10.1.1 Coursework Content

The coursework task continues linearly with the examples worked out thus far. Readers must 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 behaviour.

Experience in the industry has shown that weight can only grow. Aircraft performance is extremely sensitive to weight because it must defy gravity. Aerodynamicists 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 by weights‐control engineers – a difficult task to perform. The aircraft operators also have loadmasters to ensure loading of aircraft for safe operation for the full operational envelope, on ground and in flight.

10.2.1 From ‘Concept Definition’ to ‘Concept Finalisation’

Progress of aircraft conceptual design study up to Chapter 9 led to presenting a preliminary configuration based on confident ‘guesstimates’ of maximum takeoff mass (MTOM), wing and other geometric parameters, and engine size from the statistics of past designs. From this chapter and onwards, work progresses towards refinement to accurate values as much as is possible, through iterations to update design to a better definition to arrive at the end to concept definition, in order to freeze aircraft to, that is, concept finalisation in Chapter 14. This completes Phase I of the project and enters into Phase II, the ‘Detailed Design Phase’; beyond the scope of this book.

In this chapter, aircraft and component weight estimation are carried out using reliable and well‐established semi‐empirical formulae, followed by establishing the aircraft CG position. It is expected that the estimated MTOM and CG position will be different from the guessed values. This will necessitate returning to a point to iterate through the subsequent stages to arrive at an improved concept definition. The iterative process may require repositioning of wing and undercarriage with respect to the fuselage. The last iteration will be after the aircraft is sized (Chapter 14) with a matched engine that could necessitate altering of wing reference area (like zooming in/out) yet maintain geometric parameters of sweep, taper ratio, dihedral and twist unaffected. Change of wing reference area will also affect empennage reference areas. Geometric changes will also change aircraft drag affecting engine size. All these changes will also change aircraft component weights. Therefore, at least, another iteration is required to fine‐tune aircraft configuration to, that is, concept finalisation, to freeze the design. Experience designers can minimise iterative process. This book avoids iterations to save from repetitive work.

Once again, the authors recommend that the readers prepare spreadsheets for repeated calculations because iterations will ensue after the CG is established and the aircraft is sized.

10.5.1 Neutral Point, NP

Aircraft NP is the aerodynamic centre of a weightless aircraft (see Section 3.7). NP is the point where dM/dα is constant, that is, aircraft pitching moment is invariant to attitude changes. In‐flight CG has to take into account of the influence wing position relative to fuselage and H‐tail size that determine the aircraft aerodynamic centre, a shape dependent aerodynamic parameter. Therefore, the position of NP shifts if external geometric shape changes, for example, between clean and dirty (hi‐lift devices and/or undercarriage are in extended position) conditions. Also, the position of CG changes depending on the fuel and payload/cargo loading status; seen as CG travel. The CG position with respect to the aircraft neutral is an important consideration to maintain aircraft stability.

Figure 10.4 shows that the aircraft NP moves backward near ground during field performance (takeoff and landing at dirty configuration) 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 pre‐flight aft‐most CG location is forward of the in‐flight NP centre by a convenient margin, which should be as low as possible to minimise trim. The distance between aircraft CG and the aircraft NP is known as the stability margin of the aircraft. (To repeat that the main‐wheel contact point (and strut line) is aft of the aft‐most CG, the subtending angle, β, should be greater than the fuselage‐rotation angle, γ, as described in Section 9.6.)

10.5.2 Aircraft CG Travel

Aircraft approaches to unstable flight as CG moves close to the NP. CG behind NP makes an aircraft unstable (see Chapter 18). The designers must ensure safe operation for the full operating envelope. (Note that an aircraft can be flown at neutral stability. This requires the pilot to struggle all the time to maintain steady attitude that will tire them out, which is not a safe way to fly. A bicycle ride is a good example where one has to learn to stay in balance.) Advanced military combat aircraft with Fly‐by‐Wire (FBW) technology can have relaxed static stability to provide quicker responses. That is, the margin between the aft‐most CG and the in‐flight NP is reduced (it may be even slightly negative), but design considerations relative to the undercarriage position remain unaffected. The following are some points for consideration that determine the extremities of CG travel.

Forward‐Most CG Position

1. If the aircraft is loaded to MTOW or close to it in a manner that makes CG stay forward then load on the nose becomes high and steering problems could show up. The nose gear limit line for high aircraft is shown in Figure 10.2.
2. Aircraft in a dirty configuration with a forward CG position have a high nose down moment requiring adequate elevator power with downward force to keep the aircraft in balance to maintain the desirable attitude. In the case of bad design, trim may run out at the landing configuration with a forward CG when the pilot may not be able to raise the nose up sufficiently to the landing attitude.
3. A larger sized H‐tail may overcome this problem at the expense of high stick force. It can now be realised that a compromise is required to size the H‐tail.

Aft‐Most CG Position

1. The aft‐most CG limit is concerned with aircraft safety. As the aircraft CG moves closer to the NP, which must be always behind, the pitch actuation force becomes lighter and the aircraft becomes more responsive with the pilot requiring little force to change pitch attitude. In loose terms, the control stick feels lighter. It suites military pilots, but there is a limit beyond which the pilot will have to struggle with constant attention. A limit has to be given for safe operation, which is the aft limit of the CG.
2. During field performance, aft‐most CG at a high load will have a high main gear load (Section 9.7). The main gear limit line for a highly loaded aircraft is shown in Figure 10.2.
3. For pitch stability consideration, a minimum stability margin has to be given suited to the H‐tail size. Trim drag is minimum at the aft‐most CG position.

CG at MTOM is not necessarily at the forward‐most limit. Similarly, the CG of an empty aircraft at OEW (Operator's Empty Weight) is not necessarily at the aft‐most position. The CG is always forward of the NP point. The loadmaster's responsibility is to make sure the aircraft CG travel stays within the specified limits for the full operational envelope.

Initially, locations of some of the components (e.g. the wing) were arbitrarily chosen based on designers' past experience, which works well (see Chapter 8). 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 case of supersonic Concorde).

10.6.1 Aircraft Components

10.6

1. Fuselage group (M_FU)
2. Wing group (M_W) – includes all structural items, for example, flaps, winglets and so on
3. Horizontal tail group (M_HT)
4. Vertical tail group (M_VT)
5. Nacelle group (M_N and M_PY – nacelle and pylon)
6. Undercarriage group (M_UC)
7. Miscellaneous (M_MISC) – dorsal, ventral fins and so on

Military aircraft have the following to be accounted for:

1. Fixed armament (M_FARM), for example, internal guns and so on
2. Pylon (M_PYL) – to carry armament load/drop tank.

This is the basic structure of the aircraft, for example, the fuselage shell (seats are separated under furnishing group) and so on.

10.7

1. Dry equipped engine (M_E)
2. Thrust reverser (M_TR)
3. Engine control system (M_EC)
4. Fuel system (M_FS)
5. Engine oil system (M_OI)

Military aircraft have the following to be accounted for:

1. Retarding devise (M_RD), for example, brake parachute (M_RD).

These come as a package and are items dedicated to power plant installation. They are mostly bought‐out items supplied by specialists.

10.8

1. Environmental control system (M_ECS)
2. Flight control system (M_FC)
3. Hydraulic/pneumatic system (M_HP) – sometimes lumped in with other systems
4. Electric system (M_ELEC)
5. Instrument system (M_INS)
6. Avionics system (M_AV)
7. Oxygen system (M_OX).

Military aircraft have the following to be accounted for:

1. Ejection seat system – this is no more treated as furnishing (M_EJ).

A wide variety of equipment, these are all vendor‐supplied bought‐out items.

10.9

1. Seat, galleys and other furnishing (M_SEAT)
2. Paints (M_PN).

Note that most of the weight goes to the fuselage, yet this is itemised under a different heading. Paints can be quite heavy. A well‐covered B737 with airline livery can use up 75 kg of paint.

Contingencies (M_CONT)

1. Contingencies (M_CONT) – a margin to allow unspecific weight growth

MEM – the total of these items.

This is the weight of the complete aircraft coming out from the production line for the first time to be airborne. The following items are added to MEM to obtain OEM.

1. Crew – Flight and cabin crews (M_CREW)
2. Consumables – Food, water and so on. (M_CON)

   (long duration military sortie have consumables for the crew)

OEM – aircraft is now ready for operation.

To make it sortie ready, add the payload and requisite fuel to obtain MRM.

1. Fuel (M_FUEL) – for the design range, which may not fill all tanks

Civil aircraft specific payload is the passengers and/or cargo is the following:

1. Payload (M_PL passengers @ 90 kg per passenger – includes baggage)

Military aircraft specific payload is the armament/electronic devices as the following:

1. Armaments (M_ARM) – missiles, bombs, firing rounds, drop tanks, electronic pods and so on.

MTOM (Maximum Takeoff Mass) = (MRM – taxi fuel)

That is the aircraft at the edge of runway ready to takeoff.

Civil Aircraft MTOM weight is the sum of weights of all component groups as totalled next.

where subscript ‘i’ stands for each component group listed here.

For civil aircraft:

10.10

For military aircraft:

10.11

10.7.1 Use of Semi‐Empirical Weight Relations versus Use of Weight Fractions

Section 1.11.2 discusses the merit of using the semi‐empirical weights formulae. The weight fraction data and related graphs are generated from the statistics of the latest aircraft in operation may not be close enough. Statistical based weight fractions data offer a good check to examine if the results obtained using the semi‐empirical weights formulae are in agreement or not. Industries maintain databases of existing aircraft and possible competition aircraft to give some idea of what is expected and serve to initiate the conceptual design study and make comparisons. After that, the typical trend in industry is to use their in‐house semi‐empirical weights formulae for aircraft components weight estimation with substantiated accuracy within 5% to begin with and gets refined to less than to 2% before the final assembly of the first aircraft is completed. Once manufactured, accurate weights data for components and the whole aircraft can be quickly weighed to obtain the exact value.

10.7.2 Limitations in Use of Semi‐Empirical Formulae

The first thing is to understand how semi‐empirical weight prediction equations are formulated and practised. Industrial practices are very different from academic studies in aircraft weight prediction. Industrial weights data are the real data by actually weighing the components and sub‐components. These real data are the backbones to formulate the weights equations. Using the real weights data, industrial formulated semi‐empirical relations and estimation procedures are very elaborate and not suitable in undergraduate classroom usage. They generate different sets of semi‐empirical formulae for each type of technology used to maintain high degree of accuracy. Their semi‐empirical equations are not mere regression analyses. The regression analyses are backed up by theories and factors to capture the dispersion. It is evolving all the time.

Industries have a separate dedicated group devoted to predicting aircraft weights. They interact with other groups, most importantly the structures group and production planning group, to feed aerodynamic group for aircraft performance analyses. They work full time from the inception to the end of the new aircraft design programme. They have many internal company documents on methodologies, some are over 300 pages. Industrial practices are proprietary information and kept commercial in confidence. Some older industrial methods and some technical reports are quoted in some publications but these are not easily available in public domain. It is difficult to find actual weights components weights data of currently manufactured in public domain. Reference [2] is a good source to obtain aircraft weights of older aircraft, mostly they are not in production but it still gives good values for academic usage.

In academies, various authors have formulated their own semi‐empirical relations with the data they generate and from some actual data. The work involves analytical derivations taking into account different technologies adopted in designing aircraft structures, suited to technologies adopted, manufacturing philosophies, aircraft performance goals and cost benefits. Due to these complexities, academic methods inherently do not carry high fidelity. In fact, Roskam [2] demonstrated three methods (i.e. Torenbeek, Cessna and US DATCOM) that yield different values, which is typical when using semi‐empirical relations. It is suggested that the readers work out using other available methods to examine any discrepancy. This is the real problem associated with different methods used in weight estimations and one has to live with it.

The SAWE was formed to address these issues in [6]. The SAWE [3] consolidated various equations and presented some generic equations with indices that can accept a wide range of user defined values. These served as baseline equations for many users. Also, [4] gives good relations in a simpler form that yield good values compared to many complex ones. The authors used these two sources and presented the semi‐empirical relations incorporating some changes in the expressions and adjusted the indices to match with the data in hand.

It is suggested that the instructors collect as many data, as much as possible, of aircraft component weights within the class of aircraft taken up for the study and adjust the available semi‐empirical relations for the technology differences, mainly in the choice of material. The state‐of‐the‐art in weight prediction has room for improvement. The advent of solid modelling (i.e. CAD) of components improved the accuracy of the mass‐prediction methodology.

Aircraft weight estimates will undergo several iterations. The final one will be after aircraft sizing and engine matching (Chapter 14).

The development of civil aircraft design has remained in the wake of military aircraft evolution. Competition within these constraints kept civil aircraft designs close to one another. As a result, variation in statistics is lower compared to military aircraft designs that are kept in secrecy. It is for this reason civil and military aircraft weight estimation methods are dealt with separately.

CIVIL AIRCRAFT

The following are some general comments with respect to civil aircraft mass estimation.

1. In the case of single engine propeller driven aircraft, the fuselage starts from aft of the engine bulkhead, as the engine nacelle is separately accounted for. These are mostly small aircraft. For wing‐mounted nacelles, this is not the case.
2. A fuselage‐mounted undercarriage is shorter and lighter for the same MTOM. Fixed undercarriage mass fraction is lower than the retractable type (typically 10% higher). The extent depends on retraction type.
3. Three‐engine aircraft are not dealt with here. Also fuselage‐mounted turboprop powered aircraft are not discussed here. Not many of these kinds of aircraft are manufactured. There is enough information given herein for the readers to adjust mass accordingly for the classes of current operating aircraft.

Turbofan aircraft with a higher speeds would have a longer range compared to turboprop aircraft and, therefore, would have a higher fuel fraction (typically, 2000 nm range will have around 0.26). However, with rigorous analyses using semi‐empirical prediction, better accuracy can be achieved that captures the influence of other parameters that remain obscure in mass fraction method.

10.8.1 Mass Fraction Analyses

Using Eqs. 10.1 and 10.2, the MTOM can be written as follows.

In terms of weight with shorter symbols (MTOM = W₀), this relation can be rearranged as follows.

10.12

The problem arises on how get accurate OEM and fuel mass fractions. The OEM fraction (W_OEW/W₀ – Figure 7.3) and fuel weight fraction (W_f/W₀ – Figure 7.4) are range dependent. There is a wide spread, primarily on account of different design considerations and it becomes difficult to choose the right value.

Equation (15.38) gives the cruise sector range, , which gives,

10.13

where W_{cruise ‐ fuel} = (W_ini − W_final)_cruise.

From which the fuel weight fraction (W_fuel/W₀) can be computed as follows.

Estimations of MTOM using mass fraction analyses have some inherent inadequacies due to not having the fuel consumption during takeoff, climb, descent and landing. These are estimated from statistical values. The fuel fractions for these segments varies to the extent that using a generic factor will be at the loss of fidelity to determine MTOM. The pitfalls of using the statistical values of the mass fractions are highlighted in Section 10.7.1. One has to depend on the tabulated statistical values given in Section 10.8. If such values are picked up from the tabulated values or any empirical relations, then the merits of each design superiority gets masked.

Since the statistics of mass fraction has a wide error band masking the design merits, it is better to avoid using Eq. 10.13 to get the preliminary MTOM of a new aircraft. Estimation of aircraft component mass requires an initial guesstimated MTOW to start with. Therefore, relying on the statistics of MTOM within the class of aircraft for the technology adopted, this book progresses with initial guess of MTOM, in finer resolution. Thereafter, using proven semi‐empirical equations that can be made to reflect the merits of the structural efficiency with suitable material selection, aircraft component masses are evaluated. After computing all the aircraft component masses, a more accurate MTOM is obtained to replace the guessed MTOM.

Sections 10.7.2 and 10.10 outline how to use semi‐empirical equations to avoid inherent complexities as there are different sets of semi‐empirical equations proposed by different authorities. Figure 7.2 shows in a better resolution the relation between payload versus MTOM separated for the range class. Payload (Eq. 10.12) and range (Eq. (15.38)) dictate the aircraft design. It is suggested that the readers prepare statistical graphs specific to the class of aircraft types that their project embraces.

However, mass fraction analyses using equations such as Eqs. 10.12 and 10.13 offer good insights to make trend analyses, optimisation, rapid sensitivity studies and so on, as the nature of such studies is equation‐based, allowing analytical explorations.

10.10.1 Fuselage Group (M_F) – Civil Aircraft

A fuselage is essentially a hollow shell to accommodate a payload/crew. The drivers for the fuselage group mass are its length, L(↑), diameter, D_ave (↑), shell area/volume (↑), maximum permissible aircraft velocity, V(↑), pressurisation (↑), aircraft load factor and n(↑) and mass increases with engine/undercarriage installation. Maximum permissible aircraft velocity is the dive speed in the V‐n diagram explained in Chapter 17. For a non‐circular fuselage it is the average diameter obtained by taking half of the sum of width and depth of the fuselage. For a rectangular cross‐section (invariably unpressurised), it is the average of the fuselage width and height. Length and diameter give the fuselage shell area – the larger the area the greater the weight. Higher velocity and limit load ‘n’ would require more material for structural integrity. Installation of engines and/or undercarriage on the fuselage would require additional reinforcement mass. Pressurisation of the cabin increase the fuselage shell hoop stress requiring reinforcement and a rear mounting cargo door is a large addition of mass. (Note again that non‐structural items in fuselage, e.g. the furnishing, systems etc., are separately computed.)

Given next are several sets of semi‐empirically derived equations by various authors for the transport category of aircraft (the readers may bear in mind the remarks made in the first paragraph that the results may yield different values). Nomenclatures are rewritten in line with what is in this book. The equations are for all‐metal (aluminium) aircraft.

By Niu in FPS:

10.14

where

• k₁ = 1.05 for a fuselage‐mounted undercarriage
• = 1.0 for a wing‐mounted undercarriage
• k₂ = 1.1 for a fuselage‐mounted engine
• = 1.0 for a wing‐mounted engine
• S_{net_fus_wetted_area} = fuselage shell gross area minus cut‐outs

Two of Roskam's suggestions are as follows [2]:

1. The General Dynamic method:

10.15

where

• K_inlet = 1.25 for inlets in or on the fuselage; otherwise, 1.0
• q_D = dive dynamic pressure in psf
• L = fuselage length
• D = fuselage depth

1. The Torenbeek method:

10.16

where

• K_f = 1.08 for a pressurised fuselage
• = 1.07 for the main undercarriage attached to the fuselage
• = 1.1 for a cargo aircraft with a rear door
• V_D = design dive speed in knots equivalent air speed (KEAS)
• L_{H tail} = tail arm of the H‐tail
• S_{net_fus_wetted_area} = fuselage shell gross area less the cut outs

By Jenkinson (taken from Howe) [8] in SI is given here,

10.17

The authors are not in a position to compare which one is better. As mentioned earlier, that depends on the type: weight equations do show some inconsistency. Torenbeek's equation has been in use for a long time. The Eq. 10.17 is the simplest one. Classroom usage may make the choice.

The authors modify Eq. 10.17 as shown in Eq. 10.18. The results appear to have yielded to satisfactory results. This allows capturing more details of the technology level. The authors suggest using Eq. 10.18 for coursework.

10.18

where c_fus is a generalised constant to fit the regression, as follows:

• c_fus = 0.038 for small unpressurised aircraft (leaving the engine bulkhead forward)
• = 0.041 for a small transport aircraft (≤19 passengers)
• = 0.04 for 20–100 passengers
• = 0.039 for a midsized aircraft
• = 0.0385 for a large aircraft
• = 0.04 for a double‐decked fuselage
• = 0.037 for an unpressurised, rectangular‐section fuselage

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

• k_e = for fuselage‐mounted engines = 1.05–1.07
• k_p = for pressurisation = 1.08 up to 40 000‐ft operational altitude
• = 1.09 above 40 000‐ft operational altitude
• k_uc = 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
• k_VD = 1.0 for low‐speed aircraft below Mach 0.3
• = 1.02 for aircraft speed 0.3 < Mach < 0.6
• = 1.03–1.05 for all other high‐subsonic aircraft
• k_door = 1.1 for a rear‐loading door

The value of index x depends on the aircraft size: 0 for aircraft with an ultimate load (nult) < 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 × nult)^x = 1. The value of index y is very sensitive. Typically, the last term index, 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 (n_ult < 5) Eq. 10.18 can be simplified to:

10.19

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

10.20

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:

10.21

In a simpler form, if there is some reduction in mass due to lighter material then reduce by that factor. Say there is 10% mass saving then M_Fcivil = 0.9 × M_{Fcivil_all metal}.

10.10.2 Wing Group – Civil Aircraft

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 in‐flight wing‐load relief; however, for small turbofan aircraft, they may not be possible due to the lack of ground clearances (unless engines are mounted over the wing (Hondajet) or it is a high‐wing aircraft (BAe146) – few are manufactured).

The drivers for the wing group mass are its planform reference area, S_W (↑); aspect ratio, AR(↑); quarter‐chord wing sweep, Λ_¼, is (↑); wing‐taper ratio, (↑); mean‐wing t/c ratio, (↓); maximum permissible aircraft velocity, V(↑); aircraft limit load, n(↑); fuel carried, (↓) and wing‐mounted engines, (↓). The AR and wing area give the wing span, b. Because the quarter‐chord wing sweep, Λ_¼ is 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 due to having better stiffness.

A well‐established general analytical wing weight equation published by SAWE [4] is as follows (others are not included):

10.22

where

• C = wing root chord
• B = width of box beam at wing root
• S_CS = wing‐mounted control surface reference area
• M_dg = MTOM

The authors make some modification to the equation for classroom usage. The term (M_dgN_Z)^x1 in the book nomenclature is (MTOM × n_ult)^0.48. The term (B/C)_t^x7 S_CS^x8 is replaced by the factor 1.005 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, Eq. 10.22 is replaced by Eq. 10.23 in SI (the MTOM is estimated):

10.23

where

• c_w = 0.0215 and flaps standard fitments to the wing
• k_uc = 1.02 for wing‐mounted undercarriage, otherwise 1.0
• k_sl = 1.04 for use of a slat
• k_sp = 1.01 for a spoiler
• k_wl = 1.01 for a winglet (a generalised approach is to have a standard size)
• k_re = 1 for no engine, 0.98 for two engines and 0.95 for four engines (generalised)

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:

10.24

In a simpler form, if there is some reduction in mass due to lighter material then reduce by that factor. Say there is 10% mass saving then M_Wcivil = 0.9 × M_{Wcivil_all metal}.

10.10.3 Empennage Group – Civil Aircraft

Empennages are also 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 10.23 is modified to suit empennage mass estimation. Both the horizontal and vertical tail plane mass estimations have a similar form but differ in the values of constants used.

10.25

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:

10.26

In a simpler form, if there is some reduction in mass due to a lighter material then reduce by that factor. Say there is 10% mass saving, then M_EMPcivil = 0.9 × M_{EMPcivil_all metal}.

Writing the modified equations in terms of the book's notation, Eq. 10.25 is changed to (empennage is now split into a H‐tail and V‐tail):

Horizontal Tail:

For all tail movement, use k_emp = 0.02, k_conf = 1.05, otherwise 1.0.

10.27

Vertical Tail:

For V‐tail configurations, use k_emp = 0.0215 (jet) to 0.025 (prop), k_conf = 1.5 for T‐tail, 1.2 for mid‐tail and 1.0 for low tail.

10.28

10.10.4 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 follows.

Jet engine type (includes pylon mass)

The preferred method is to use the factor of the manufacturer supplied dry engine weight for the bypass ratio (BPR) as follows.

10.29

where

• k_nac = 0.28 to 0.3 for BPR 4.0
• = 0.30 to 0.32 for BPR 6.0
• = 0.32 to 0.34 for BPR 8.0

If dry engine weight is not available, then use the factor for the T_SLS in kN per engine as follows.

10.30

where

• k_nac = 6.2 for a BPR 4.0
• = 6.4 for BPR 6
• = 6.6 for BPR 8

Thrust reverser increases the nacelle weight by about 40–50%.

Turboprop Engine Nacelle

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. the 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:

10.31

For a turboprop nacelle housing an undercarriage:

10.32

For a fuselage‐mounted turboprop nacelle with a pylon:

10.33

Piston Engine Nacelle

For fuselage‐mounted (single) engine, the nacelle is integral to the fuselage.

For wing‐mounted engine, the piston engine nacelle has as follows:

10.34

For a wing‐mounted, piston engine nacelle:

10.35

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:

10.36

In a simpler form, if there is some reduction in mass due to lighter material then reduce by that factor. Say there is a 10% mass saving, then M_Fcivil = 0.9 × M_{Fcivil_all metal}.

10.10.5 Undercarriage Group – Civil Aircraft

Chapter 9 describes undercarriages and their types in detail. Undercarriage size depends on an aircraft's MTOM. Mass estimation is based on a generalised 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:

10.37

For a mid‐wing‐mounted undercarriage:

10.38

For a high‐wing‐mounted undercarriage:

10.39

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.

10.40

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

10.10.6 Miscellaneous Group – Civil Aircraft

Carefully examine which structural parts are omitted (e.g. delta fin). Use the mass per unit area for a comparable structure (i.e. a lifting surface or a body of revolution; see Section 10.4). If any item does not fit into the standard groups listed up to Section 10.10.5, then it is included in this group. Typically, this is expressed as:

10.41

10.10.7 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 mass.

Turbofans

1. Equipped dry engine mass (M_{E_dry})
2. Thrust‐reverser mass (M_TR), if any – mostly installed on bigger engines
3. Engine control system mass (M_EC)
4. Fuel system mass (M_FS)
5. Engine oil system mass (M_OI)

Turboprops

1. Equipped dry engine mass (M_{E_dry}) – includes reduction in gear mass to drive the propeller
2. Propeller (M_PR)
3. Engine control system mass (M_EC)
4. Fuel system mass (M_FS)
5. Engine oil system mass (M_OI)

Piston Engines

1. Equipped dry engine mass (M_E) – includes reduction gear, if any
2. Propeller mass (M_PR)
3. Engine control system mass (M_EC)
4. Fuel system mass (M_FS)
5. Engine oil system mass (M_OI)

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. with fuel pump and generator). The engine thrust‐to‐weight ratio (T/M_{dry engine}; thrust is measured in Newton) is a measure of dry engine weight in terms of rated thrust (T_SLS). Typically, T/M_{dry engine} varies between 4 and 8 (special‐purpose engines can be more than 8). For turboprop engines, the mass is expressed as (SHP/M_{dry engine}); for piston engines, it is (HP/M_{dry 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:

For the variant engine mass, only add the increment/decrement of the dry engine mass to the baseline power plant group mass.

Turbofan

Civil aircraft power plant (with no thrust reverser):

10.42

Civil aircraft power plant (with thrust reverser):

10.43

Turboprop

Civil aircraft power plant:

10.44

where 1.5 ≤ k_tp ≤ 1.8. (due to propeller and gear box)

Piston Engine

Civil aircraft power plant:

10.45

where 1.4 ≤ k_p ≤ 1.5.

APU (if any)

10.46

10.10.8 Systems Group – Civil Aircraft

The systems group includes flight controls, hydraulics and pneumatics, electrical, instrumentation, avionics, and environmental controls (see Section 10.6.1).

10.47

10.48

10.49

10.10.9 Furnishing Group – Civil Aircraft

This group includes the seats, galleys, furnishings, toilets, oxygen system and paint (see Section 10.6.1). At the conceptual design stage, they are grouped together to obtain the furnishing group.

10.50

10.51

10.52

10.10.10 Contingency and Miscellaneous – Civil Aircraft

A good designer plans for contingencies; that is:

10.53

Miscellaneous items should also be provided for; that is:

10.54

10.10.11 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 10.4.

10.10.12 Payload – Civil Aircraft

A civil aircraft payload is basically the number of passengers at 100 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.

10.10.13 Fuel – Civil Aircraft

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. At this design stage, statistical data are the only means to estimate fuel load, which is then revised in Chapter 15.

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

10.12.1 Fuselage Group Mass

Consider a 5% weight reduction due to composite usage in non‐load‐bearing structures (e.g. floorboards).

Use Eq. 10.19:

For civil aircraft having n_ult < 0, then (MTOM × n_ult)^x = 1; therefore:

There is a 5% reduction of mass due to use of composites M_Fcivil = 0.95 × M_{Fcivil_all metal} ≈ 930 kg

Checking with Torenbeek's method (Eq. 10.16)

The higher value of the two (930 kg) is kept. This gives a safer approach to begin with.

10.12.2 Wing Group Mass

Consider that it has a 10% composite secondary structure, that is, k_mat = (0.9 + 0.9 × 0.1) = 0.99.

It has no slat making k_sl = 1.0 and is without a winglet, k_wl = 1. For the spoiler, k_sl = 1.01 and for a wing‐mounted undercarriage k_uc = 1.02. Equation 10.23 gives

where c_w = 0.0215 and flaps are standard fitment to the wing.

S_W = 30 m², n_ult = 4.125, b = 15 m, AR = 6.75, λ = 0.375, Fuel in wing, M_WR = 1140 kg, Λ = 14° and t/c = 0.10. Load factor, n = 3.8.

Equation 10.23 becomes:

10.12.3 Empennage Group Mass

Use Eq. 10.25.

Horizontal Tail:

A conventional all moving tail has k_emp = 0.2, k_conf = 1.0. Consider that it has a 20% composite secondary structure, that is, k_mat = (0.8 + 0.9 × 0.2) = 0.98.

MTOM = 9500 kg, n_ult = 3.8, S_HT = 5.5 m² (exposed), AR = 3.5, λ = 0.3, Λ = 16° and t/c = 0.10.

Use Eq. 10.27.

Vertical Tail (T‐tail):

k_conf = 1.5. Consider that it has a 20% composite secondary structure, that is, k_mat = (0.8 + 0.9 × 0.2) = 0.98.

MTOM = 9200 kg, n_ult = 3.8, S_VT = 3.5 m² (exposed), AR = 2.0, λ = 0.5, Λ = 20°, t/c = 0.1 and V_D = 380 kt = 703.76 km/h = 195.5 m s⁻¹.

Use Eq. 10.28.

10.12.4 Nacelle Group Mass

The Honeywell TFE731 at dry weight is 379 kg per engine.

Engine thrust = 17 230 N (3800 lb) per engine with a BPR 4.

Use Eq. 10.29: M_{nac + pylon} = 0.3 × 379 = 113.7 ≈ 114 kg/nacelle

Two nacelles = 228 kg.

10.12.5 Undercarriage Group Mass

MTOM = 9500 kg low‐wing mount; M_U/C wing = 0.04 × 9500 = 380 kg.

10.12.6 Miscellaneous Group Mass

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

10.12.7 Power Plant Group Mass

This is determined from statistics until it is sized in Chapter 14. A typical engine is of the class Allison TFE731–20 turbofan with thrust per engine = 15 570 N to 17 230 N (3500–3800 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. M_DRYENG = 379 kg per engine. This gives:

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

M_ENG = 1.4 × M_DRYENG per engine = 1.4 × 379 = 530.5 kg ≈ 530 kg. For two engines, M_ENG = 1060 kg.

10.12.8 Systems Group Mass

MTOW = 9500 kg (use Eq. 10.47), M_SYS = 0.11 × 9500 = 1045 kg.

10.12.9 Furnishing Group Mass

MTOW = 9500 kg (use Eq. 10.50), M_FUR = 0.065 × 9500 = 618 kg.

10.12.10 Contingency and Miscellaneous Group Mass

MTOW = 9500 kg (use Eq. (10.53 and 10.54)), M_CONT = 0.016 × 9500 = 152 kg.

10.12.11 Crew Mass

There are two flight crew members and no cabin crew. Therefore, M_CREW = 2 × 90 = 180 kg.

10.12.12 Payload Mass

There are 10 passengers. Therefore, M_PL = 10 × 100 + 100 = 1100 kg.

10.12.13 Fuel Mass

The range requirement is 2000 nm carrying 10 passengers. From statistical data given in Table 10.1, take the highest value of 26% of MTOW, M_FUEL = 0.26 × 9500 = 2470 ≈ 2500 kg.

10.12.14 Weight Summary

Table 10.7, later, is the weight summary obtained by using the coursework example worked out thus far. The last column provides the estimation as a result of the graphical solution (in bracket – kg).

This computation requires further iterations to be fine‐tuned for better accuracy. The CG position is established in the next section, where further iterations will yield a better picture.

Variant Aircraft in the Family

For simplification, linear variations are considered, which should be worked out as a coursework exercise.

Long Variant. Increase payload = 400 kg, fuselage = 300 kg, furnish = 200 kg, fuel = 300 kg and others = 200 kg. Total increase = +1400 kg; MTOM_Long = 10 900 kg (no structural changes).

Small Variant. Decrease payload = 500 kg, fuselage = 350 kg, furnish = 250 kg, fuel = 350 kg and others = 250 kg (lightening of the structures). Total decrease = −1700 kg; MTOM_Small = 7800 kg.

10.12.15 Bizjet Aircraft Mass and CG Location Example

The x‐axis is the fuselage axis and ZRP is the y‐axis. Table 10.6 and Eqs. 10.55 and 10.56 are used to locate the CG. SI units are used. Determination of aircraft NP is not done in this book but taken at around 55% of wing MAC.

Item group	Mass (kg)	X (m)	Moment	Z (m)	Moment
Fuselage @ 40%	930	6.08	5655	1.4	1 302
Wing @ 30% of MAC	854	7.4	6320	1	854
H‐tail	118	14	1652	8	944
V‐tail	83	15	1245	3	249
Undercarriage (nose)	110	1.8	198	0.4	44
Undercarriage (main)	270	8	1840	0.5	135
Nacelle/pylon	228	10.5	2415	2	460
Miscellaneous
Power plant	1060	10.7	11 342	2	2120
Systems	1045	6.5	6793	1	1045
Furnishing	618	6	3708	2	1236
Contingencies	152	3	456	1.2	182
MEM	5470
Crew	180	3	540	1.4	252
Consumable	120	4	480	1.5	180
OEM	5770
Payload	1100a	6.2 (6.4)	7040	1.1	1210
Fuel	2500	7.8	19 500	1	2500
Total moment at MTOM of 9370 ≈ 9400			69 184		12 713
		7.36 m		1.352 m
MRM	9450

Bold values highlight the most important parameters.

^aNo frills 10 passengers or a luxury arrangement for eight passengers.

The CG should lie in the plane of symmetry (there are unsymmetrical aircraft).

The CG locations for the CG_{Long Variant} and the CG_{Short Variant} may be worked out by the readers.

10.12.16 First Iteration to Fine‐Tune CG Position Relative to an Aircraft and Components

The preliminary aircraft configuration begins in Chapter 8 with a guesstimated MTOM, engine size and CG position. It is unlikely that the computed aircraft mass as worked out in this chapter will match the estimated mass. In fact, the example shows that it is lighter, with a more accurate CG position; therefore, this replaces the estimated values in Chapter 8. The differences from guesstimate and revised in chapter are as follows.

Also note that both mass and CG location are slightly different from preliminary data.

	Estimated at concept definition	Revised in Chapter 9 (Figure 9.22)
	(Chapter 8)	First iteration in Chapter 10
MTOM – kg	9500	9400
MAC leading edge location – m from ZRP	6.77	6.76
Aft CG limit, from ZRP	7.4 m	7.36 m
CG position, from ground – m	1.38 m	1.352 m

The differences are small. From now onwards, the revised values will used; most importantly the MTOM is now revised to 9400 kg. In principle, the aircraft configuration must be revised at this stage of progress as the first iteration. The worked‐out example of the Bizjet shows small differences, hence no iteration is carried out as a simple repositioning of battery, black‐boxes and so on will do the job. In that case the angle, β, is kept satisfied. Final sizing is accomplished in Chapter 14, when iteration may be required.

MILITARY AIRCRAFT

10.15.1 Military Aircraft Fuselage Group (SI System)

Military aircraft do not have a constant section fuselage and are more densely packed. A combat aircraft fuselage is not pressurised (only cockpit is). In a blended body it is not easy to lineate fuselage line from the wing. The best would be construction specific where the joining line of the wing is the line for fuselage. It is much simpler for the example of the Advanced Jet Trainer (AJT) in hand. In this case take D_ave = average of the shape (see Figure 8.14) around the engine. The expression includes fuselage‐mounted side/chin intakes. (The AJT does not have a pod – intake weight is taken as integral with fuselage weight.)

where c_fus = 0.175

• k_uc = 1.05 for a fuselage‐mounted undercarriage with bulge and 1.03 without bulge.
• k_mat = 1.0 for metal, otherwise make a percentage weight reduction.
• k_para = 1.002 if there is a brake parachute, otherwise 1.
• k_intake = 1.005

10.58

10.15.2 Military Aircraft Wing Mass (SI System)

The equation is simplified by the authors for classroom usage. The same remarks as made for civil aircraft are also applicable to the military case, except that the values of factors change. Military wings are thinner. In many cases, the wing does not carry fuel (e.g. F104) and does not have winglets.

Writing the modified equation in terms of the book's notation, Eq. 10.21 is changed to:

10.59

• c_w = 0.0215 for AR < 5 (high subsonic – jet)
• = 0.023 for AR > 5 (low subsonic – propeller)
• k_mat = effect of material change as in the case of the fuselage
• k_uc = 1.002 for a wing‐mounted undercarriage, otherwise 1.0
• k_sl = 1.004 for use of a slat and k_sp = 1.005 for a spoiler. Flaps are standard.

10.15.3 Aircraft Empennage

The equation is simplified by the authors for classroom usage. The same remarks as made for civil aircraft are also applicable to the military case, except that the values of factors change as follows. Equation 10.25 for civil aircraft gives:

k_mat = Effect of material change as in the case of fuselage.

The modified equations, in terms of the notations in Eq. 10.25, are as follows.

Horizontal Tail:

For all tail movement, use k_emp = 0.018, k_conf = 1.05, otherwise 1.0.

10.60

Vertical Tail:

For tail configurations, use k_emp = 0.022, k_conf = 1.1 for T‐tail, 1.05 for mid‐tail and 1.0 for low tail.

10.61

10.15.4 Nacelle Mass Example – Military Aircraft

A typical combat aircraft does not have a nacelle and a pod – intake weight is taken to be integral with fuselage weight.

10.15.5 Power Plant Group Mass Example – Military Aircraft

From the engine T_SLS the engine dry mass can be obtained from its thrust‐to‐weight ratio, this typically varies from 5 to 8. It is always better to use engine manufacturer's supplied data that is freely available in the public domain. Otherwise, refer to Table 10.8 to obtain engine dry mass.

For the variant engine mass, only add the increment/decrement of the dry engine mass to the baseline power plant group mass.

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

Jet engine

10.62

Turboprop

Civil aircraft power plant:

10.63

where 1.5 ≤ k_tp ≤ 1.8. (due to propeller and intake)

APU (if any)

10.64

10.15.6 Undercarriage Mass Example – Military Aircraft

Undercarriage size depends on an aircraft's MTOM. A military aircraft undercarriage is heavier in order to withstand hard landings.

Tricycle type (retractable) – wing‐mounted (nose and main gear estimated together)

For a low‐wing‐mounted undercarriage:

10.65

For a mid‐wing‐mounted undercarriage:

10.66

For a high‐wing‐mounted have fuselage‐mounted :

10.67

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

10.15.7 System Mass – Military Aircraft (Higher Mass Fraction for Lighter Aircraft)

System mass comprises of avionics, electrical, plumbing, ECS (oxygen system) and so on.

10.68

10.15.8 Aircraft Furnishing (Ejection Seat) – Military Aircraft

10.69

As such, military aircraft do not have furnishings but the following can be added:

10.15.9 Miscellaneous Group (M_MISC) – Military Aircraft

Carefully examine what structural parts are left out, for example, delta fin and so on. If any item does not fit into the standard groups as listed here – include it in this group. Typically, this be expressed as

10.70

10.15.10 Contingency (M_CONT) – Military Aircraft

A good designer keeps provision for contingency, that is,

10.71

10.15.11 Crew Mass

There are trainer and trainee pilots. Take a mass of 90 kg per crew member.

10.15.12 Fuel (M_FUEL)

Fuel load is mission specific. This can be determined by proper performance estimation shown in Chapters 11 and 13. At this stage, statistical data is the only means to guesstimate fuel load that will be revised in Chapters 11 and 13.

Military aircraft is mission specific and also has to depend on statistical data at this stage of design until accurate fuel load can be estimated through performance estimation.

10.15.13 Payload (M_PL)

Military aircraft payload is the armament specified by the user requirements. Internal guns are taken as systems weight.

10.17.1 AJT Fuselage (Based on CAS Variant)

Consider that it has a 10% weight saving on account of composite usage in secondary structures (10%), that is, k_mat = 0.99. MTOM = 6500 kg. L = 12.1 m, D_ave = 0.5 × (1.8 + 0.9) = 1.35 m and V_D = 620 kt = 1150 km/h = 319.4 m s⁻¹, n_ult = 6 g. Equation gives:

10.17.2 AJT Wing (Based on CAS Variant)

Consider that it has a 10% weight saving on account of composite usage in secondary structures (10%), that is, k_mat = 0.99. MTOM = 6500 kg. n_ult = 6 g, S_W = 17 m², AR = 5.3, λ = 0.3, M_WR = 828 kg, Λ = 20°, t/c = 0.10 and = 1150 km/h = 319.4 m s⁻¹. It has no slat making k_sl = 1 and has 828 kg fuel in the wing. For the spoiler, k_sl = 1.005 and the undercarriage is not wing‐mounted k_uc = 1. Equation gives:

10.17.3 AJT Empennage (Based on CAS Variant)

Horizontal Tail: Equation gives:

Consider that it has a 10% weight saving on account of composite usage in secondary structures (10%), that is, k_mat = 0.99. MTOM = 6500 kg.

AJT horizontal tail is all moving, that is, configuration factor k_conf = 1.05. Note that the exposed area constitutes as the empennage mass as constructed.

Vertical Tail: Equation gives:

10.17.4 AJT Nacelle Mass (Based on CAS Variant)

The AJT does not have a pod – the intake weight is taken to be integral to the fuselage weight.

10.17.5 AJT Power Plant Group Mass (Based on AJT Variant)

Suitable engines in the class are the following. Both have BPR = 0.75 and engine = 22 in.

For AJT – Adour 861 producing 26.66KN (6000 lb), dry weight = 590 kg (1300 lb).

For CAS – Adour 871 producing 31.4KN (7060 lb), dry weight = 610 kg (1330 lb).

Taking Adour 871 for AJT (weights will not change with final engine sizing in Chapter 14).

Total power plant group mass can be expressed semi‐empirically as (Eq. gives):

10.17.6 AJT Undercarriage Mass (Based on CAS Variant)

To maintain commonality, use the heaviest fully loaded CAS mass of 7500 kg. This is the only time CAS maximum mass has been applied to get component mass.

For fuselage‐mounted undercarriage use Eq. 10.68.

10.17.7 AJT Systems Group Mass (Based on AJT Variant)

Equation 10.65 gives:

10.17.8 AJT Furnishing Group Mass

Ejection seat plus the oxygen system are taken in this group to be 90 kg per crew member.

10.17.9 AJT Contingency Group Mass

Equation 10.71 gives:

10.17.10 AJT Crew Mass

There are trainer and trainee pilots. Therefore, M_CREW = 2 × 90 = 180 kg

10.17.11 AJT Fuel Mass (M_FUEL)

Fuel load is mission specific. From AJT statistics in Section 6.12.1, M_{FUEL_NTC} = M_NFC − OEW = 4800–3700 = 1100 kg (tank not full).

It is assumed that practise range is less than 100 miles away and an AJT reaches the range at economic cruise to make two to three passes for armament training at a sortie duration of about 50 minutes. The tankage capacity can hold 1400 kg fuel for ferry flights and, if required, an AJT armament practice sortie can be made longer at a slightly higher MTOM.

10.17.12 AJT Payload Mass (M_PL)

Military aircraft payload is its armament, which is specified by the user requirements. Internal guns are taken as systems weight. Practice armament (includes gun ammunition) = 1800 kg.

10.19.1 TPT Fuselage Example

Consider that it has a 10% weight saving on account of composite usage in secondary structures (10%), that is, k_mat = 0.99. MTOM = 2800 kg. L = 9.5 m, D_ave = 0.5 × (1.8 + 0.9) = 1.35 m and V_D = 300 kt = 556.5 mph = 155 m s⁻¹, n_ult = 6 g. Equation 10.58 gives:

10.19.2 TPT Wing Example

Consider that it has a 10% weight saving on account of composite usage in secondary structures (10%), that is, k_mat = 0.99. MTOM = 2800 kg. n_ult = 6 g, S_W = 16.5 m², AR = 6.5, λ = 0.5, Λ = 0°, t/c = 0.15 and = 280 kt = 144 m s⁻¹. It has no slat making k_sl = 1.0, k_uc = 1.002 and has 470 kg fuel in the wing. For the spoiler, k_sl = 1.005 and the undercarriage is not wing‐mounted k_uc = 1. Equation 10.59 gives:

10.19.3 TPT Empennage Example

Horizontal Tail:

Equation 10.60 gives:

Consider that it has a 10% weight saving on account of composite usage in secondary structures (10%), that is, k_mat = 0.99. MTOM = 6600 kg.

AJT horizontal tail is all‐moving, that is, configuration factor k_conf = 1.05. Note that the exposed area constitutes as the empennage mass as constructed.

Vertical Tail: Equation 10.61 gives:

10.19.4 TPT Nacelle Mass Example

The TPT does not have a pod‐intake weight taken to be integral with fuselage weight.

10.19.5 TPT Power Plant Group Mass Example

Use Eq. 10.62:

Engine selected is Garratt TPT‐331‐12B (include propeller) derated to 1000 SHP.

For TPT – Garratt TPT‐331‐12B, dry weight = 175 kg (385 lb).

For COIN – Garratt TPT‐331‐12B at the rated power of 1100 SHP.

10.19.6 TPT Undercarriage Mass Example (Based on CAS Variant)

To maintain commonality, use the heaviest fully loaded COIN mass of 3100 kg. This is the only time COIN maximum mass has been applied to get component mass.

For a fuselage‐mounted undercarriage use Eq. 10.65.

10.19.7 TPT Systems Group Mass Example

Use Eq. 10.68

MTOW = 2800 kg, M_SYS = 0.12 × 2800 = 336 kg.

10.19.8 TPT Furnishing Group Mass Example

Ejection seat plus oxygen system are taken in this group to be 80 kg per crew member.

10.19.9 TPT Contingency Group Mass Example

Use Eq. 10.71:

M_CONT = 0 for TPT.

10.19.10 TPT Crew Mass Example

There are trainer and trainee pilots. Therefore, M_CREW = 2 × 90 = 180 kg.

10.19.11 TPT Fuel (M_FUEL)

Fuel load is mission specific. The tankage capacity can hold 1000 kg fuel.

10.19.12 TPT Payload (M_PL)

Military aircraft payload is its armament, which is specified by the user requirements. Internal guns are taken as systems weight. Practice armament (includes gun ammunition) = 1800 kg.

10.20.1 COIN Variant

The role of COIN aircraft is to take airborne military action to encounter insurgent (hence the name) movements against the government, for example, terrorism, paramilitary activities, civil war, guerrilla actions, crime syndicate and so on, within the country or at the borders. These are not exactly in CAS role but carried out at low intensity engagements where heavy armament is not required nor the radius of action is far away from conveniently located airbases. The aircraft should be capable of finding, tracking, and attacking targets (observation/attack – US nomenclature OA‐X role). The measures are primarily strafing and smaller missiles firing to take vehicles out of action. This type of aircraft need not have very speed but must be capable of high manoeuvre with tight turning capability. A typical high performance COIN armament load is around ±2000 lb. COIN variant of TPT is in many air force inventories, but a COIN variant carrying higher load than 1000 kg would penalise the baseline TPT variant to grow too big for ab initio/intermediate level training to gain proficiency in airmanship, as well as be expensive to operate for the intended role at that level. A true COIN aircraft should be a new design fit for the specific role, possibly be able to keep some components commonality to a considerably lesser extent with the TPT fleet in the air force inventory.

A COIN variant of the TPT example is suggested at the end of Section 8.17.4, as shown in Figure 8.25. Here, the suggestion extends to assist, if the readers would like to refine to a sized COIN variant by taking out the second pilot, ejection seat and flight‐deck equipment to the extent of reducing OEM by 250 lb to bring down to a bare‐bone OEM of 1650 lb. Add Δ50 lb for COIN avionics, fuel load to 550 lb and armament load of 750 lb to COIN MTOM ≈ 3000 lb, the same as the TPT MTOM. The engine is the same Garratt TPT‐331‐12B at the rated power of 1100SHP (the TPT engine is derated to 950 SHP). Tail‐arms are kept about the same so that the empennage areas retain commonality. Shortening of fuselage length will also shorten the wheel base requiring to shortening the wheel track as well. This can easily be done if the baseline TPT wing structure keeps provision to make minor modifications to move the attachment support members and the wheel well slightly inboard. This is merely a suggestion and is not worked out. The readers may give this a try.