Aircraft Design Projects - part 5 doc

“chap05” — 2003/3/10 — page 128 — #28 128 Aircraft Design Projects Climb rate (m/s) sea level 25 000 ft 36 000 ft 30 40 50 60 0 –20 –10 10 20 Aircraft speed (m/s) 0 50 100 150 200 250 300 Fig. 5.19 Aircraft rate of climb 0 5000 10 000 15 000 20 000 25 000 30 000 35 000 40 000 Aircraft altitude (ft) Max. climb rate (m/s) 20 30 40 50 60 0 10 Fig. 5.20 Aircraft climb and ceiling evaluation • The maximum speed even at 85 per cent thrust is 505 kt. This easily exceeds the specified requirement of 450 kt. • All the turn performance criteria are easily met. • Take-off ground run at 1856 ft is below the specified 2000 ft but with a derated engine of 85 per cent thrust, this increases to 2184 ft. • The approach speed requirement of 100 kts cannot be met except by a lighter aircraft (no weapon load). In this case an approach speed of 95 kt is achieved. • Landing ground run at 2215 ft also exceeds the specification of 2000 ft. Only with aircraft at lighter landing mass can the specification be met. • The ferry mission of 1000 nm cannot be met with internal fuel but can be achieved if 833 kg of fuel is carried externally. The maximum range that could be flown is estimated at 1464 nm. This is substantially less than the 2000 nm specified. It is “chap05” — 2003/3/10 — page 129 — #29 Project study: military training system 129 suggested that this requirement be reviewed as in its present form it would seriously compromise the overall aircraft design. • Climb and ceiling requirements are easily achieved. 5.8 Constraint analysis From the project brief there are six separate constraints to be considered in this analysis: 1. Take-off distance less than 2000 ft. 2. Approach speed no greater than 100 kt. 3. Landing distance less than 2000 ft. 4. Combat turn, at least 4g at sea level. 5. Combat turn, at least 2g at 25 000 ft. 6. Climb rate to provide for 7 min climb to 25 000 ft. 5.8.1 Take-off distance The equations to be used to determine the effect of the take-off criterion can be found in most textbooks (e.g. reference 4) as shown below: (T /W ) = (constant) (W /S)/(S take-off .C Ltake-off ) Obviously this represents a straight line on the (T /W ) versus (W /S) graph. For our aircraft the lift coefficient in the take-off configuration (C Ltake-off ) is assumed to be 1.7. The value S take-off represents the total take-off distance (i.e. ground roll plus climb distance to 50 ft). Assuming a climb gradient from zero to 50 ft of 5 ◦ gives a ground distance covered of 571 ft. Adding this to the specified ground roll of 2000 ft gives S take-off = 2571 ft (784 m). The constant in the above equation is assessed from Nicholi’s book 4 as 1.27 (in SI units with wing loading in kg/m 2 ), so (T /W ) = 1.27/(784 × 1.7)(W /S) = 0.00095(W /S) 5.8.2 Approach speed Assuming the approach speed V A = 1.2V STALL then: (W /S) landing = β(W/S) = 0.5 × ρ(V A /1.2) 2 × C Llanding /g V A is specified at 100 kts (52 m/s). β is the ratio of landing mass to take-off mass. At a maximum landing weight β = 0.9. At minimum landing weight (i.e. empty aircraft plus pilot plus 10 per cent fuel = 3311 + 136 + 90 = 3537 kg) β = 0.62. Assuming the lift coefficient in the landing configuration (C Llanding ) = 2.1 (W /S) = (0.5 × 1.225 × 52 × 52 × 2.1)/(1.2 × 1.2 × 9.81) = 273.6 @ β = 0.9 = 397.1 @ β = 0.62 Note: these constraints are constant (vertical) lines on the (T /W ) versus (W /S) graphs. “chap05” — 2003/3/10 — page 130 — #30 130 Aircraft Design Projects 5.8.3 Landing distance The approximate equation to determine ground run in landing can be rewritten as shown below: (W /S) = (S landing run × C Llanding )/(constant × β) The landing ground run S landing run is specified as 2000 ft (610 m). The lift coefficient in the landing configuration (C Llanding ) is assumed to be 2.1 (as above). The expression will be evaluated for the two landing mass fractions used above (i.e. β = 0.9 and 0.62). The (constant) in the expression above (in SI units with W /S in kg/m 2 )is5.0. (W /S) = (610 × 2.1)/(5.0 × β) = 284.7 @ β = 0.9 and 413.2 @ β = 0.62 Note: these are also constant vertical lines on the constraint diagram. 5.8.4 Fundamental flight analysis The fundamental equation used in the flight cases can be found in most textbooks. In terms of sea level, take-off thrust loading the equation is: (T /W ) TO = (β/α)[(q/β){C DO /(W /S) TO + k 1 (nβ/q) 2 (W /S) TO } + (1/V )(dh/dt) + (1/g)(dV /dt)] where (T /W ) TO is the take-off thrust loading α 1 = T/T SLS T SLS = sea level static thrust (all engines) β = W /W TO C DO and k 1 are coefficients in the aircraft drag equation, see below D = qS(C DO + k 1 C 2 L ) (W /S) TO is the take-off wing loading (N/m 2 ) n is the normal acceleration factor = L/W g = gravitational acceleration V is the aircraft forward speed q is the dynamic pressure = 0.5ρV 2 (dh/dt) = rate of climb (dV /dt) = longitudinal acceleration 5.8.5 Combat turns at SL In this flight condition the aircraft is in ‘sustained’ flight with no change in height and no increase in speed therefore the last two terms in the fundamental equation are both zero. At sea level α = 1 Assume that the turn requirement is appropriate to the mean combat mass (i.e. air- craft empty + pilot + half fuel + half weapon load = 3311 + 136 + 450 + 680 = 4577 kg/10 092 lb) Hence β = 4577/5707 = 0.8. “chap05” — 2003/3/10 — page 131 — #31 Project study: military training system 131 From previous analysis (in SI units) the best speed for turning at SL is about 150 m/s. ∴ q = 0.5 × 1.225 × 150 2 = 13 781 From the drag analysis done earlier (at 4577 kg with an increase in drag coefficient to represent the stores on the wing) at a speed of 150 m/s, C D = 0.03 + 0.017C 2 L . As specified, the aircraft is subjected to a normal acceleration n = 4 in the turn. T /W = 13 781{(0.03/(W /S) + 0.017 ×[4/13 781] 2 × (W /S)} 5.8.6 Combat turn at 25 000 ft This is similar to the analysis above but with α = 0.557/1.225 = 0.455. At 25 000 ft the best speed for excess power is 200 m/s (in SI units) ∴ q = 0.5 × 0.557 × 200 2 = 11 140 With β and C D values the same but with load factor n = 2 gives: T /W = (0.8/0.445)[(11 140/0.8){(0.03/(W /S)+0.017×[(2×0.8)/11 140] 2 ×(W /S)} 5.8.7 Climb rate This criterion assumes a non-accelerating climb, so the last term in the fundamental equation is zero but the penultimate term assumes the value relating to the specified rate of climb. We will use an average value of climb rate of 18.15 m/s (i.e. 25 000 ft in 7 min) and make the calculation at the average altitude of 12 500 ft, at a best aircraft speed of 150 m/s. At 12 500 ft α = 0.841/1.225 = 0.686 At 150 m/s q = 0.5 × 0.841 × 150 2 = 9461 Using the standard values for β at mean combat mass, and the drag coefficients (C DO and K) previously specified, we get: T /W = (0.8/0.686)[(9461/0.8){(0.03/(W /S) + 0.017 ×[(1 × 0.8)/9461] 2 × (W /S)} + 18.15(1/150) 5.8.8 Constraint diagram The above equations have been evaluated for a range of wing loading values (150 to 550 kg/m 2 ). The resulting curves are shown in Figure 5.21. The constraint diagram shows that the landing constraints (approach speed and ground run) present severe limits on wing loading. To identify the validity of the constraints relative to other aircraft, values appropriate to specimen (competitor) aircraft that were identified earlier in the study have been plotted on the same constraint diagram Figure 5.21. Some interesting conclusions can be drawn from this diagram: • The S212, T45, MiG, L159 and, to a lesser extent, the Hawk aircraft appear to fit closely to the climb constraint line. This validates this requirement. “chap05” — 2003/3/10 — page 132 — #32 132 Aircraft Design Projects Thrust /weight ratio Take-off run Climb rate FL250 turn SL turn Initial design point New design point Landing run and approach speed 62% MTOM Landing run and approach speed 90% MTOM 0.000 0.100 0.200 0.300 0.400 0.500 0.600 0.700 100 200 300 400 500 60 0 0 Wing loading (kg/m 2 ) Fig. 5.21 Aircraft constraint diagram • None of the existing aircraft satisfy the landing conditions at M LAND = 0.9M TO . This suggests that this requirement is too tight. • The turn requirements do not present critical design conditions for any of the aircraft. The 25 000 ft turn criteria is seen to be the most severe. Some further detailed analysis suggests that the aircraft is capable of a 3g turn rate at this altitude. Warning: The constraint analysis described above is a very approximate analytical tool as it does not take into account some of the finer detail of the design (e.g. detailed changes in engine performance with speed). It can only be used in the form presented in the initial design phase. Later in the development of the layout more detailed analysis of the performance will enable the effect of the various constraints on the aircraft design to be better appreciated. However, with this consideration in mind it is possible to use the constraint diagram to direct changes to the original baseline layout as discussed below. 5.9 Revised baseline layout The main conclusion from the constraint analysis and aircraft performance estimations is that the aircraft landing requirements are too tight and should be renegotiated with the customers. To provide evidence on the effects of the landing constraints, the revised baseline layout will ignore them. The new design can be analysed to show what landing characteristics are feasible. With the above strategy in mind the design point for the aircraft will be moved closer to the intersection of the take-off and climb constraint lines, i.e.: (T /W ) = 0.38 and (W /S) = 390 kg/m 2 (80 lb/sq. ft) Anticipating the need to increase aircraft mass to allow more fuel to be carried, the max- imum take-off mass is increased to 5850 kg (and the structural design mass increased “chap05” — 2003/3/10 — page 133 — #33 Project study: military training system 133 to 6100 kg). Using the new values for (T /W ) and (W /S) the new thrust and wing area become: T = 0.38 × 5850 = 4900 lb (SSL) S = 5850/400 = 14.65 m 2 (136 sq. ft) For an aspect ratio (AR) of 5, the new area gives a wing span (b) = 8.56 m and a mean chord = 1.71 m. For an aspect ratio of 4.5 the wing geometry becomes b = 8.12 m and mean chord = 1.80 m. Rounding these figures for convenience of the layout drawing gives: c mean = 1.75 m (5.75 ft) and b = 8.5 m (28 ft) ∴ gives, AR = 4.86 and S = 14.87 sq. m/160 sq. ft This geometry will be used in the new layout. Also, since the tip chord on the previous layout seemed small, the taper ratio will be increased to 0.33. Hence C mean = (C tip + C root )/2 = 1.75 m (assumed) With, (C tip /C root ) = 0.33 This gives C root = 2.63 m/8.6 ft, C tip = 0.87 m/2.8 ft 5.9.1 Wing fuel volume It is now possible to check on the internal fuel volume of the new wing geometry. Assume 15 per cent chord is occupied by trailing edge devices and 33 per cent span is taken by ailerons (assume no fuel in the wing tips ahead of the ailerons). Although previously the wing thickness was assumed to be 10 per cent, it has now become clear that the aircraft will require substantial internal volume for fuel storage. To anticipate this, the wing thickness will be increased to 15 per cent in the expectation that supercritical wing profiles can be designed to assist in the transonic flow conditions particularly for the high-speed development aircraft. With the above geometry (see Figure 5.22) and assuming 66 per cent of the enclosed volume is available for fuel, gives an internal wing fuel capacity of 0.5 m 3 . A total fuel load of 1050 kg equates to a volume of 305 Imp. gal. This requires a volume of 1.385 m 3 . It is therefore necessary to house some fuel in the aircraft fuselage (namely 1.385 − 0.5 = 0.885 m 3 ). This is not uncommon on this type of aircraft. The preferred place to keep the fuel is in the space behind the cockpit and between the engine air intakes. This is close to the aircraft centre of gravity, therefore fuel use will not cause a large centre of gravity movement. For our layout it would be preferable to keep the fuel tank below the wing structural platform to make the wing/fuselage joint simpler. From the original aircraft layout this fuselage space would provide a tank volume of about 1 × 2 × 0.5 = 1m 3 . This is satisfactory to meet the internal fuel requirement. Using all of this space for fuel may present a problem for the installation of aircraft systems. To anticipate the need for extra space in the fuselage to house the electronic and communication systems an extra 0.5 m will be added to the length of the fuselage. Moving the engine and intakes back to rebalance the aircraft will also provide a cleaner installation of the intake/wing junction (i.e. moving the intake behind the wing leading edge). “chap05” — 2003/3/10 — page 134 — #34 134 Aircraft Design Projects Wing LE extension Wing LE fuel tank 25% MAC Front spar line MAC 25%C 50%C Rear spar line FL A P AILERON Aircraft centre line Fuselage bodyside Fig. 5.22 Revised aircraft wing planform Lengthening the fuselage has the effect of increasing the tail effectiveness. This may permit either a traditional low tailplane/fin arrangement, or more likely, a twin fin/tail butterfly layout. Subsequent wind tunnel tests and CFD modelling would be necessary to define the best tail arrangement. In the revised layout a butterfly tail will be shown to illustrate this option. It is now possible to redraw the baseline layout to account for the above changes. At the same time it is possible to add more details to the geometry (Figure 5.23). 5.10 Further work With the new baseline aircraft drawing available and increased confidence in the aircraft layout it is possible to start a more detailed analyses of the aircraft. We start this next stage by estimating the mass of each component using the new aircraft geometry as input data for detailed mass predictions. Such equations can be found in most aircraft design textbooks. These formulae have, in general, been derived from data of existing (therefore older) aircraft. As our aircraft will be built using mate- rials and manufacturing methods that have been shown to provide weight savings it will be necessary to apply technology factors to reduce the mass predicted by these older aircraft related methods. The factors that are applied must correspond to the expected degree of mass reduction. Different structural components will require indi- vidual factors depending on their layout. For example, the wing structure is more likely to benefit from a change to composite material than the fuselage. The fuselage has many more structural cut-outs and detachable access panels than the wing which makes it less suitable. The mass reduction factors for composite materials may vary between 95 and 75 per cent. The lower value relates to an all-composite structure (e.g. as used for control surfaces and fin structure). “chap05” — 2003/3/10 — page 135 — #35 Project study: military training system 135 metres Extra equipment Extra fuel tank 17° 15° 15° Conformal systems pack 0 1 2 fuel Fig. 5.23 Revised baseline aircraft layout Aspects other than the choice of structural material may also influence the estimation of component mass. Such features may include the requirement for more sophistication in aircraft systems to accommodate the remote instructor concept, the requirements related to the proposal for variability in the flight control and handling qualities of the aircraft to suit basic and advanced training, and the adoption of advanced technology weapon management systems. All such issues and many more will eventually need to be carefully considered when finalising the mass of aircraft components. When all the component mass estimations have been completed it will be possible to produce a detailed list in the form of an aircraft mass statement. Apart from identifying various aircraft load states, the list can be used to determine aircraft centre of gravity positions. As the aircraft will be used in different training scenarios (e.g. basic aircraft handling experience to full weapon training) it is necessary to determine the aircraft centre of gravity range for different overall loading conditions. With this information it will be possible to balance the aircraft (see Chapter 2, section 2.6.2) and to accurately “chap05” — 2003/3/10 — page 136 — #36 136 Aircraft Design Projects position the wing longitudinally along the fuselage. Up to this point in the design process the wing has been positioned by eye (i.e. guessed). With the wing position suitably adjusted and a knowledge of the aircraft masses and centre of gravity positions, it is now possible to check the effectiveness of the tail surfaces in providing adequate stability and control forces. Until now the tail sizes have been based on the area ratio and tail volume coefficient values derived from existing aircraft. It is now possible to analyse the control surfaces in more detail to see if they are suitably sized. The previously crude methods used to determine the aircraft drag coefficients can now be replaced by more detailed procedures. Using the geometry and layout shown in Figure 5.23 it is possible to use component drag build-up techniques or panel methods to determine more accurate drag coefficients for the aircraft in different configurations (flap, undercarriage and weapon deployments). Aircraft design textbooks adequately describe how such methods can be used. Likewise, more accurate predictions can now be made for the aircraft lift coefficient at various flap settings. Before attempting to reassess aircraft performance it is necessary to produce a more accurate prediction of engine performance. If an existing engine is to be used it may be possible to obtain such data from the engine manufacturer. If this is not feasible it will be necessary to devise data from textbooks and other reference material. It may be possible to adapt data available for a known engine of similar type (e.g. equivalent bypass and pressure ratios) by scaling the performance and sizes. Design textbooks suggest suitable relationships to allow such scaling. More detailed aircraft performance estimations will be centred on point performance. The results will be compared to the values specified in the project brief and subsequent considerations. The crude method used previously will be replaced by flight dynamic calculations (e.g. the take-off and landing estimations will be made using step-by-step time methods). It is also possible at this stage to use the drag and engine performance estimations to conduct parametric and trade-off studies. These will be useful to confirm or adjust the values used in the layout of the aircraft geometry (for example, the selection of wing aspect ratio, taper, sweepback and thickness). Further detailed work on the aircraft layout will include: • The identification and specification of the aircraft structural framework. • The installation of various aircraft system components. This will require some additional data on the size and mass of each component in the system (e.g. APU). • A more detailed understanding of the engine installation. This will include the mounting arrangement and access requirements. It will also be necessary to consider the intake and nozzle geometry in more detail. • Investigate the landing gear mountings and the required retraction geometry. • Make a more accurate evaluation of the internal fuel tank volumes (wing and fuselage tanks). • Detailed considerations of the layout requirements for wing control surfaces including flap geometry. It is obvious that the above list of topics requires a great deal of extra work. All of this is necessary in order to draw the final baseline layout. It would be wasteful to do all of this work without first reviewing the project and considering the overall objectives against the predicted design. The following section outlines the nature of such a review process. “chap05” — 2003/3/10 — page 137 — #37 Project study: military training system 137 5.11 Study review There are several different ways in which a design review can be conducted. At the higher level a technique known as a SWOT (strengths, weaknesses, opportunities, threats) analysis can be used. At a lower (more detailed) level an analysis similar to that described in section 2.10.2 could be followed. In this study we will adopt the SWOT method as this will illustrate the use of this technique in a design context. It must be emphasised that the low- and high-level methods of review are not mutually exclusive and that in some projects it is advisable to use both. Before starting the review it must be mentioned that the descriptions below do not constitute a complete analysis. A project of this complexity has many facets and it would be too extensive to cover all of them here. The intention is to provide a guide to the main issues that have arisen in the preceding work. 5.11.1 Strengths The most obvious advantage of this project lies in the overall life cycle cost (LCC) savings that are expected from introducing a new advanced technology, training system, approach. If such savings cannot be shown it will be difficult to ‘sell’ the new system to established air forces. The savings will accrue from the lighter modern aircraft. The use of composites will increase the purchase cost of the aircraft based on the price per unit weight. This would also require extra stringency in inspection of the structure. More elaborate systems will also increase the aircraft first cost. However, the new concept would avoid duplicity of aircraft types in the basic to advanced phase and this will reduce life cycle costs. In addition, the aircrew will have received a higher standard of training from the advanced training system, a consequential reduction of OCR training cost. The second most powerful advantage for the new concept lies in the ability of the aircraft to more closely match modern fast-jet performance than is currently possible with training aircraft that were originally conceived and designed in the 1970s. Another strength of the new system is the total integration of modern flight and ground-based systems into a total system design approach. Upgraded older aircraft types are not capable of achieving this aspect of the training system. Many more advantages could be listed for the system. How many can you identify? 5.11.2 Weaknesses There are three principal weaknesses to the project as currently envisaged. To reduce these deficiencies, if at all possible, it will be necessary to devise strategies or modifications to our design. The main and intrinsic difficulty lies in the conservative nature of all flight train- ing organisations. This is a natural trait as they take responsibility of human life and national security. As such they will be highly sceptical of the potential advantages of conducting advanced training in a single seat aircraft with a remote instructor. For our concept, as we currently envisage it, this difficulty is insurmountable. Therefore a change of design strategy must be considered to save the credibility of the project. It will be necessary to extend the design concept to encompass a two-seat trainer through- out the full (basic to advanced) training programme. The remote instructor concept can be developed as a separate part of the aircraft/system development programme (i.e. flight testing the aircraft without the instructor present as a proof of concept). [...]... 4.42 5. 08 5. 03 – 4.78 7.79 4.88 5. 15 4.67 4 .57 6.41 6.4 5. 08 8.08 8.16 8 .53 7.79 4 .57 5. 85 5.08 8.23 6.22 5. 12 6.97 6.06 6.28 7 .57 7.88 6.27 7.21 4.27 7.49 6.6 8.0 3.7 10.8 10.6 9.6 7.7 3.3 4.7 6.0 9.0 236 1 65 199 2 95 1 65 172 440 227 2 15 191 118 340 290 340 473 290 326 7 15 363 3 25 2 95 238 0.69 0 .57 0 .59 0.62 0 .57 0 .53 0.62 0.63 0.66 0. 65 0 .50 53 6 55 6 479 766 453 422 890 56 8 442 678 312 75 49 45 88 19... point References Textbooks for military aircraft design and performance: 1 Raymer, D P., Aircraft Design: A Conceptual Approach, AIAA Education Series, 1999, ISBN 1 -5 634 7-2 8 1-0 “chap 05 — 2003/3/10 — page 141 — #41 141 142 Aircraft Design Projects 2 Brandt, S A et al., Introduction to Aeronautics: A Design Perspective, AIAA Education Series, 1997, ISBN 1 -5 634 7-2 5 0-3 3 McCormick, B W., Aerodynamics, Aeronautics... Aerodynamics, Aeronautics and Flight Mechanics, Wiley and Sons, 1979, ISBN 0-4 7 1-0 303 2 -5 4 Nicolai, L M., Fundamentals of Aircraft Design, METS Inc., San Jose, California 951 20, USA, 1984 5 Mattingly, J D., Aircraft Engine Design, AIAA Education Series, 1987, ISBN 0-9 3040 3-2 3-1 The following publication is also useful in collecting data on existing aircraft: Aviation Week Source Book, published annually in January... 0 feet 5 0 metres 1 .5 Fig 6.4 ‘Conventional’ initial general arrangement “chap06” — 2003/3/10 — page 159 — #17 159 160 Aircraft Design Projects Aircraft data O/A length O/A height Wing span Wing area Wing AR Wing sweep LE Wing thickness Canard area Canard span Prop diam 12 ft /3.7 m 4 .5 ft /1.4 m 22 .5 ft /6.9 m 66 sq ft /6.1 sq m 7.7 23° 15% 9 sq ft /0.84 sq m 10 ft /3 m 4.0 ft /1.2 m 0 feet 5 0 metres... fuel tank 0 1 2 metres + Fig 5. 25 Single-seat aircraft variant 5. 12 Postscript This study has demonstrated how project design decisions may change as the aircraft is more thoroughly understood This demonstrates the iterative nature of conceptual design It is possible for students to continue this project into the next iterative stage using the final aircraft drawings (Figure 5. 25) as the starting point... the low-voltage DC supply from the fuel cells to the inverter (7) and then to the electric motor (9) The controller (10) provides the overall system control This includes: • the input to the motor (via the converter and inverter), • the start-up sequence, “chap06” — 2003/3/10 — page 155 — #13 155 156 Aircraft Design Projects Methane 1 Water 2 Reformed gas oxidiser Fuel reformer 3 Flue gas H 5 Fuel... when designing aircraft that incorporate advanced technology features “chap06” — 2003/3/10 — page 156 — #14 Project study: electric-powered racing aircraft Table 6.2 Component Mass (kg) Dimensions (m) Fuel cell Reformer Compressor Inverter Motor Battery Total 28.8 53 .6 4 .5 5.0 26.0 3.0 120.9 0.44 × 0.20 × 0.13 0.61 × 0.31 × 0.31 0.08 dia × 0.04 0.24 × 0.24 × 0.12 0.23 dia × 0.20 0. 15 × 0. 15 × 0. 15 112... surplus of high-powered, “chap06” — 2003/3/10 — page 1 45 — #3 1 45 146 Aircraft Design Projects mass-produced, ex-military aircraft These were introduced into air racing but the purists argued that this was not in the spirit of the original sport Such aircraft were raced in an open classification category that still exists To return to the original racing concept and to make the sport more pilot-centred,... annually in January This handbook is a useful source of general aeronautical data: AIAA Aerospace Design Engineers Guide, 1998, ISBN 1 -5 634 7-2 83X 1 “chap 05 — 2003/3/10 — page 142 — #42 6 Project study: electric-powered racing aircraft Existing Formula 1 racers “chap06” — 2003/3/10 — page 143 — #1 144 Aircraft Design Projects 6.1 Introduction This project is the direct result of collaboration between aeronautical... Nemesis designers have paid great attention to detail design (e.g one piece wheel fairings) This is a good overall strategy that is worth considering by all new racing aircraft design teams AR -5 This aircraft is another example of a modern low drag design Although powered by only a 65 hp motor it set a world record in 1992 of 213 mph (which gave a remarkable 3.28 mph per hp) It has a completely, all-composite, . “chap 05 — 2003/3/10 — page 128 — #28 128 Aircraft Design Projects Climb rate (m/s) sea level 25 000 ft 36 000 ft 30 40 50 60 0 –20 –10 10 20 Aircraft speed (m/s) 0 50 100 150 200 250 300 Fig. 5. 19. aircraft design and performance: 1 Raymer, D. P., Aircraft Design: A Conceptual Approach, AIAA Education Series, 1999, ISBN 1 -5 634 7-2 8 1-0 . “chap 05 — 2003/3/10 — page 142 — #42 142 Aircraft Design. and Sons, 1979, ISBN 0-4 7 1-0 303 2 -5 . 4 Nicolai, L. M., Fundamentals of Aircraft Design, METS Inc., San Jose, California 951 20, USA, 1984. 5 Mattingly, J. D., Aircraft Engine Design, AIAA Education