends on the engine size and type. The nacelle mass semi-empirical relations are as follow.

Jet Type (Includes Pylon Mass)

For a BPR greater than 4.0, MNAC_jet = 6.7 x thrust (kN) per nacelle. (8.28)

For a BPR less than 4.0, MNAC_jet = 6.2 x thrust (kN) per nacelle. (8.29)

Turboprop Type

Pods are slung under the wing or placed above the wing with little pylon, unless it is an aft-fuselage-mounted pusher type (e. g., Piaggio Avanti). For the same power, turboprop engines are nearly 20% heavier, requiring stronger nacelles; however, they have a small or no pylon.

For a wing-mounted turboprop nacelle:

MNAC_pr0p = 6.5 x SHP per nacelle (8.30)

For a turboprop nacelle housing an undercarriage:

MNAC^prop-U. c = 8 x SHP per nacelle (8.31)

For a fuselage-mounted turboprop nacelle with a pylon:

MNAC_prop = 7 x 4 x SHP per nacelle (8.32)

Piston-Engine Nacelle

For tractor types, the nacelle is forward of the engine bulkhead; for pusher types, it is aft of the engine bulkhead – both have an engine mount. This mass is not considered a fuselage mass, even when it is an extension of the fuselage mould line.

For a fuselage-mounted, piston-engine nacelle:

Mnacpiston = 0.4 x HP per nacelle (8.33)

For a wing-mounted, piston-engine nacelle:

Mnacpiston = 0.5 x HP per nacelle (8.34)

If a nonmetal is used, then mass changes by the factor of usage. For example, x% mass is nonmetal that is y% lighter, the component mass would be as follows:

The wing is a thin, flat, hollow structure. The hollow space is used for fuel storage in sealed wet tanks or in separate tanks fitted in; it also houses control mechanisms – accounted for separately. As an option, the engines can be mounted on the wing. Wing-mounted nacelles are desirable for wing-load relief; however, for small turbo­fan aircraft, they may not be possible due to the lack of ground clearances (unless the engine is mounted over the wing or it is a high-wing aircraft – few are manufac­tured).

The drivers for the wing group mass are its planform reference area, Mt); aspect ratio, AR(t); quarter-chord wing sweep, A/(t); wing-taper ratio, A(t); mean-wing t/c ratio, (q); maximum permissible aircraft velocity, V(t); aircraft limit load, n(t); fuel carried, (q); and wing-mounted engines, (q). The aspect ratio and wing area give the wing span, b. Because the quarter-chord wing sweep, Л/, is expressed in the cosine of the angle, it is placed in the denominator, as is the case with the t/c ratio because the increase in the t/c ratio decreases the wing weight by having better stiffness.

A well-established general analytical wing-weight equation published by SAWE [2] is as follows (others are not included):

Mw — K( MdgNz)x1 Swx2 ARx3(t/c)x4(1 + Л)х5(^Лі/4)х6(В/С)Х7 Sc/8 (8.19)

where C — wing-root chord, B — width of box beam at wing root, SCS — wing – mounted control-surface reference area, and Mdg — MTOM.

The equation is modified for coursework. The term (MdgNZ)x1 in this book’s nomenclature is (MTOM x nult)0 48. The term (B/C)tx7 SCSx8 is replaced by the factor

1.5 and included in the factor K. The lift load is upward; therefore, mass carried by the wing (e. g., fuel and engines) would relieve the upward bending (like a bow), resulting in stress relief that saves wing weight. Fuel is a variable mass and when it is emptied, the wing does not get the benefit of weight relief; but if aircraft weight is reduced, the fixed mass of the engine offers relief. Rapid methods should be used to obtain engine mass for the first iteration.

Writing the modified equation in terms of this book’s notation, Equation 8.19 is replaced by Equation 8.20 in SI (the MTOM is estimated; see Chapter 6):

Mw — cw X kuc X ksl X ksp X kwl X kre x (MTOM X Hult)0’48 X SW78 X Ar

X (1 + k) X (1 WFuel_massjH_ 0)

where cw — 0.0215 and flaps are a standard fitment to the wing.

kuc — 1.002 for a wing-mounted undercarriage; otherwise, 1.0

ksl — 1.004 for the use of a slat

ksp = 1.001 for a spoiler

kwi = 1.002 for a winglet (a generalized approach for a standard size) kre = 1 for no engine, 0.98 for two engines, and 0.95 for four engines (general­ized)

If nonmetal is used, then mass changes by the factor of usage. For example, x% mass is nonmetal that is y% lighter, the component mass would be as follows:

MWciviLnonmetal = Mwcivil — x/y x MWdvu + x x MWdva (8.21)

In a simpler form, if there is reduction in mass due to lighter material, then mass is reduced by that factor. If there is a 10% mass saving, then:

MWcivil_nonmetal — 0-9 x MWcivil_all metal