Category AIRCRAF DESIGN

Fuselage Group – Civil Aircraft

A fuselage is essentially a hollow shell designed to accommodate a payload. The drivers for the fuselage group mass are its length, L(t); diameter, Dave (t); shell area and volume, (t); maximum permissible aircraft velocity, V(t); pressurization, (t); aircraft load factor, n(t); and mass increases with engine and undercarriage installation. The maximum permissible aircraft velocity is the dive speed explained in the V-n diagram in Chapter 5. For a noncircular fuselage, it is the average diam­eter obtained by taking half the sum of the width and depth of the fuselage; for a

rectangular cross-section (invariably unpressurized), it is obtained using the same method. Length and diameter give the fuselage shell area: the larger the area, the greater is the weight. A higher velocity and limit load n require more material for structural integrity. The installation of engines and/or the undercarriage on the fuse­lage requires additional reinforcement mass. Pressurization of the cabin increases the fuselage-shell hoop stress that requires reinforcement, and a rear-mounting cargo door is also a large increase in mass. (The nonstructural items in the fuse­lage – e. g., the furnishings and systems – are computed separately.)

Following are several sets of semi-empirically derived relations by various authors for the transport aircraft category (nomenclature is rewritten according to the approach of this book). The equations are for all-metal (i. e., aluminum) aircraft.

By Niu [6] in FPS:

WFcivil = k1k2

2,446.4

ncrw, w л Л, 1.5ДP^°’5 ”|

0.5(Wflight-gross-weight + ‘Wlanding_weight) x ( 1 + 4 ) x

lSnet. fus. wetted. area X [0.5 (W + D)]0’5 X L0’6 X 10-4 – 678 J

(8.11)

where k1 = 1.05 for a fuselage-mounted undercarriage

= 1.0 for a wing-mounted undercarriage where k2 = 1.1 for a fuselage-mounted engine

= 1.0 for a wing-mounted engine Snet-fus-wettedjarea = fuselage-shell gross area less cutouts

Two of Roskam’s suggestions are as follows [5]:

1. The General Dynamic method:

WPcivU = 10.43 (Kiniet)1A2 (qB/100)0’283 (MTOW/1,000)0’95 (L/D)071 (8.12)

where Kiniet = 1.25 for inlets in or on the fuselage; otherwise, 1.0 qD = dive dynamic pressure in psf L = fuselage length D = fuselage depth

2. The Torenbeek method:

Wpcivii = 0.021 Kf {VdLht/ (W + D)}° ‘ 5 (f s_gross_area )12 (8.13)

where Kf = 1.08 for a pressurized fuselage

= 1.07 for the main undercarriage attached to the fuselage = 1.1 for a cargo aircraft with a rear door Vd = design dive speed in knots equivalent air speed (KEAS)

LH_tail = tail arm of the H-tail Sfus. gross. area = fuselage-shell gross area

By Jenkinson (from Howe) [7] in SI:

Mfcmi = 0.039 X (2 X L x Dave x VD’5)1’5 (8.14)

The author does not compare the equations here. As mentioned previously, the best method depends on the type – weight equations show inconsistency. Toren – beek’s equation has been used for a long time, and Equation 8.14 is the simplest one.

The author suggests using Equation 8.14 for coursework. The worked-out exam­ple appears to have yielded satisfactory results, capturing more details of the tech­nology level.

MFcivil = cfus X ke X kp X kuc X kdoor X (MTOM X HuttY X (2 X L X Dave X v°’5)y,

(8.15)

where cjUs is a generalized constant to fit the regression, as follows:

cfus = 0.038 for small unpressurized aircraft (leaving the engine bulkhead for­ward)

= 0.041 for a small transport aircraft (<19 passengers)

= 0.04 for 20 to 100 passengers = 0.039 for a midsized aircraft = 0.0385 for a large aircraft = 0.04 for a double-decked fuselage = 0.037 for an unpressurized, rectangular-section fuselage

All k-values are 1 unless otherwise specified for the configuration, as follows:

ke = for fuselage-mounted engines = 1.05 to 1.07 kp = for pressurization = 1.08 up to 40,000-ft operational altitude = 1.09 above 40,000-ft operational altitude kuc = 1.04 for a fixed undercarriage on the fuselage = 1.06 for wheels in the fuselage recess = 1.08 for a fuselage-mounted undercarriage without a bulge = 1.1 for a fuselage-mounted undercarriage with a bulge kVD = 1.0 for low-speed aircraft below Mach 0.3 = 1.02 for aircraft speed 0.3 < Mach < 0.6 = 1.03 to 1.05 for all other high-subsonic aircraft kdoor = 1.1 for a rear-loading door

The value of index x depends on the aircraft size: 0 for aircraft with an ultimate load (nutt) < 5 and between 0.001 and 0.002 for ultimate loads of (nult) >5 (i. e., lower values for heavier aircraft). In general, x = 0 for civil aircraft; therefore, (MTOM x nult)x = 1. The value of index y is very sensitive. Typically, y is 1.5, but it can be as low as 1.45. It is best to fine-tune with a known result in the aircraft class and then use it for the new design.

Then, for civil aircraft (nult <5), Equation 8.15 can be simplified to:

MFcivil = cfus X ke X kp X kuc X kdoor X (2 X L X Dave X VD5)!-5 (8.16)

For the club-flying-type small aircraft, the fuselage weight with a fixed undercar­riage can be written as:

MFsmalla/c = 0-038 X 1.07 X kuc X (2 X L X Dave X У£5)!-5 (8.17)

If new materials are used, then the mass changes by the factor of usage. For example, x% mass is new material that is y% lighter; the component mass is as follows:

MFcivil _new-material — MFcivil x/y X MFcivil + x X MFcivil (8.18)

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

MFcivil — °-9 X MFciviLall metal

Semi-empirical Equation Method (Statistical)

Semi-empirical relations are derived from theoretical formulation and then refined with statistical data to estimate aircraft component mass. It is an involved process to capture the myriad detailed parts. Mass estimation using semi-empirical relations can be inconsistent until a proper one is established. Several forms of semi-empirical weight-prediction formulae have been proposed by various analysts, all based on key drivers with refinements as perceived by the proponent. Although all of the propositions have similarity in the basic considerations, their results could differ by as much as 25%. In fact, in [5], Roskam describes three methods that yield different values, which is typical when using semi-empirical relations. One of the best ways is to have a known mass data in the aircraft class and then modify the semi-empirical relation for the match; that is, first fine-tune it and then use it for the new design. For a different aircraft class, different fine-tuning is required; the relations provided in this chapter are amenable to modifications (see [5] and [6]).

For coursework, the semi-empirical relations presented in this chapter are from [2] through [7]; some have been modified by the author and are satisfactory for con­ventional, all-metal (i. e., aluminum) aircraft. The accuracy depends on how closely aligned is the design. For nonmetal and/or exotic metal alloys, adjustments are made depending on the extent of usage.

To demonstrate the effect of the related drivers on mass, their influence is shown as mass increasing by (t) and decreasing by (^) as the magnitude of the driver is increased. For example, L(t) means that the component weight increases when the length is increased. This is followed by semi-empirical relations to fit statistical data as well as possible. Initially, the MTOM must be guesstimated from statistics as in Chapter 6. When the component masses are more accurately estimated, the MTOW is revised to the better accuracy.

Graphical Method for Predicting Aircraft Component Weight: Civil Aircraft

The graphical method is based on regression analyses of an existing design. To put all the variables affecting weight in graphical form is difficult and may prove imprac­tical because there will be separate trends based on choice of material, maneuver loads, fuselage layout (e. g., single or double aisle; single or double deck), type of engine integrated, wing shape, control architecture (e. g., FBW is lighter), and so forth. In principle, a graphical representation of these parameters can be accom­plished at the expense of simplicity, thereby defeating the initial purpose. The sim­plest form, as presented in this section, obtains a preliminary estimate of component and aircraft weight. At the conceptual design stage – when only the technology level to be adopted and the three-view drawing are available to predict weights – the

Table 8.1. Smaller aircraft mass fraction (fewer than or 19 passengers -2 abreast seating)

Rapid mass estimation method: Summary of mass fraction of MTOM for smaller aircraft. A range of applicability is shown; add another ± 10% for extreme designs.

Group

Small-piston

aircraft

Agriculture

aircraft

Small aircraft 2-engine (Bizjet, utility)

1-Engine

2-Engine

(1-Piston)

(Turboprop)

(Turbofan)

Fuselage

Ffu = Mfu/MTOM

12 to 15

6 to 10

6 to 8

10 to 11

9 to 11

Wing

Fw = MW/MTOM

10 to 14

9 to 11

14 to 16

10 to 12

9 to 12

H-tail

Fht = MHT/MTOM

1.5 to 2.5

1.8 to 2.2

1.5 to 2

1.5 to 2

1.4 to 1.8

V-tail

Fvt = Mvt/MTOM

1 to 1.5

1.4 to 1.6

1 to 1.4

1 to 1.5

0.8 to 1

Nacelle

Fn = MN/MTOM

1 to 1.5

1.5 to 2

1.2 to 1.5

1.5 to 1.8

1.4 to 1.8

Pylon

Fpy = MPY/MTOM

0

0

0

0.4 to 0.5

0.5 to 0.8

Undercarriage

Fuc = Me/MTOM

4 to 6

4 to 6

4 to 5

4 to 6

3 to 5

Engine

Fuc = Muc/MTOM

11 to 16

18 to 20

12 to 15

7 to 10

7 to 9

Thrust rev.

Ftr = Mtr/MTOM

0

0

0

0

0

Engine control

Fec = Mec/MTOM

1.5 to 2.5

2 to 3

1 to 2

1.5 to 2

1.7 to 2

Fuel system

Ffs = Mfs/MTOM

0.7 to 1.2

1.4 to 1.8

1 to 1.4

1 to 1.2

1.2 to 1.5

Oil system

Fos = Mos/MTOM

0.1 to 0.3

0.25 to 0.4

0.1 to 0.3

0.3 to 0.5

0.3 to 0.5

APU

0

0

0

0

0

Flight con. sys.

Ffc = Mfc/MTOM

1.5 to 2

1.4 to 1.6

1 to 1.5

1.5 to 2

1.5 to 2

Hydr./pneu. sys.

Fhp = MHP/MTOM

0 to 0.3

0.3 to 0.6

0 to 0.3

0.5 to 1.5

0.7 to 1

Electrical

Felc = Melec/MTOM

1.5 to 2.5

2 to 3

1.5 to 2

2 to 4

2 to 4

Instrument

Fins = Mins/MTOM

0.5 to 1

0.5 to 1

0.5 to 1

0.5 to 1

0.8 to 1.5

Avionics

Fav = Mav/MTOM

0.2 to 0.5

0.4 to 0.6

0.2 to 0.4

0.3 to 0.5

0.4 to 0.6

ECS

Fecs = Mecs/MTOM

0 to 0.3

0.4 to 0.8

0 to 0.2

2 to 3

2 to 3

Oxygen

Fox = MOx/MTOM

0 to 0.2

0 to 0.4

0

0.3 to 0.5

0.3 to 0.5

Furnishing

Ffur = Mfur/MTOM

2 to 6

4 to 6

1 to 2

6 to 8

5 to 8

Miscellaneous

Fmsc = Mmsc/MTOM

0 to 0.5

0 to 0.5

0 to 0.5

0 to 0.5

0 to 0.5

Paint

Fpn = Mpn/MTOM

0.01

0.01

0 to 0.01

0.01

0.01

Contingency

Fcon = McOn/MTOM

1 to 2

1 to 2

0 to 1

1 to 2

1 to 2

MEW (%)

57 to 67

60 to 65

58 to 62

58 to 63

55 to 60

Crew

6 to 12

6 to 8

4 to 6

1 to 3

1 to 3

Consumable

0 to 1

0 to 1

0

1 to 2

1 to 2

OEM (%)

65 to 75

65 to 70

62 to 66

60 to 66

58 to 64

Payload and fuel are traded

Payload

12 to 25

12 to 20

20 to 30

15 to 25

15 to 20

Fuel

8 to 14

10 to 15

8 to 10

10 to 20

18 to 28

MTOM (%)

100

100

100

100

100

Notes: Lighter/smaller aircraft would show a higher mass fraction.

A fuselage-mounted undercarriage is shorter and lighter for the same MTOM.

Turbofan aircraft with a higher speed would have a longer range as compared to turboprop aircraft and, there­fore, would have a higher fuel fraction (typically, 2,000-nm range will have around 0.26).

prediction is approximate. However, with rigorous analyses using semi-empirical prediction, better accuracy can be achieved that captures the influence of various parameters, as listed previously.

Not much literature in the public domain entails graphical representation. An earlier work (1942; in FPS units) in [3] presents analytical and semi-empirical treat­ment that culminates in a graphical representation. It was published in the United States before the gas-turbine age, when high-speed aircraft were nonexistent; those graphs served the purpose at the time but are now no longer current. Given herein

Table 8.2. Larger aircraft mass fraction (more than 19 passengers – abreast and above seating). Rapid Mass Estimation Method: Summary of mass fraction of MTOM for larger aircraft. A range of applicability is shown; add another ± 10% for extreme designs.

RJ/Midsized aircraft 2 engines

Large aircraft turbofan

Group

Turboprop

Turbofan

2-engine

4-engine

Fuselage

Ffu = Mfu/MTOM

9 to 11

10 to 12

10 to 12

9 to 11

Wing

Fw = MW/MTOM

7 to 9

9 to 11

12 to 14

11 to 12

H-tail

Fht = MHT/MTOM

1.2 to 1.5

1.8 to 2.2

1 to 1.2

1 to 1.2

V-tail

Fvt = Mvt/MTOM

0.6 to 0.8

0.8 to 1.2

0.6 to 0.8

0.7 to 0.9

Nacelle

Fn = Mn/MTOM

2.5 to 3.5

1.5 to 2

0.7 to 0.9

0.8 to 0.9

Pylon

Fpy = MPY/MTOM

0 to 0.5

0.5 to 0.7

0.3 to 0.4

0.4 to 0.5

Undercarriage

Fuc = MUC/MTOM

4 to 5

3.4 to 4.5

4 to 6

4 to 5

Engine

Feng = M ENG/MTOM

8 to 10

6 to 8

5.5 to 6

5.6 to 6

Thrust rev.

Ftr = MTR/MTOM

0

0.4 to 0.6

0.7 to 0.9

0.8 to 1

Engine con.

Fec = M EC/MTOM

1.5 to 2

0.8 to 1

0.2 to 0.3

0.2 to 0.3

Fuel system

Ffs = Mfs/MTOM

0.8 to 1

0.7 to 0.9

0.5 to 0.8

0.6 to 0.8

Oil system

Fos = MOS/MTOM

0.2 to 0.3

0.2 to 0.3

0.3 to 0.4

0.3 to 0.4

APU

0 to 0.1

0 to 0.1

0.1

0.1

Flight con. sys.

Ffc = MFC/MTOM

1 to 1.2

1.4 to 2

1 to 2

1 to 2

Hydr./pneu. sys.

Fhp = MHP/MTOM

0.4 to 0.6

0.6 to 0.8

0.6 to 1

0.5 to 1

Electrical

Felc = Melec/MTOM

2 to 4

2 to 3

0.8 to 1.2

0.7 to 1

Instrument

Fins = MINS/MTOM

1.5 to 2

1.4 to 1.8

0.3 to 0.4

0.3 to 0.4

Avionics

Fav = MAV/MTOM

0.8 to 1

0.9 to 1.1

0.2 to 0.3

0.2 to 0.3

ECS

Fecs = M ECS/MTOM

1.2 to 2.4

1 to 2

0.6 to 0.8

0.5 to 0.8

Oxygen

Fox = MOX/MTOM

0.3 to 0.5

0.3 to 0.5

0.2 to 0.3

0.2 to 0.3

Furnishing

Ffur = MFUR/MTOM

4 to 6

6 to 8

4.5 to 5.5

4.5 to 5.5

Miscellaneous

Fmsc = MMSC/MTOM

0 to 0.1

0 to 0.1

0 to 0.5

0 to 0.5

Paint

Fpn = MPN/MTOM

0.01

0.01

0.01

0.01

Contingency

Fcon = MCON/MTOM

0.5 to 1

0.5 to 1

0.5 to 1

0.5 to 1

MEW (%)

53 to 55

52 to 55

50 to 54

48 to 50

Crew

0.3 to 0.5

0.3 to 0.5

0.4 to 0.6

0.4 to 0.6

Consumable

1.5 to 2

1.5 to 2

1 to 1.5

1 to 1.5

OEW (%)

Payload and fuel are traded

54 to 56

53 to 56

52 to 55

50 to 52

Payload

15 to 18

12 to 20

18 to 22

18 to 20

Fuel

20 to 28

22 to 30

20 to 25

25 to 32

MTOM (%)

100

100

100

100

Notes: Lighter aircraft would show higher mass fraction.

A fuselage-mounted undercarriage is shorter and lighter for the same MTOM.

Turbofan aircraft with a higher speed would have a longer range as compared to turboprop aircraft and, therefore, would have a higher fuel fraction.

Large turbofan aircraft have wing-mounted engines: 4-engine configurations are bigger.

are updated graphs based on the data in Table 8.3; they are surprisingly represen­tative with values that are sufficient to start the sizing analysis in Chapter 11. Most of the weight data in the table are from Roskam [4] with additions by the author notated with an asterisk (these data are not from the manufacturers). The best data is obtained directly from manufacturers.

In all of the graphs, the MTOW is the independent variable. Aircraft – component weight depends on the MTOW; the heavier the MTOW, the heavier

Table 8.3. Aircraft component weights data

Aircraft

MTOW

Weight (lb) Fuse Wing

Emp

Nacelle

Eng

U/C

n

Piston-engined aircraft

1. Cessna182

2,650

400

238

62

34

417

132

5.70

2. Cessna310A

4,830

319

453

118

129

852

263

5.70

3. Beech65

7,368

601

570

153

285

1,008

444

6.60

4. Cessna404

8,400

610

860

181

284

1,000

316

3.75

5. Herald

37,500

2,986

4,365

987

830

1,625

3.75

6. Convair240

43,500

4,227

3,943

922

1,213

1,530

3.75

Gas-turbine-powered aircraft

7. Lear25

15,000

1,575

1,467

361

241

792

584

3.75

8. Lear45 class

20,000

2,300

2,056

385

459

1,672

779

3.75

9. Jet Star

30,680

3,491

2,827

879

792

1,750

1,061

3.75

10. Fokker27-100

37,500

4,122

4,408

977

628

2,427

1,840

3.75

11. CRJ200 class

51,000

6,844

5,369

1,001

1,794

5.75

12. F28-1000

65,000

7,043

7,330

1,632

834

4,495

2,759

3.75

13. Gulf GII (J)

64,800

5,944

6,372

1,965

1,239

6,570

2,011

3.75

14. MD-9-30

108,000

16,150

11,400

2,780

1,430

6,410

4,170

3.75

15. B737-200

115,500

12,108

10,613

2,718

1,392

6,217

4,354

3.75

16. A320 class

162,000

17,584

17,368

2,855

2,580

12,300

6,421

3.75

17. B747-100

710,000

71,850

86,402

11,850

10,031

34,120

31,427

3.75

18. A380 class

1,190,497

115,205

170,135

24,104

55,200

52,593

3.75

are the component weights (see Chapter 4). Strictly speaking, wing weight could have been presented as a function of the wing reference area, which in turn depends on the sized wing-loading (i. e., the MTOW) (see Chapter 11).

To use the graph, the MTOW must first be guesstimated from statistics (see Chapters 4 and 6). After the MTOW is worked out in this chapter, iterations are necessary to revise the estimation.

Figure 8.3 illustrates civil aircraft component weights in FPS units. The first pro­vides the fuselage, undercarriage, and nacelle weights. Piston-engine-powered air­craft are low-speed aircraft and the fuselage group weight shows their lightness. There are no large piston-engine aircraft in comparison to the gas-turbine type.

Figure 8.3 Aircraft component weights in pounds

The lower end of the graph represents piston engines; piston-engine nacelles can be slightly lighter in weight.

The second graph in Figure 8.3 shows the wing and empennage group weights. The piston – and gas-turbine engine lines are not clearly separated. FBW-driven con­figurations have a smaller wing and empennage (see Chapter 13), as shown in sep­arate lines with lighter weight (i. e., A320 and A380 class). The newer designs have composite structures that contribute to the light weight.

Figure 8.3 shows consistent trends but does not guarantee accuracy equal to semi-empirical relations, which are discussed in the next section.

Rapid Mass Estimation Method: Civil Aircraft

A rapid mass estimation method is used to quickly determine the component weight of an aircraft by relating it in terms of a fraction given in the percentage of maximum takeoff mass (Mi/MTOM), where the subscript i represents the ith component. With a range of variation among aircraft, the tables in this section are not accurate and serve only as an estimate for a starting point of the initial configuration described in Chapter 6. Roskam [4] provides an exhaustive breakdown of weights for aircraft of relatively older designs. A newer designs show improvements, especially because of the newer materials used.

Because mass and weight are interchangeable, differing by the factor g, wing­loading can be expressed in either kg/m2 or N/m2; this chapter uses the former to be consistent with mass estimation. To obtain the component mass per unit wing area (Mi/SW, kg/m2), the Mi/MTOM is multiplied by the wing-loading; that is, Mi/SW = (Mi/MTOM) x (MTOM/SW). Initially, the wing-loading is estimated (multiply 0.204816 to convert kg/m2 to lb/ft2).

Tables 8.1 and 8.2 summarize the component mass fractions, given in a percent­age of the MTOM for quick results. The OEM fraction of the MTOM fits well with the graphs (see Figures 4.7 and 4.8). This rapid method is not accurate and only pro­vides an estimate of the component mass involved at an early stage of the project. A variance of ±10% is allowed to accommodate the wide range of data.

It is better to use more accurate semi-empirical relations (see Section 8.10) to obtain the component mass at the conceptual design phase. The tables are useful for estimating fuel mass and engine mass, for example, which are required as a starting point for semi-empirical relations.

Military Aircraft (Combat Category)

This extended section of the book can be found on the Web site www. cambridge .org/Kundu and lists generic military aircraft-component mass as required in the conceptual design stage. The list covers aircraft components in the following groups. Structure Group Power Plant Group Systems Group Furnishing

Manufacture’s Empty Mass Operators Empty Mass Maximum Takeoff Mass Maximum Ramp Mass

8.3 Aircraft Component Mass Estimation

Mass estimation at the conceptual design stage must be predicted well in advance of detailed drawings of the parts being prepared. Statistical fitment of data from the past designs is the means to predict component mass at the conceptual design phase. The new designs strive for improvement; therefore, statistical estimation is the starting point. During the conceptual design stage, iterations are necessary when the configuration changes.

Typically, there are three ways to make mass (i. e., weight) estimations at the conceptual design stage:

1. Rapid Method. This method relies on the statistical average of mass one level below major aircraft components (i. e., in more detail). The mass is expressed in terms of percentage (alternatively, as a fraction) of the MTOM. All items should total 100% of the MTOM; this also can be expressed in terms of mass per wing area (i. e., component wing-loading). This rapid method is accomplished at the price of considerable approximation.

2. Graphical Method. This method consists of plotting component weights of vari­ous aircraft already manufactured to fit into a regression curve. Graphs are gen­erated from analytical considerations (see [3]), superimposed by actual data. The graphical method does not provide fine resolution but it is the fastest method without the next level of mass estimation, as explained previously. It is difficult to capture the technology level (and types of material) used because there is considerable dispersion. Obtaining details of component mass for sta­tistical analysis from various industries is difficult.

3. Semi-Empirical Method. This method is a considerable improvement, in that it uses semi-empirical relations derived from a theoretical foundation and backed by actual data that have been correlated statistically. The indices and factors in the semi-empirical method can be refined to incorporate the technology level and types of material used. The expressions can be represented graphically, with separate graphs for each class. When grouped together in a generalized manner, they are the graphs in the graphical method described previously.

The first two methods of component mass estimation provide a starting point for the design progression.

The state-of-the-art in weight prediction has room for improvement. The advent of solid modeling (i. e., CAD) of components improved the accuracy of the mass- prediction methodology; with CAD, weight change due to a change in material can be easily captured. As soon as the component drawing is completed, the results are instantaneous and carry on through subassembly to final assembly. CAD modeling of parts occurs after the conceptual design phase has been completed.

The design drivers for civil aircraft have always been safety and economy. Civil – aircraft design developed in the wake of military aircraft evolution. Competition within these constraints kept civil aircraft designs similar to one another. Following are general comments relative to civil aircraft mass estimation:

1. For a single-engine, propeller-driven aircraft, the fuselage starts aft of the engine bulkhead because the engine nacelle is accounted for separately. These are mostly small aircraft; this is not the case for wing-mounted nacelles.

2. The fixed-undercarriage mass fraction is lower than the retractable type. The extent depends on the retraction type (typically 10% higher).

3. Neither three-engine aircraft nor fuselage-mounted, turboprop-powered air­craft are discussed in this book. Not many of these types of aircraft are man­ufactured. Sufficient information has been provided herein for readers to adjust mass accordingly for these aircraft classes.

The three methods are addressed in more detail in the following sections.

Aircraft Component Groups

The recognized groups of aircraft components are listed in exhaustive detail in the ATA’s publication. This section presents consolidated, generalized groups (for both civil and military aircraft) suitable for studies in the conceptual design phase. Both aircraft classes have similar nomenclature; the difference in military aircraft is

described in Section 8.6.2. Each group includes subgroups of the system at the next level. Care must be taken that items are not duplicated – accurate bookkeeping is essential. For example, although the passenger seats are installed in the fuselage, for bookkeeping purposes, the fuselage shell and seats are counted separately.

8.6.1 Civil Aircraft

Structure group (MstR = Mfu + Mw + Mht + MvtMn + Mpy + Muc + MmiSC)

(8.6)

• Fuselage group (Mfu)

• Wing group (Mw): includes all structural items (e. g., flaps and winglets)

• H-tail group (Mht)

• V-tail group (Mvt)

• Nacelle group (Mn and Mpy) (nacelle and pylon)

• Undercarriage group (MuC)

• Miscellaneous (MMisC) (e. g., delta wing)

The basic structure of the aircraft – the fuselage shell (seats are listed separately under the Furnishing group) is as follows:

Power plant group (Mpp = Me + MtR + MeC + Mfs + Moi) (8.7)

• Dry-equipped engine (Me)

• Thrust reverser (MtR)

• Engine control system (MeC)

• Fuel system (Mfs)

• Engine oil system (Moi)

The power plant group comes as a package, with all items dedicated to the power plant installation. These are mostly bought-out items supplied by specialists:

Systems group (Msys = Mecs + MFc + Mhp + Melec + Mins + Mav) (8.8)

• Environmental control system (MeCs)

• Flight-control system (MFC)

• Hydraulic and pneumatic system (Mhp) (sometimes grouped with other sys­tems)

• Electrical system (MeleC)

• Instrument system (Mins)

• Avionics system (Mav)

The systems group includes a variety of equipment, all vendor-supplied, bought-out items:

Furnishing group (Mfur = Mseat + Mox + Mpn) (8.9)

• Seat, galleys, and other furnishings (Mseat)

• Oxygen system (Mox)

• Paint (Mpn)

Most of the weight is in the fuselage, yet the furnishings are itemized under different headings. Paint can be quite heavy. A well-painted B737 with airline livery can use as much as 75 kg of paint:

Contingencies (MCont)

• This is a margin to allow unspecific weight growth (Mcont).

The MEM is the total of the previous twenty-two items. This is the weight of the complete aircraft as it comes off the production line to be come airborne for the first time.

Add the following items to the MEM to obtain the OEM:

• Crew: flight and cabin crews (McREw)

• Consumables: food, water, and so forth (Mcon)

The OEM is when the aircraft is ready for operation.

Add the payload and requisite fuel to obtain the MRM. At the takeoff point at the edge of the runway, the MRM becomes the MTOM = (MRM – taxi fuel):

• Payload (Mpl) (passengers at 90 kg per passenger, including baggage)

• Fuel (MfuEl) (for the design range, which may not fill all tanks)

MTOM: The aircraft at the end of the runway is ready for takeoff. The civil – aircraft MTOM is the total weight of all component groups, as shown in Equa­tion 8.10.

The MTOM = f M(x) dx = J2Mi, where the subscript i stands for each compo­nent group listed previously.

For civil aircraft, the MTOM is equal to

( Mfu ) + ( Mw) + ( Mht) + (Mvt) + (Mn) + ( Mpy ) + ( Muc ) + ( Mmisc )

+ ( Me) + ( Mtr) + ( Mec) + ( Mfs) + ( Moi ) + ( Mecs) + ( Mfc ) + (Mhp)

+( Melec ) + ( Mins) + ( Mav ) + ( Mseat ) + (Mox ) + ( Mpn ) + (Mcont)

+ ( Mcrew ) + ( Mcons) + Mpl + Mfuel (8.10)

Desirable CG Position

Proper distribution of mass (i. e., weight) over the aircraft geometry is key to estab­lishing the CG. It is important for locating the wing, undercarriage, engine, and empennage for aircraft stability and control. The convenient method is to first esti­mate each component weight separately and then position them to satisfy the CG

Aerodynamic center range

in flight »i Vcn ground

CG range

I j-i – rotation stability – maximize

trim crag – minimize

ground

7777777

ground maneuver margin – maximize

Figure 8.1. Aircraft CG position showing stability margin location relative to overall geometry. A typical aircraft CG margin that affects air­craft operation is shown in Figure 8.1.

The aircraft aerodynamic center moves backward on the ground due to the flow field being affected by ground constraints. There is also movement of the CG loca­tion depending on the loading (i. e., fuel and/or passengers). It must be ensured that the preflight aftmost CG location is still forward of the in-flight aerodynamic center by a convenient margin, which should be as low as possible to minimize trim. Where the main-wheel contact point (and strut line) is aft of the aftmost CG, the subtend­ing angle, в, should be greater than the fuselage-rotation angle, a, as described in Section 7.6. The main wheel is positioned to ease rotation as well as to assist in good ground handling.

Advanced military combat aircraft can have relaxed static stability to provide quicker responses. That is, the margin between the aftmost CG and the in-flight aerodynamic center is reduced (it may be even slightly negative), but the other design considerations relative to the undercarriage position are the same.

Initially, locations of some of the components (e. g., the wing) were arbitrar­ily chosen based on designers’ past experience, which works well (see Chapter 6). Iterations are required that, in turn, may force any or all of the components to be repositioned. There is flexibility to fine-tune the CG position by moving heavy units (e. g., batteries and fuel-storage positions). It is desirable to position the payload around the CG so that any variation will have the least effect on CG movement. Fuel storage should be distributed to ensure the least CG movement; if this is not possible, then an in-flight fuel transfer is necessary to shift weight to maintain the desired CG position (as in the Supersonic Concorde).

Fuel loads and payloads are variable quantities; hence, the CG position varies. Each combination of fuel and payload results in a CG position. Figure 8.2 shows variations in CG positions for the full range of combinations. Because it resembles the shape of a potato, the CG variation for all loading conditions is sometimes called the “potato curve.” Designers must ensure that at no time during loading up to the MTOM does the CG position exceed the loading limits endangering the aircraft to tip over on any side. Loading must be accomplished under supervision. Whereas

10 20 30 40

Percent wing MAC (typically 10 to 50%)

(b) Range of CG variations – vertical limits

passengers have free choice in seating, cargo and fuel-loading are done in prescribed sequences, with options.

It has been observed that passengers first choose window seats and then, depending on the number of abreast seating, the second choice is made. Figure 8.2 shows the window seating first and the aisle seating last; note the boundaries of front and aft limits. Cargo – and fuel-loading is accomplished on a schedule with the locus of CG travel in lines. In the figure, the CG of the OEM is at the rear, indicating that the aircraft has aft-mounted engines. For wing-mounted engines, the CG at the OEM moves forward, making the potato curve more erect.

For static-stability reasons, it must be ensured that the aircraft has a static mar­gin at all loading conditions. With the maximum number of passengers, the CG is not necessarily at the aftmost position. Typically, the CG should be approximately 18% of the MAC when fully loaded and approximately 22% when empty. The CG is always forward of the neutral point (i. e., the aircraft’s aerodynamic center, estab­lished through CFD and wind-tunnel tests). The aerodynamic center is assumed to be 50% of the MAC and must be iterated until the final configuration is reached.

Figure 8.2 represents a typical civil aircraft loading map, which indicates the CG travel to ensure that the aircraft remains in balance within horizontal and vertical limits. Loading starts at the OEM point; if the passengers boarding first opt to sit in the aft end, then the CG can move beyond the airborne aft limit, but it must remain within the ground limit. Therefore, initial forward cargo-loading should pre­cede passenger boarding; an early filling of the forward tank fuel is also desirable.

Aircraft Mass (Weight) Breakdown

Definitions of various types of aircraft mass (i. e., weight) (see Section 4.5) are repeated here for the convenience of readers.

MEM (manufacturer’s empty mass) (8.1)

is the mass of an aircraft as it rolls out of the factory before it is taken to a flight hangar for the first flight.

OEM (operator’s empty mass) = MEM + Crew + Consumable (8.2)

The aircraft is now ready for operation (residual fuel from the previous flight remains).

MTOM (maximum takeoff mass) = OEM + Payload + Fuel (8.3)

The MTOM is the reference mass loaded to the rated maximum. This is also known as the brake release mass (BRM) ready for takeoff.

Aircraft are allowed to carry a measured amount of additional fuel for taxiing to the end of the runway, ready for takeoff at the BRM (MTOM). This additional fuel mass would result in the aircraft exceeding the MTOM to the maximum ramp mass (MRM). Taxiing fuel for midsized aircraft would be approximately 100 kg, and it must be consumed before the takeoff roll is initiated – the extra fuel for taxiing is not available for the range calculation. On busy runways, the waiting period in line for takeoff could extend to more than an hour in extreme situations.

MRM (maximum ramp mass) « 1.0005 x MTOM (very large aircraft) to

1.1 x MTOM = MTOM + fuel to taxi to end of runway for takeoff (8.4)

This is also known as the maximum taxi mass (MTM) and it is heavier than the MTOM.

ZFM (zero fuel mass) = MTOM minus all fuel (nonusable residual fuel remains)

(8.5)

The Weight Drivers

The factors that drive aircraft weight are listed herein. References [4] through [6] discuss more detail on aircraft material, stress, and structures. Aircraft material properties given herein are typical for comparing relative merits. Material elasticity, E, and density, p, provide the strength-to-weight ratio. In the alloys and material categories, there is variation.

1. Weight is proportionate to size, indicated by geometry (i. e., length, area, and volume).

2. Weight depends on internal structural-member density – that is, the denser, the heavier.

3. Weight depends on a specified limit-load factor n (see Chapter 5) for structural integrity.

4. Fuselage weight depends on pressurization, engine and undercarriage mounts, doors, and so forth.

5. Lifting-surface weight depends on the loading, fuel carried, engine and under­carriage mounts, and so forth.

6. Weight depends on the choice of material. There are seven primary types used in aircraft, as follows:

(a) Aluminum alloy (a wide variety is available – in general, the least expen­sive)

typical E = 11 x 106 lb/in2; typical density = 0.1 lb/in3

(b) Aluminum-lithium alloy (fewer types available – relatively more expen­sive)

typical E = 12 x 106 lb/in2; typical density = 0.09 lb/in3

(c) Stainless-steel alloy (hot components around engine – relatively inexpen­sive)

typical E = 30 x 106 lb/in2; typical density = 0.29 lb/in3

(d) Titanium alloy (hot components around engine – medium-priced but lighter)

typical E = 16 x 106 lb/in2; typical density = 0.16 lb/in3

(e) Composite type varies (e. g., fiberglass, carbon fiber, and Kevlar); therefore, there is a wide variety in elasticity and density (price relatively inexpensive to expensive). (For details, refer to [5] and [6].)

(f) Hybrid (metal and composite “sandwich” – very expensive; e. g., Glare).

(g) Wood (rarely used except for homebuilt aircraft; is not discussed in this book – price increasing).

In this book, the primary load-bearing structures are constructed of metal; sec­ondary structures (e. g., floorboard and flaps) could be made from composites. On the conservative side, it generally is assumed that composites and/or new alloys

comprise about 10 to 15% of the MEM for civil aircraft and about 15 to 25% of the MEM for military aircraft. The use of composites is increasing, as evidenced in current designs. Although composites are used in higher percentages, this book remains conservative in approach. All-composite aircraft have been manufactured, although only few in number (except small aircraft). The metal-composite sandwich is used in the Airbus 380 and Russia has used aluminum-lithium alloys. In this book, the consequences of using newer material is addressed by applying factors.

Aircraft Weight and Center of Gravity Estimation

8.1 Overview

An aircraft must ascend to heights by defying gravity and sustain the tiring task of cruise – naturally, it is weight-sensitive. Anyone who has climbed a hill knows about this experience, especially if one has to carry baggage. An inanimate aircraft is no exception; its performance suffers by carrying unnecessary mass (i. e., weight). At the conceptual design stage, aircraft designers have a daunting task of creating a structure not only at a low weight but also at a low cost, without sacrificing safety. Engineers also must be accurate in weight estimation, well ahead of manufacture. This chapter presents a formal method to predict an aircraft and its component mass (i. e., weight), which results in locating the CG during the conceptual design phase. The aircraft inertia estimation is not within the scope of this book.

In the past, aircraft weight was expressed in FPS units in pound (lb) weight in the United Kingdom and the United States. With the use of kg as mass in SI, the unit for weight is a Newton, which is calculated as the mass multiplied by gravitational acceleration (9.81 m/s2). This book uses both the FPS and SI systems; this chapter addresses mass in SI and weight in FPS, sometimes interchangeably.

Material strength contributes to structural integrity. As stated previously, air­craft conceptual designers must have broad-based knowledge in all aspects of tech­nology; in this case, they must have a sound knowledge in material properties (e. g., strength-to-weight and strength-to-cost ratios). Higher strength-to-weight and strength-to-cost ratios are the desired qualities, but they act in opposition. Higher strength-to-weight-ratio material is more expensive, and designers must stay cur­rent about materials technology to choose the best compromises.

In the early days, designers had no choice but to use the best quality wood for aircraft construction material. Today, it is not a viable option for the type of load encountered and it also poses an environmental issue. Fortunately, the advent of duralumin (i. e., an aluminum alloy) in the 1930s resolved the problem, providing a considerably higher strength-to-weight ratio than wood. Having a mass-produced aluminum alloy also offers a lower material cost-to-strength ratio. Wood is eas­ier to work with, having a low manufacturing infrastructure suitable for homebuilt aircraft, but other civil and military aircraft use predominantly metal alloys and

composites. The last two decades have seen a growing use of composite material, and more exotic metal alloys offer still better strength-to-weight ratios.

Composites are basically fabric and resin bonded together, generally formed to shape in moulds. The manufacturing process associated with composites is yet to achieve the quality and consistency of metal; hence, at this point, the certifying authorities are compelled to apply reduced values of stress levels to allow for dam­age tolerance and environmental issues, as well as to keep the factor of safety at

1.5 (see Section 5.6). The manufacturing process also plays a role in deciding the allowable stress level. These considerations can erode the benefits of weight sav­ings. Research on new material, whether metal alloys (e. g., lithium-aluminum and beryllium alloy) or composites (e. g., fabric and resin) or their hybrid is an area where there is potential to reduce aircraft weight and cost. New materials are still relatively expensive, and they are steadily improving in both strength and lower costs.

8.1.1 What Is to Be Learned?

This chapter covers the following topics:

Aircraft mass, component mass, and CG position Parameters that act as drivers for aircraft mass Aircraft mass breakdown sequence Desirable CG location relative to aircraft Aircraft mass decomposed into component groups Aircraft component mass estimation methods Civil aircraft rapid mass estimation method Civil aircraft graphical mass estimation method Civil aircraft semi-empirical mass estimation method Bizjet example

Methodology to establish aircraft CG with Bizjet example Military aircraft rapid mass estimation method Military aircraft graphical method for mass estimation Military aircraft semi-empirical mass estimation method AJT and CAS examples (military aircraft)

Methodology to locate aircraft CG with AJT and CAS examples

8.1.2 Coursework Content

The coursework task continues linearly with the examples worked out thus far. Readers must now estimate aircraft-component mass, which gives the aircraft mass and its CG location. This is an important aspect of aircraft design because it deter­mines aircraft performance, stability, and control behavior.

Experience in the industry has shown that weight can only grow. Aircraft per­formance is extremely sensitive to weight because it must defy gravity. Aerodynam – icists want the least weight, whereas stress engineers want the component to be strong so that it will not fail and have the tendency to beef up a structure. The struc­ture must go through ground tests when revisions may be required. It is easy to omit an item (there are thousands) in weights estimation. Most aeronautical companies

have a special division to manage weights – weights-control engineers – a difficult task to perform.

8.2 Introduction

Because aircraft performance and stability depends on aircraft weight and the CG location, the aircraft weight and its CG position are paramount in configuring an air­craft. The success of a new aircraft design depends considerably on how accurately its weight (mass) is estimated. A pessimistic prediction masks product superiority and an optimistic estimation compromises structural integrity.

Once an aircraft is manufactured, the component weights can be easily deter­mined by actual weighing. The aircraft CG then can be accurately determined. How­ever, the problem in predicting weight and the CG is at the conceptual design stage, before the aircraft is built. When the first prototype is built, the weights engineers have the opportunity to verify the predictions – typically, a 4-year wait! Many of the discrepancies result from design changes; therefore, weights engineers must be kept informed in order to revise their estimations. It is a continuous process as long as the product is well supported after the design is completed.

Mass is the product of the solid volume and average density. For an aircraft com­ponent (e. g., wing assembled from a multitude of parts and fasteners), it is a labori­ous process to compute volumes of all those odd-shaped parts. In fact, the difficulty is that the mass prediction of complex components is not easily amenable to theoret­ical derivations. The typical approach to estimate weights at the conceptual design stage is to use semi-empirical relationships based on theory and statistical data of previously manufactured component masses. (A 3-D CAD model of parts provide the volume but may not be available in the early stages of conceptual design.)

The mass of each component depends on its load-bearing characteristics, which in turn depend on the operational envelope (i. e., the V-n diagram). Each manufac­turer has a methodology developed over time from the statistics of past products combined with the physical laws regarding mass required for the geometry to sus­tain the load in question. These semi-empirical relations are proprietary informa­tion and are not available in the public domain. All manufacturers have developed mass-prediction relationships yielding satisfactory results (e. g., an accuracy of less than ±3% for the type of technology used). The semi-empirical relations of various origins indicate similarity in the physical laws but differ in associated coefficients and indices to suit their application domain (e. g., military or civil, metal or nonmetal, and level of desired accuracy). Nowadays, computers are used to predict weight through solid modeling – this is already in conjunction with semi-empirical relations. The industry uses more complex forms with involved and intricate manipulations that are not easy to work with in a classroom.

The fact is that no matter how complex academia may propose semi-empirical relations to improve accuracy in predicting component mass, it may fall short in sup­planting the relationship available in the industry based on actual data. Of necessity, the industry must keep its findings “commercial in confidence.” At best, the indus­try may interact with academia for mutual benefit. An early publication by Toren – beek [3] with his semi-empirical relations is still widely used in academic circles. Roskam [4] presented three methods (i. e., Torenbeek, Cessna, and U. S. Datcom)

that clearly demonstrate the difficulty in predicting mass. Roskam’s book presents updated semi-empirical relations, corroborated with civil aircraft data showing sat­isfactory agreement (this may be useful to homebuilt aircraft designers). The equa­tions are not complex – complexity does not serve the purpose of coursework. Read­ers will have to use industrial formulae when they join a company. This chapter explains the reasons associated with formulating the relationships to ensure that readers understand the semi-empirical relations used in the industry.

The author recommends the Society of Allied Weights Engineers (SAWE) (U. S.) as a good source for obtaining semi-empirical relations in the public domain. Some of the relations presented herein are taken from SAWE, Torenbeek, Sechler, Roskam, Niu, and Jenkinson ([2] through [7]). Some of the equations are modified by the author. It is recommended that readers collect as much component weights data as possible from various manufacturers (both civil and military) to check and modify the correlation and to improvise if necessary.

Revision of mass (i. e., weight) data is a continuous process. In each project phase, the weight-estimation method is refined for better accuracy. During the con­ceptual design phase, semi-empirical relations based on statistical data are used; in subsequent phases, more detailed analytical and statistical methods are used. CAD solid models offer accurate geometric representations to improve volume pre­diction. Actual mass is known when components are manufactured, providing an opportunity to assess the mass-prediction methodology. The unavoidable tendency is that aircraft weight grows over time primarily due to modifications (e. g., reinforce­ments and additions of new components per user requirements). Although strength­testing of major aircraft components is a mandatory regulatory requirement before the first flight, structural-fatigue testing continues after many aircraft are already in operation. By the time results are known, it may not prove cost-effective to lighten an overdesigned structural member until a major retrofit upgrading is implemented at a later date.

The importance of the Six Sigma approach to make a design right the first time is significant to weights engineers. Many projects have suffered because of proto­types that were heavier than prediction or even experienced component failure in operation resulting in weight growth. The importance of weight prediction should not be underestimated due to not having an analytical approach involving high-level mathematical complexity, as in the case of aerodynamics. Correct weight estimation and its control are vital to aircraft design. One cannot fault stress engineers for their conservatism in ensuring structural integrity – lives depend on it. Weight-control engineers check for discrepancies throughout project development.

Mass prediction methodology starts with component weight estimation catego­rized into established groups, as described in Section 8.6. The methodology culmi­nates in overall aircraft weight and locating the CG and its range of variation that can occur in operation. Estimations of aircraft inertia are required to assess dynamic behavior in response to control input but then are not needed until completion of the conceptual design study – hence, inertia is not addressed in this book. Iteration of the aircraft configuration is required after the CG is located because it is unlikely to coincide with the position guesstimated from statistics in Chapters 6 and 7. A spreadsheet is recommended for calculations.