Aerodynamic, dynamic and conceptual design of a fire-fighting aircraft

125

Aerodynamic, dynamic and conceptual design of a fire-fighting aircraft Z Goraj1 , A Frydrychewicz2 , E C P Ransom 3 ¤ , A Self 3 and P Wagstaff 3 1 Warsaw University of Technology, Poland 2 PZL-WSK-Okecie, Poland 3 Faculty of Technology, School of Engineering, Kingston University, London, UK

Abstract: This paper presents an evaluation of available aircraft types and demonstrates the lack of a low cost aircraft optimized for fighting forest fires, spreading viscous liquids for land reclamation and spraying pesticides over forest areas. It shows that among critical areas demanding special consideration are (a) the development of mathematical models for the water bomb–aircraft separation and the aircraft transient dynamics following separation, (b) the identification of parameters influencing the coherence of the water column and effectiveness of water delivery for fire fighting, (c) choice of aircraft configuration and (d) hopper configuration. Design and numerical analysis have led to the selection of the biplane as the best aircraft. The main aerodynamic characteristics for the selected aircraft have been computed by means of panel methods, the so-called modified Hess method for thick wings and bodies and/or the vortex lattice method for thin lifting surfaces. Different gaps and staggers and their influence on aerodynamic characteristics have been analysed. It has been found that shifting the lower wing rearwards (positive stagger) while keeping the angle of attack constant results in a small increase of induced drag and an almost constant value of the lift curve slope as well as an increase in the pitching moment curve slope. The increase of drag is disadvantageous, whereas the increase of the pitching moment curve slope means that the neutral point of stability is moved forward (a disadvantage from the stability point of view). The influence of the biplane configuration on downwash in the vicinity of the horizontal plane and aircraft dynamic stability is also discussed. Another important concept — developed at PZL-Okecie and presented in this paper — consists in using parts from existing aircraft. The pilot’s cabin, the rear part of the fuselage with control surfaces and wings originate from the PZL-106 ‘Kruk’. This diminishes the cost of design and prototype construction as well as of the cost of aircraft production. It has been shown that an important cost factor in the operation of such a fire-fighting aircraft is the weight of the agent which may be carried for the same fuel consumption. This cost factor, representing the economical efficiency of a fire-fighting aircraft, has been computed and compared for a number of fire-fighting aircraft. The design under consideration (called the PZL-240 ‘Pelikan’) has the above-mentioned factor equal to 14, whereas the average value for other aircraft is about 8. Keywords: aircraft design, biplane, panel method, flight dynamics, aerodynamics

NOTATION Ae Aef b c cD cL

cm equivalent monoplane aspect ratio effective aspect ratio wing span wing chord drag coefficient lift coefficient

The MS was received on 14 August 2000 and was accepted after revision for publication on 25 April 2001. ¤ Correspondin g author: Faculty of Technology, School of Engineering, Kingston University, Friars Avenue, Roehampton Vale, London SW15 3DW, UK. G03200 # IMechE 2001

Ca Ck , Cl, Cn Cp Di D, L, M H Jy K L MAC n

pitching moment coefficient, around the mean quarter-chord point A MAC aerodynamic influence coefficients pressure coefficient ˆ ( p ¡ p1 )=q induced drag lift, drag and pitching moment for whole aircraft gap moment of inertia about y axis Munk’s span factor lift force mean aerodynamic chord load coefficient Proc Instn Mech Engrs Vol 215 Part G

126

Z GORAJ, A FRYDRYCHEWICZ, E C P RANSOM, A SELF AND P WAGSTAFF

N q Q S U, W x x0 , z0 xA, xC , zC , xN y á ä1 , ä2

number of panels dynamic pressure ˆ 0:5rV 2 pitch rate wing area speed components stagger coordinates of aircraft position in the ground-fixed axis system coordinates of points A, C, N with respect to nose of MAC coordinate along wing span angle of attack influence coefficient of sweep angle and taper ratio respectively downwash pitch angle shorter span to longer span ratio; doublet strength coordinate along wing chord Prandtl’s interference factor; source strength disturbance velocity potential velocity potential

å õ í ê ó j ¼ Subscripts A C e E H i l N u

1

mean quarter-chord point mass centre of the whole aircraft equivalent empty aircraft hopper induced lower neutral point of static stability upper

INTRODUCTION

Forest fires are a major problem in many parts of the world. Fire areas are often very far from centres of communication

and access may be difficult. For this reason aircraft have proved to be a powerful means of dealing with this continuing problem. The potential market for specialized fire-fighting aircraft is small by comparison with other types and with few exceptions existing aircraft used to control fires are conversions. These may be small agricultural aircraft that are adapted for fire control duties or very large aircraft, such as the Hercules C130, that are capable of delivering many tonnes of water to the fire zone. In the former case the aircraft are often too small to be effective and in the latter while often very effective they are expensive in initial capital cost and operational costs and require base facilities which are costly to provide and operate. Such expenditure is beyond the resources of many countries. Current dedicated agricultural aircraft are very different from the original crop-dusters which were often converted war-surplus aircraft. A good example of a very efficient agricultural aircraft today is the Gippsland GA200. By making use of advanced computer-aided design analysis, the designers have achieved a remarkable performance for this category of aircraft. The GA-200 is able to carry 760 l in its hopper despite the aircraft being powered by no more than one 240 h.p. engine [1]. There is a tendency for the heavier multi-engined agricultural aircraft to be replaced by smaller, more economical single turbine engined aircraft. Typical examples of agricultural aircraft, their significant technical data and some economic factors are compared in Table 1 [2]. The last three columns of Table 1 give (a) hopper capacity (l) divided by the maximum take-off mass (kg), (b) hopper capacity (l) divided by the empty mass (kg) and (c) hopper capacity (l) divided by the available engine power (kW). The most economically efficient aircraft have the highest values in these columns. Some features which are desirable for agricultural aircraft are not desirable for fire fighting. A typical example is the wing span (or aspect ratio if the aircraft weight is assumed to be the same). One of the main tasks of agricultural aircraft is to spread fertilizing granulates or to spray protective, anti-pest liquids. The larger the span the more effective is the distribution of fertilizer or liquid spray. A large span can be

Table 1 Technical data and economical factors of different 14 aircraft, mostly agricultural, used in fire fighting Aircraft

Power P (kW)

Empty mass m (kg)

Maximum takeoff mass M (kg)

Hopper capacity H (l)

H/M (l/kg)

H/m (l/kg)

H/P (l/kW)

Ag-Cat SuperB Air Tractor AT-502B An-2 Ayres 660 Turbo Thrush Cessna A188B AG Husky Cresco 600 Embraer 201A Ipanema Gippsland GA-200 Fatman Let Z-37A Cmelak Piper PA-36 M-18B Dromader PZL-106BT ‘Kruk’ Weatherley 620B Canadair CL-215T

448 507 745 788 224 447 224 194 235 213 721 544 338 2 3 1 720

1 656 1 996 1 996 2 700 982 1 270 1 011 770 1 043 987 2 800 1 750 1 288 12 400

3 184 4 300 5 500 5 682 1 905 3 175 1 550 1 315 1 850 1 769 5 300 3 500 1 814 21 000

1 514 1 892 1 960 2 508 1 060 1 847 950 776 650 852 2 500 1 500 1 268 7 000

0.47 0.44 0.35 0.44 0.56 0.58 0.61 0.59 0.35 0.48 0.47 0.43 0.70 0.3

0.91 0.95 0.98 0.92 1.08 1.45 0.94 1.0 0.62 0.86 0.89 0.86 0.98 0.57

3.37 3.73 2.63 3.18 4.73 4.13 4.24 4.0 2.76 4.0 3.47 2.76 3.75 2.03

Proc Instn Mech Engrs Vol 215 Part G

G03200 # IMechE 2001

AERODYNAMIC, DYNAMIC AND CONCEPTUAL DESIGN OF A FIRE-FIGHTING AIRCRAFT

dangerous during fire fighting operations because of the presence of high thermal gradients and vertical gusts. Some aircraft such as the Cessna AG Husky, Cresco 600 and Gippsland GA-200, which appear to be the most efficient aircraft by the criteria of Table 1, have smaller capacity hoppers. Thus they cannot be very effective in the firefighting role. Aircraft safety during hazardous operations is of paramount importance. The configuration has to be such that the aircraft is stable and controllable in the fire zone, particularly at the moment of water release, and immediately afterwards, that is under transient conditions. The aircraft should be agile, very stable with respect to roll and yaw and not very sensitive to temperature gradient. It is recognized that the hopper is the key component of the aircraft since it is the principal load-bearing structure of the aircraft. It is attached between the front and rear fuselage and supports both the wings and the undercarriage. In addition it transports and delivers the cargo. The higher the hopper is, the greater the initial water speed of outflow, owing to the greater hydrostatic pressure. Also, the section ratio of the hopper may have an influence on the characteristics of the water-dropping process. Two interrelated effects are apparent. The cohesion of the water column can be destroyed if the air speed is too great and a decrease in lateral section decreases the drag force. Figure 1 shows a light agricultural aircraft, the PZLM-18, releasing water. It is clearly seen that the water bomb structure, at a cruise speed of 47 m=s, is destroyed. The air flow velocity seems to be too large and has a decisive effect on the coherence of the liquid column, which is dispersed into fine droplets.

Fig. 1

PZLM-18 agricultural aircraft (hopper of 1850 kg capacity)

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127

In this case the quenching effect on the fire is negligible. This aircraft usually drops water on a fire at V ˆ 40 m=s. An adjustment of the shape of the discharge orifice to give the column a reduced dimension transverse to the flow and an increased length parallel to the flow would reduce the tendency for the column to disperse by reducing drag. Another factor which accelerates the dispersion of water column is too small a water pressure at the hopper outlet, which results in too small a velocity at water outflow. The only way to increase the pressure at the bottom of the hopper is to construct it as tall as possible. However, because the hopper height is limited by the fuselage vertical dimension, considerable compromise is inevitable in the overall design. A number of tests were conducted by the US Forest Service [3] in California during 1993. It was found that effectiveness of fire fighting, measured in coverage level by number of litres per square metre, depends on various parameters, including the amount of agent, flowrate, type of agent (its density and viscosity), wind speed and its direction, aircraft speed and drop height. For example, the Air Tractor AT-802F is endowed with a computerized pilot interference system, allowing the pilot to select an average coverage level ranging from 0.2 to 2.5 l/m2 on salvo drops. These factors aid the selection of the preferred aircraft and hopper configuration. The biplane is considered as a possible design configuration for fire-fighting applications. Widely used at the beginning of heavier than air flight, it was rapidly superseded by cantilevered monoplane designs, owing to strong interference effects especially at high speeds. A number of papers [4–11] have been published which provide comparisons between different two- and threesurface configurations. The biplane may be considered a three-surface arrangement and the monoplane as a twosurface arrangement. References [4] to [7] focus on minimum aircraft induced drag versus gap and stagger. These analyses are based on Munk’s stagger theorem [12] and Prandtl’s relation for the induced drag [13]. Because of the limitation of Prandtl’s formula to elliptically loaded wings, Laitone [7, 8] generalized it to include more realistic distributions and to describe the wing mutual downwash which rotates the lifting force vectors. Butler [9] showed that the maximum ‘induced thrust’ can be reached for zero gap with a canard configuration. Kroo [10] found that interference terms can be beneficial and that the induced drag is lower than if the two wings are infinitely far apart. However, he wrote that ‘in the context of preliminary design optimisation, model panel codes may prove prohibitively time-consuming and expensive’. Kendall [11] considered longitudinal trim and static stability in addition to an analysis of induced drag. He analysed a number of three-surface configurations for a range of gaps and concluded that such a design can attain minimum induced drag without compromising the conditions for longitudinal trim and static stability over a useful range of locations for the centre of gravity. Proc Instn Mech Engrs Vol 215 Part G

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2


FIRE-FIGHTING AIRCRAFT — CHARACTERISTICS REQUIRED

The most important, general characteristics of fire-fighting aircraft are listed below:

It is also worth mentioning that Norton [18], as early as in 1918, noticed that ‘. . .possible stagger greatly restricts the centre of pressure travel’, thus making the stability problem simpler. Biplanes can be very attractive at lower speeds, for the following reasons [19]:

(a) a hopper capacity of at least 4 t of water is the minimum for effective action; (b) the fire-extinguishing agent should be discharged at a controlled rate, including the possibility of releasing a coherent column of fluid in the form of a water bomb; (c) three hopper filling systems are required, i.e. a groundbased system, an onboard system and an in-flight system to refill from a water surface; (d) good cockpit visibility, particularly forward and downwards; (e) a minimal wing span to reduce the effects of vertical turbulent gusts over the fire zone; (f) ability to operate from uneven or unprepared landing fields, leading to the need for aircraft to have a high lift coefficient, a strong undercarriage and large diameter, low pressure tyres; (g) mild stall characteristics—when at critical angles of attack, the aircraft should lose altitude without stalling (this feature is relevant for agricultural aircraft too); (h) small turn radius.

1. A relatively big wing area with a high effective wing aspect ratio can be obtained for a moderate wing span. 2. A relatively stiff wing structure can be built for a large wing area and moderate structure mass. 3. A relatively small longitudinal moment of inertia at rather high mass of structure can be attained. 4. Smaller overall span makes the biplane more manoeuvrable. 5. Take-off and landing distance can be shorter than that for a monoplane of the same weight. 6. For the same wing area, the biplane can be aerodynamically more efficient than the equivalent monoplane, i.e. lower drag at the same lift or at the same equivalent wing aspect ratio.

The above-mentioned characteristics lead to a number of specific design requirements:

4

(a) a tall hopper in order to maximize the hydrostatic pressure at the base of the tank, which allows faster water release and encourages the formation of a coherent column of fluid; (b) a moderate wing span for a relatively large wing area, for which the biplane configuration is a possible option; (c) long span slots or flaps, to maintain high lift in turns. The second of these requirements is directly opposite to the desired features for agricultural aircraft. Most have a relatively high aspect ratio wing in order to facilitate the spread of granulates or liquid spray over a wide area. The weaknesses of the biplane configuration are its poor cockpit visibility and difficulties in arranging for the rapid loading of the hopper. Both drawbacks may be minimized by using wing stagger.

3

BIPLANES—BACKGROUND AND ADVANTAGES

Many papers [12– 14] and project descriptions [15,16], related to biplanes, have been published. A paper of particular interest is one by Nenadovitch, published in 1936 in Saint-Cyr [17], in which two-dimensional characteristics of a biplane versus gap, stagger and decalage can be found. Proc Instn Mech Engrs Vol 215 Part G

All of the above factors show that the biplane configuration for certain special cases may have superior characteristics.

INITIAL DESIGN CONSIDERATIONS—AREAS WHICH NEED TO BE INVESTIGATED PRIOR TO DESIGN

The above analysis suggests that a biplane configuration would provide many desirable characteristics needed for a fire-fighting aircraft. However, a number of design decisions have to be taken and it is necessary to specify some basic parameters. The areas which need to be investigated before design decisions can be taken are listed below: 1. Selection of specific mathematical models. As usual, this is a compromise between simple models which operate more rapidly or more sophisticated models which yield more credible results but take longer to run. For example, gap and stagger can be obtained from Prandtl–Munk’s stagger theorem, from panel or from field methods [20]. Similarly, investigation of dynamic response after the water drop can be performed either by including the fact that water outflow from the hopper changes lift and pitching moment (unsteady aerodynamic model) or under the assumption that the water outflow does not change the aerodynamic characteristics (quasi-steady aerodynamic model). Another problem of that type which has to be solved is the modelling of the shape of the water column. This can be done either using computational fluid dynamics (CFD) for the simulation of two-phase flow (water and air) or by experimental methods, for example wind tunnel tests or in-flight measurements. G03200 # IMechE 2001


2. Choice of specific values of the important design parameters, based on the results of the above analyses. Among them are gap, stagger, decalage, tailplane volume ratio and position of the centre of mass. 3. The hopper design. This key component has a multifunction role which needs careful analysis. The factors to be considered include structural design, material, particularly with reference to corrosion aspects, construction method, filling techniques for fluid or granulates and release modes.

Di ˆ

L21 2ó L1 L2 L22 ‡ ‡ ðqb21 ðqb1 b2 ðqb22

129

(2)

If it is assumed that the lift coefficient is the same for both wings (there is no decalage) and the symbols c1 , c2 are introduced for the chords of both wings, the induced drag coefficient, obtained from equation (2), is C Di ˆ

C 2L 2 (c ‡ 2ó c1 c2 ‡ c22 ) ðS 1

(3)

where the expression 5

AERODYNAMIC CONSIDERATIONS

Most of the published design data related to the biplane are based on classical lifting-line theory and Munk’s stagger theorem [4,13]. For a biplane with elliptic planform wings the added induced drag on one wing due to the proximity of the other is given by ¢Di ˆ

ó L1 L2 ðqb1 b2

(1)

where L1 , L2 and b1 , b2 are the lifts and spans respectively, q is the dynamic pressure and ó is the Prandtl interference factor (shown in Fig. 2), which depends on the ratio of gap to average span and on the ratio of the shorter to longer span [21,22]. In practical configurations the ratio of gap to mean span ( H/b) is never less than 0.05 and never greater than 0.25. The total added drag has twice the value of that for single wing, so the total induced drag of a biplane is

Fig. 2 G03200 # IMechE 2001

Aˆ

c21

S ‡ 2ó c1 c2 ‡ c22

(4)

is called the equivalent monoplane aspect ratio and (5)

S ˆ c1 b1 ‡ c2 b2

is the total wing area. The equivalent monoplane span for the special case of equal wing spans (í ˆ 1) can be computed [21] from the relation (6)

be ˆ Kb where Munk’s span factor is equal to Kˆ

r 2 1‡ó

(7)

Prandtl’s interference factor over a limited range of the gap to mean span ratio [21] Proc Instn Mech Engrs Vol 215 Part G

130


With all the approximations of the classical lifting-line theory and Munk’s stagger theorem, which is the basis of the above considerations, its application is simple. The results obtained can be compared with more accurate results generated by panel methods and in this sense provide a check for accuracy, especially for higher aspect ratio wings. In this paper all design decisions and recommendations have been based on results obtained from the application of panel methods for thick surfaces. Panel methods originate from Laplace’s equation, ¢¼ ˆ 0, for the velocity potential ¼ and disturbance velocity potential j [23, 24]. The solution of Laplace’s equation for the full velocity potential has the following form: 1 ¼(x, y, z) ˆ 4ð

³ ´ … 1 body 1 @ 1 dS ¡ í ó dS ‡ ¼1 4ð r wake @ n r (8)

… body

The boundary conditions are as follows: 1. The inner Dirichlet condition on the surface of the body is 1 4ð

³ ´ … 1 body 1 @ 1 dS ¡ í ó dS ˆ 0 4ð r wake @ n r

… body

(9)

where Doublet strength:

í ˆ ¡(¼ ¡ ¼i )

(10)

Source strength:

ó ˆ @ í=@ n

(11)

2. The Kutta–Zhukovsky condition at the trailing edge is ¢ p(x, y)TE ˆ 0

(12)

3. The condition on the wake is @j(x, y) ˆ0 @x

(13)

If it is assumed that the inner velocity potential ¼i is equal to the potential at infinity, ¼1 , then from equation (8) it is possible to obtain an integral equation in the form of equation (9). Equation (9) can be approximated by a set of linear equations with unknown strength of doublets í (being constant over each panel), i.e. N X kˆ1

C k ík ‡

Nw X lˆ1

C l íl ‡

N X kˆ1

Bkó k ˆ 0

(14)

where C k , Cl and B k are aerodynamic influence coefficients: Proc Instn Mech Engrs Vol 215 Part G

1 Ck ˆ 4ð

Bk ˆ ¡

³ ´ @ 1 dS k 1234 @ n r k

…S

1 4ð

…S

1234

(15a)

1 dS k rk

(15b)

and N ˆ number of panels over the whole aircraft N w ˆ number of panels over the wake S1234 ˆ area of the kth panel

To investigate the effect of configuration changes, i.e. gap, stagger and tailplane position, a number of calculations have been carried out to determine lift, induced drag, pitching moment and downwash. These results assisted in arriving at the preliminary configuration. For these calculations it was assumed that both wings were of the same geometry (constant section NACA 4409, untwisted with an aspect ratio of 6, zero sweep, parallel chord). The variables investigated were gap and stagger. The decalage angle was initially assumed to be zero, although reference [25] suggests that a negative decalage of about ¡68 gives the best aerodynamic efficiency. In Figs 3 and 4 the horizontal axis shows the dimensionless coordinate ê/c, measured parallel to the wing chord, and the vertical axis denotes the dimensionless pressure coefficient C p ˆ ( p ¡ p1 )/0.5 r1 V 21. The upper graphs of Figs 3 and 4 show the pressure distribution over the upper and lower surfaces of the top wing; the lower part of figure shows the same over the lower wing. Curves in Fig. 3, obtained at the root of wing (2 y=b ˆ 0) and at angle of attack á ˆ 08, correspond to different gaps, H=b ˆ f0:0415, 0.083, 0.166}, and are compared with the curves obtained for a single wing. It is seen that the negative pressure coefficient distribution over the upper surface of the top wing and the positive pressure coefficient distribution over the lower surface of bottom wing do not depend on gap H/b since they are the same as for a single wing. However, the positive pressure coefficient distribution over the lower surface of the top wing decreases and for the case when the gap is equal to H =b ˆ 0:0415 the pressure coefficient becomes negative. Overpressure occurs also at H =b ˆ 0:083 and 0.116. This phenomenon may be explained by the interference effect that one wing has on the other. For the largest gap ( H=b ˆ 0:166) there is little difference from the isolated wing. However, as the gap is reduced, a strong outflow develops, which causes a decrease in pressure. The effect on the upper surface of the lower wing is negligible, particularly at the larger gaps. Only at the smallest gap ( H =b ˆ 0:0415) is the effect noticeable and then over the centre chord region. Stagger strongly influences the pressure distribution over the lower and upper surfaces of the top wing and the upper surface of the bottom wing (see Fig. 4). Shifting the lower wing rearwards with respect to the top wing (x . 0) G03200 # IMechE 2001


Fig. 3

Pressure distribution over the biplane (at the root of wing) versus wing gap, á ˆ 08, x=c ˆ 0

involves a simultaneous increase in negative pressure coefficient (over the upper surface) and of positive pressure coefficient over the lower surface. The corresponding effect on the lower wing is a decrease in upper surface negative pressure coefficient, with little change to the positive pressure coefficient on the lower surface. The overall effect is an increase of lift on the top wing and a decrease of lift on the bottom wing. Figures 5 and 6 show spanwise lift distributions, obtained after the integration of local pressures. For unstaggered wings it is seen that the upper wing generates slightly more lift than the lower wing. Stagger, positive or negative, significantly affects the pressure distribution, and hence lift, on the upper wing with respect to the unstaggered arrangement. The effect of stagger on the lower wing is to modify the pressure distributions on the upper and lower surfaces. However, as indicated in the lower part of Fig. 5 the changes are less significant than for the upper wing. Figure 6 shows the effect of stagger on the total lift generated by both wings acting together. Positive stagger (upper wing ahead of lower wing) shows a significant but not large increase in lift. The integration of the incremental lift against span yields ¢CL ˆ 0:025. Stagger may be employed to improve visibility on landing, to aid stability or for aesthetic reasons [21]. G03200 # IMechE 2001

Fig. 4

131

Pressure distribution over the biplane (at the root of the wing) for three different staggers, á ˆ 08, H=b ˆ 0:083

Some general characteristics for aircraft having a biplane configuration are shown in Figs 7 to 10. For simplicity an uncambered, symmetrical wing section has been chosen (NACA 0012). Figure 7 illustrates the effect of gap on the pitching and drag coefficients, when the lift coefficient CL is equal to 0.5. Two wing designs are examined, one rectangular and one tapered with a taper ratio of 0.38. Both are unswept, untwisted and have an aspect ratio AR ˆ 12. The pitching moment and drag coefficients decrease monotonically with relative gap, H/b. The single isolated points at the edge of the figure denote the boundary values where the gap goes to infinity. It is also clear from Fig. 7 that the pitching moment coefficient for a biplane, when the relative gap is greater than 0.5, is approximately equal to the corresponding value for the equivalent single wing. This applies to both rectangular and tapered wing configurations. The induced drag coefficient for a rectangular winged biplane with a relative gap H =b ˆ 0:84 is very close to the value for the equivalent single wing. However, for tapered wings, the induced drag coefficient differs from the corresponding value for the equivalent monoplane by approximately 12 per cent. For small relative gaps ( H =b , 0:08) the induced drag coefficient of a biplane at equilibrium (for CL ˆ 0:5) is almost twice the corresponding value for a monoplane. Proc Instn Mech Engrs Vol 215 Part G

132


Fig. 5

Spanwise lift distribution, for top and bottom wings acting separately. Two different angles of attack (á ˆ 58 for upper curves and á ˆ 08 for lower curves) and three different staggers, x=c ˆ ¡0:5, 0, 0.5; H=b ˆ 0:083

The induced drag coefficients for biplanes with tapered and rectangular wings are compared with those computed on the basis of Prandtl– Munk’s biplane theory [equations (2) and (3)] and are shown in Fig. 8 for a limited range of relative gap. From this figure (corresponding to a wing aspect ratio Am ˆ 12) it can be concluded that Prandtl– Munk’s theory gives results consistent with those obtained from panel methods. In particular there is good correlation for tapered wings. Unswept, tapered wings have a spanwise lift distribution very close to the ideal elliptic case and that is the reason why the induced drag of tapered wings is closer to that obtained from Prandtl–Munk’s theory. However, even for rectangular wings, the difference between results obtained from panel methods and the Prandtl–Munk theory is small, the largest difference being less than 2 per cent. Figures 9 and 10 compare the lift, drag, pitching moment coefficients and the position of the centre of pressure for a biplane of zero stagger (full curve) and for a monoplane (single, isolated points), for both tapered (Fig. 9) and rectangular (Fig. 10) wings. These have been computed using the Hess panel method for thick wings [equations (8) to (15)] with corresponding coefficients for thin wings (broken curves), computed by use of a vortex lattice method (VLM) [26]. Consistency of lift is satisfactory. To Proc Instn Mech Engrs Vol 215 Part G

explain why the induced drag (Figs 9 and 10) first increases and then decreases, it is necessary to consider the graph of lift against gap and how the aspect ratio of the equivalent wing (see Section 6) varies with gap (see Table 2). The graph of lift against gap (Figs 9 and 10) increases strongly at small gaps and less steeply for gaps greater than 0.2. Equivalent wing aspect ratios increase with gap uniformly. At small gaps the induced drag increases because lift increases and then decreases because of an increase in the equivalent wing aspect ratio. Downwash for thick wings has been computed by means of panel methods (Figs 11 and 12). Firstly, the distribution of doublets and sources was found. Then the velocities in the vicinity of the horizontal tail (dimensionless tail arm lH/c was assumed to be 5), induced by the distribution of doublets and sources over the whole configuration, were computed. Components of velocities, normal to the undisturbed flow velocity vector and divided by its value, give the local downwash. Isolines of downwash (positive value means that flow streams are deflected down, i.e. that they decrease the angles of attack) are shown in Figs 11 and 12 and correspond to an angle of attack á ˆ 58. Figure 11 shows the downwash behind a monoplane. Isolines corresponding to the edge vortices as well as an increase of local angle of attack, outside of the wing, are G03200 # IMechE 2001


Fig. 6

Spanwise lift distribution for biplane, for top and bottom wings acting together. Two different angles of attack (á ˆ 58 for upper curves and á ˆ 08 for lower curves) and three different staggers (x=c ˆ ¡0:5, 0.0, 0.5); H=b ˆ 0:083

Fig. 7

Pitching moment and induced drag coefficients versus gap for fixed lift coefficient and changeable angle of attack, x=c ˆ 0

G03200 # IMechE 2001

133


134


Fig. 8

Comparison of the induced drag coefficients between the results of panel methods and Prandtl –Munk’s theory based on lifting line theory, x=c ˆ 0

Fig. 9

Lift, induced drag, pitching moment coefficients and centre of pressure location versus gap for fixed angle of attack (tapered wings), x=c ˆ 0

seen very clearly. An increase of wing gap increases the average downwash (Fig. 12).

6

EQUIVALENT WING

The aerodynamic characteristics of a biplane are related to the area and chord of an equivalent wing. It is Proc Instn Mech Engrs Vol 215 Part G

usually assumed that the area of the equivalent wing is equal to the whole area of both wings. However, the equivalent chord may be computed by different methods. The classical mean aerodynamic chord (MAC) definition, based on the assumption regarding the equivalence of lifts and pitching moments between an original wing of an arbitrary geometry and the equivalent, rectangular wing, may be calculated according to the formula G03200 # IMechE 2001


Fig. 10

135

Lift, induced drag, pitching moment coefficients and centre of pressure location versus gap for fixed angle of attack (rectangular wings), x=c ˆ 0 Table 2

Biplane aspect ratio computed according to different models 0.05

0.10

0.15

0.20

0.30

Equation (19a)

12.37

12.74

13.09

13.44

14.11

Ae ˆ b2e /Se

Equation (19b)

6.38

6.76

7.14

7.53

8.30

Figure 2 Equations (20), (5), (7) Equation (21)

0.78 6.74 12.72

0.655 7.25 13.19

0.561 7.69 13.58

0.485 8.08 13.93

0.370 8.76 14.50

CL CD,ind Aef ˆ C2L /ðCD,ind ä Ae ˆpA /(1 ef ‡ ä) be ˆ Ae Se

Panel method; Figure 10 Figure 10 Equations (22) From Engineering Sciences Data Unit [27] Equations (22) Equation (23)

0.6514 0.019 98 6.76 0.116 6.06 12.06

0.7476 0.024 25 7.34 0.116 6.58 12.57

0.7924 0.025 67 7.79 0.116 6.98 12.94

0.8183 0.025 97 8.16 0.116 7.31 13.24

0.8412 0.025 62 8.79 0.116 7.88 13.75

H/b r 4H be ˆ 1 ‡ ð b ó 2 Ae ˆp2b /S e (1 ‡ ó ) e be ˆ Ae Se

Fig. 11 G03200 # IMechE 2001

Downwashes behind the monoplane in the vicinity of horizontal tail (l H =c ˆ 5, H =b ˆ 0:0415, x=c ˆ 0, á ˆ 58) Proc Instn Mech Engrs Vol 215 Part G

136


Fig. 12

Downwashes behind the biplane in the vicinity of horizontal tail (lH =c ˆ 5, H=b ˆ 0:083, x=c ˆ 0, á ˆ 58)

… b=2

MAC ˆ Ca ˆ … 0b=2

c2 d y (16) c dy

0

but cannot be applied directly to the biplane because this formula does not include any dependence of pitching moment on induced drag and lift on gap value. Therefore, a different way to find the MAC needs to be established. The vertical location of the MAC between both wings (based on the equivalence of pitching moment) could be found from the assumption that the induced drag forces (generated on component wings) are proportional to both wing areas. Lengths d1, d2 and h1 ¡ d 1, h2 ¡ d 2 can be computed on the assumption that the moment of the drag forces about a point at equivalent chord is equal to zero (Fig. 13). The length of the MAC may be found according to the formula [21] Ca ˆ

rC au Su ‡ Cal S l rSu ‡ Sl

(17)

where r is the relative efficiency of the upper wing (loading of upper wing to loading of lower wing) and could be

Fig. 13 Proc Instn Mech Engrs Vol 215 Part G

approximated from experimental data as a function of gap and stagger. Cau and Cal denote the MACs separately computed for the upper and lower wings, from equation (16). In this model the vertical and horizontal distances of the leading edge of the MAC are found from the formulae ³ d 1 ˆ h1 1 ¡

´ eSu , eS u ‡ S i

xl ˆ

x(h1 ¡ d 1 ) h1

(18)

where x denotes stagger, x1 the horizontal location of the MAC leading edge in front of the lower wing leading edge and other parameters are shown in Fig. 13. To analyse performance it is necessary to know how drag varies with lift. The induced drag component can be computed by means of panel methods. Engineers very often estimate this component from the aspect ratio of an equivalent wing. Below two different ‘engineering approaches’ (models) and an approach based on a panel method are reviewed. Numerical results obtained from all three models are compared: 1. Prandtl’s model, based on the lifting line theory of Prandtl and Glauert. Two horseshoe vortices, representing both circulations of the real wings of a biplane, make it possible to compute the induced drag and

MAC location G03200 # IMechE 2001


pitching moment. The set of two horseshoe vortices can be replaced by one, equivalent vortex, such that the pitching moment of the biplane and that of the equivalent single wing are the same. An effective algorithm consists of computing the span of the equivalent wing, MAC and aspect ratio according to the following formulae: s 4 bl H be ˆ bu 1 ‡ ð b2u Ae ˆ

7

b2e Se

(19b)

2. The Prandtl–Munk model, based on lifting line theory [12, 13, 21]. To compute the aspect ratio of the single, equivalent wing, the Prandtl interference factor ó has to be known. The aspect ratio can be found from the following formula: 2b2 , Se (1 ‡ ó)

where b ˆ

bu ‡ bl 2

(20)

Formula (20) is equivalent to equations (4) to (7), giving the aspect ratio of the equivalent monoplane for the best lift distribution case. An equivalent aspect ratio Ae having been found, the equivalent span can be computed as be ˆ

p Ae Se

(21)

3. The model based on pressure distribution, obtained from panel methods. From the pressure distribution, the induced drag and then the equivalent effective aspect ratio can be found from the following formulae: C D,ind ˆ

symbol of the parameter calculated and the second contains the source equation or reference. The remaining columns contain the calculated data for different gaps. Assuming that the model based on the panel method gives the most accurate results it is seen that both simplified models (Prandtl’s and Prandtl– Munk’s) overestimate the aspect ratio of the equivalent monoplane wing. The relative error in the worst case is 11 per cent.

(19a)

where Ae and be are the aspect ratio and span of an equivalent monoplane wing and Se is the whole biplane wing area (Se ˆ Su ‡ Sl ).

Ae ˆ

137

C 2L C 2L Aef ! Aef ˆ ! Ae ˆ ðAef ðCD,ind 1‡ä

(22)

RESULTS OF AERODYNAMIC AND DYNAMIC INVESTIGATIONS

The following models have been selected for discussion. 7.1

Classical panel technique

Equations (8) to (15) have been used to find the steady state aerodynamic characteristics. A three-surface model (biplane ‡ tailplane) has been established to compute lift, pitching moment, induced drag and downwash. Firstly, the total weight, lifting area and wing span have been established from statistical relations and the consideration of similar aircraft. Then the wing chord and equivalent aspect ratio have been established. Selection of the gap and stagger has been carried out using panel methods (see Figs 14 to 17). From Fig. 14 it follows that increasing gap increases the lift curve slope, especially in the range 0:0415 , H =b , 0:166. Stagger has no influence on the lift curve slope (Fig. 15). The influence of gap and stagger on the pitching moment curve slope is shown in Figs 16 and 17. Pitching moment is computed around point A, which is the mean quarter-chord point. Figure 16 shows the relation between pitching moment around point A and point N (which is the neutral point of stability). Dynamic equations of motion, used in Section 7.3 for the simulation of transient processes, relate to point A. It is important to ensure the mass centre position, xC , after the water release does not go back beyond neutral point N (such a case is possible if the

where the coefficient ä ˆ ä1 ä2 includes the influence of sweep and taper ratio of an equivalent wing [24, 27]. The span of the equivalent wing may be computed from the following formula: be ˆ

p Ae Se

(23)

Aspect ratios of an equivalent monoplane wing which replaces the wings of the corresponding biplane have been computed according to the above three models. Table 2 gives the results for a biplane having the following dimensions and angle of attack: cRu ˆ cTu ˆ cRl ˆ cTl ˆ 1 m; bu ˆ bl ˆ 12 m; Se ˆ 2 3 12 m2 ˆ 24 m2 ; Au ˆ Al ˆ 12; á ˆ 108. The first column contains the G03200 # IMechE 2001

Fig. 14

Influence of gap on lift curve slope, x=c ˆ 0 Proc Instn Mech Engrs Vol 215 Part G

138


Fig. 15

Influence of H =b ˆ 0:083

stagger

on

the

lift

curve

slope,

hopper centre of gravity position is far in front of the aircraft centre of gravity position). The curves in Figs 16 and 17 indicate that as wing gap increases there is an improvement of static longitudinal stability (the slope dCm / dCL changes its value from positive to negative which means that the neutral point of stability is shifted back). Any change of stagger from the zero position, either positive or negative, results in a shifting of the neutral point of stability forward, thus reducing the margin of static stability. The general conclusion from Figs 14 and 16 is that, the bigger the gap (in the range 0:0415 , H =b , 0:166), the better the lift capacity and longitudinal stability. From Figs 15 and 17 it follows that the best choice is zero stagger. However, non-zero stagger has no influence on lift capacity and only a slight effect on stability margin. Factors other than aerodynamic strongly influence the aircraft geometry. For example, the need to have a tall hopper with provision for rapid filling and the need to provide good cockpit visibility. This has led to the selection of a gap of H =b ˆ 0:083 and a stagger of x=c ˆ 0:44. Figures 18 to 20 show average downwash curve slopes. It is seen that these slopes decrease monotonically with respect to tail arm lH /c (Fig. 18) and positive stagger (x=c . 0) (Fig. 20). However, there exists a maximum with respect to the gap H/b (Fig. 19). Longitudinal stability is improved if the downwash is reduced and the tailplane is stiff. For a compact design the tail arm must not be overlong. As a compromise the tail arm length, lh /c, is taken as 3.8. 7.2

Fig. 16

Fig. 17

Pitching moment curve slope versus gap, x=c ˆ 0

Pitching moment H =b ˆ 0:083


curve

slope

versus

stagger,

Unsteady panel technique

In order to obtain an initial understanding of aircraft behaviour under transient conditions, i.e. during water bomb release, unsteady panel methods may be used to compute the change in aerodynamic forces, moments and downwash [28]. These forces and moments are functions of

Fig. 18

Downwash curve slope versus tail arm, H =b ˆ 0:083, x=c ˆ 0 G03200 # IMechE 2001


Fig. 19

Downwash curve lH =c ˆ 3:75

Fig. 20

Downwash curve slope versus stagger, H=b ˆ 0, lH =c ˆ 3:75

slope

versus

gap,

x=c ˆ 0,

various flight parameters, i.e. angle of attack, pitch angle, trim parameters and time history. It would be difficult to present such forces in a graphical form as multivariable functions of all independent parameters, so these functions are shown as functions of time (only) under the assumption that other variables (e.g. angle of attack, pitch rate, flight speed, etc.) are fixed and equal to that of the steady state flight condition. Figure 21 shows the lifting force and pitching moment coefficients for the wing and body configuration (excluding horizontal tail) over a time interval equal to 1 s, that is the time for the contents of the hopper to be released. During this time the water column emerging from the aircraft extends to about 5 m in length.

139

assumption that the water bomb is dropped in a period of 1 s and that the aircraft is not controlled during the following 20 s (i.e. elevator deflection and the power unit thrust are the same as before water release). Unbalanced pitching moment can rapidly change aircraft pitch angle and angle of attack. In some cases this phenomenon may lead to stall. The aircraft response depends on two opposing phenomena: (a) after water release the aircraft goes up, so there is an additional downward component of velocity which decreases the angle of attack; (b) pitching moment due to the empty hopper acts either nose up (under the assumption that the hopper centre gravity position is placed forward of the mass centre), which means that the angle of attack increases, or nose down (under the assumption that the hopper centre gravity position is behind of the mass centre), which means that angle of attack decreases. Theoretically the angle of attack can increase or decrease, depending on the relationship between these two, abovementioned effects. In practice the first effect (i.e. decreasing angle of attack due to vertical motion of aircraft) is greater than the second effect corresponding to change in pitching moment. For illustrative purposes, calculations have been carried out on a concept aircraft having a biplane configuration. Apart from the hopper system with its release mechanism and strengthened undercarriage, all remaining structural components have been taken from an existing PZL agricultural aircraft, the PZL-106 ‘Kruk’. The fire-fighting aircraft has been given the designation PZL240 ‘Pelikan’. To compute a dynamic response for this aircraft, a set of dynamic equations of motion has been written in the body axis system. The origin of the axis system coincides with the mean quarter-chord point A: the AxA axis is directed forward of the aircraft along the MAC and the AzA axis is perpendicular to AxA and is directed downward. Equations of motion, together with kinematic relations, have the following form: _ ‡ m(xc ¡ xA )Q2 ˆ X ¡ mg sin õ m( U_ ‡ QW ) ¡ mzc Q _ ‡ mzc Q2 ˆ Z ‡ mg cos õ _ ‡ QU ) ‡ m(xc ¡ xA ) Q m( W _ ‡ m(xc ¡ xA ) W _ ¡ m(xc ¡ xA )UQ ¡ mzc U_ ¡ mzc QW J yQ ˆ M ‡ mgzc sin õ ‡ mg(xc ¡ xA ) cos õ x_ 0 ˆ U cos õ ‡ W sin õ z_ 0 ˆ ¡U sin õ ‡ W cos õ (24)

7.3

Flight dynamics investigation

Initial work has shown that aircraft dynamic response is strongly influenced by the change of total mass of the aircraft and the position of its centre of gravity. For this fire-fighting aircraft, the change of weight can be as high as 60 per cent of the total all-up weight. All further dynamic simulations have been computed under the G03200 # IMechE 2001

where õ denotes pitch angle, Q the pitch rate, U, W are the velocity components along x, z axes, xC , zC the coordinates of mass centre in the design system of axes (fixed to the mass centre, axis xC directed back of the aircraft parallel to MAC, axis zC directed up perpendicularly to MAC), X, Z, M the lift, drag and pitching moment and x0 , z0 the Proc Instn Mech Engrs Vol 215 Part G

140


Fig. 21

Fig. 22

Lift and pitching moment coefficient versus time (á ˆ 58, wing ‡ body only)

Influence of the hopper centre of gravity position on transient angle of attack after water release

coordinates of aircraft position in the ground fixed axis system (x0 axis is parallel to the ground, directed along the aircraft speed; z0 is vertical downwards). Figures 22 to 25 show results of the simulation determined from equations (24). Figure 22 shows the Proc Instn Mech Engrs Vol 215 Part G

variation of angle of attack after water release in four cases, corresponding to four positions of the centre of gravity, i.e. 15, 23, 27 and 32 per cent, measured with respect to MAC nose. From this figure it can be concluded that the smoothest transient response occurs when the aircraft G03200 # IMechE 2001


Fig. 23

Fig. 24

Influence of the hopper centre of gravity position on transient pitch rate after water release

Influence of the hopper centre of gravity position on transient speed after water release

centre of gravity is close to the centre of gravity of the hopper (i.e. 23 and 27 per cent). Moving the hopper centre of gravity back to the 32 per cent position changes trim conditions (á decreases from 7.38 to 78). After water is released from the hopper, the angle of attack decreases by G03200 # IMechE 2001

141

about 7o , practically independently of hopper position. Pitch rate (Fig. 23) changes very rapidly, but its maximum value is moderate (¹ 0:15 rad=s) and it quickly returns to zero. Because it has a relatively high moment of inertia, J y, the aircraft is much more liable to displace vertically than Proc Instn Mech Engrs Vol 215 Part G

142


Fig. 25

Influence of the hopper centre of gravity position on aircraft trajectory after water release

to rotate about the C y axis. This aircraft is very stable in pitch which is advantageous for fire-fighting manoeuvres. On release of the load, the aircraft speed surges and then decreases. The surge is most severe when the hopper is located well forward. This effect occurs because, following release from the hopper in the forward position, there is a large shift of the centre of gravity to the rear. The equation mH‡E xH‡E ˆ mH xH ‡ mE xE

(25)

clearly indicates the effect. The symbols xH xE denote mass centre positions, m denotes the mass value and suffixes E,H refer to the empty aircraft and to the hopper respectively. The further back the empty aircraft mass centre position is located, the smaller the stability margin and the less stable the aircraft is after water release. Figure 25 shows flight altitude after water release. It is clearly seen that altitude increases, independently of the hopper position. On the basis of the above analysis (Figs 22 to 25) it can be concluded that hopper position does not strongly influence the aircraft’s flight dynamics (under the assumption that the hopper is placed in the vicinity of the aircraft’s centre of mass). Shifting the hopper away from the aircraft’s centre of mass reduces longitudinal stability and increases the time response after water release. 7.4

Coherence of the liquid column

To find the conditions for water column coherence, giving a bomb effect, it is necessary to develop a CFD model or to perform a number of experimental tests. Because both the Proc Instn Mech Engrs Vol 215 Part G

experiments and the CFD analysis are far beyond the scope of this paper, experience gained by PZL and designers of other fire-fighting aircraft has been used to design the hopper. Its dimensions were assumed to be 1:2 m 3 2:0 m 3 2:5 m. A relatively high aspect ratio (2.08) ensures a high hydrostatic pressure at its base and allows water discharge in a short time. However, to optimize the water release a suitable CFD model needs to be developed [20].

8

SPECIFIC DESIGN OF THE PROPOSED AIRCRAFT

The concept aircraft, the PZL-240 ‘Pelikan’ referred to in Section 7.3, has been configured as a specialized forestfire-fighting aircraft although it has become apparent that other roles are possible. Consideration has been given to some of the design details required. The basic requirement has been for an aircraft with a capability of delivering 4 t of fire-extinguishing agent, normally water with additives, in a release time of 1 s. Other less demanding release modes may be employed depending on requirements. For example, division of the hopper into four compartments separated by vertical baffles to prevent sloshing has made it possible for separate release of the contents of each compartment. A novel design philosophy not available in aircraft of similar capacity used by forestry services allows the aircraft to be used for agricultural tasks, such as pest control, fertilizing, land reclamation and aerial application of specialist treatments. This may be achieved by modification of the outlet G03200 # IMechE 2001


system to accommodate granulates or liquid sprays. For example, the design of the base of the hopper and the gate allows for a controlled output of viscous liquids for land reclamation. A change of outlet system allows the aircraft to spread granulates, at a range of dosage rates, including very high rates, or to spray liquids. The aircraft may operate in a number of different firefighting roles. For example, it may operate for fire detection and location, carrying the maximum load of 4 t of fire-extinguishing agent having a flight duration of 1.5 h. Alternatively, with a payload of 3 t operating in the surveillance role it has a duration of 5 h. The aircraft has been fitted with a control system which allows additive liquid to be introduced to the hopper in flight. It may be a frothing or wetting agent as required, supplied from a separate tank or from the original commercial packing. The loading of the extinguishing agent may be performed by outboard systems, by use of the aircraft’s own pump and piping system or in flight from a suitable water surface using a retractable probe. In this design concept, the hopper with reinforced structure and landing gear make up the central and main assembly of the aircraft structure. All other assemblies have been attached to this. These include the wings, the power plant with engine mounting, the rear fuselage with pilot’s cockpit and tail unit as well as the landing gear including nose wheel. Composites have been suggested for hopper construction. Four loading ports on the upper hopper surface allow rapid loading of fire-fighting agent on the ground. The lower part of the hopper contains four gates, which may be opened in any configuration and to any degree, as required. Jacks controlling the gate opening have been located in the cockpit. A biplane configuration has been selected (Fig. 26). The pilot’s cabin, rear part of fuselage with control surfaces and the wings originate from the PZL-106 ‘Kruk’ aircraft (Fig. 27). The hermetically sealed cabin has high strength and the design has been proven over many years of operation. The cabin may take a load factor of 40 g in the flight direction. The operation of the ‘Kruk’ aircraft has demonstrated that this has contributed to pilot survival, even during very serious plane crashes. The air-tight cabin has been designed to provide the pilot with clean air delivered to the cabin through an exchangeable chemical filter cartridge. The cabin has been configured ergonomically giving good ground visibility while allowing the pilot to operate all fire-fighting systems and to maintain ‘handson’ operation of the flying controls. The high maximum lift coefficient of the wings makes it possible for the aircraft to perform tight turns—an important characteristic for fire-fighting manoeuvres. They have very mild stall characteristics and have been well proven in operation. All these components have been in production and have proven reliability. The landing gear has been configured with a highamplitude shock absorber and large-diameter, low-pressure tyres enabling the aircraft to operate from uneven or G03200 # IMechE 2001

Fig. 26

143

Configuration of the Pelikan PZL-240

unprepared landing fields. In order to ensure quick ground handling the nose wheel has been made steerable. These features, together with high static thrust and the reversing capabilities of the propeller, shorten the take-off and landing distances, thus allowing it to operate from landing fields localized within forest areas. A Pratt and Whitney type PW-120 turboprop engine has been selected as the power plant. It has a take-off rating of 2000 hp. A constant speed Hamilton STD propeller with a diameter of 3.8 m has been selected. This minimizes the landing run and enables operation from small landing strips. The engine air inlet channel has been fitted with a filtration system, which includes an anti-dust inertial filter. A container in the form of an oblong, tapered box has been provided under the rear part of the fuselage. It has been fitted with appropriate fastenings and clamps, enabling it to be used for transporting specialized auxiliary equipment, e.g. refilling hoses (Fig. 28). Significant dimensions and performance data have been computed for the aircraft at sea level, I SA [29, 30]: Proc Instn Mech Engrs Vol 215 Part G

144


Fig. 27

PZL-240 and its main elements. Assemblies originating from PZL-106 aircraft have ‘hatched’ surfaces

Span Wing chord Length Wheel track Wheel base Lifting area Tailplane area Vertical tail unit area Tailplane volume ratio Hopper for extinguishing agent Hopper for additives Capacity of fuel tanks Empty mass Maximum take-off mass (restricted category) Maximum speed Operational speed Stalling speed Rate of climb Take-off run (grass) Landing run with reverse Range Endurance (with 4000 kg of agent) Patrol endurance (with 3000 kg of agent)

9

15.5 m 2m 11 m 3.6 m 3.2 m 65 m2 10 m2 4.5 m2 0.57 6000 l 100 l 1150 l 3000 kg 7500 kg 230 km/h 170 km/h 80 km/h 4.5 m/s 270 m 250 m 1000 km 1.5 h 5h

water which may be maintained in the hopper for 1 h when consuming 1 l of fuel. The high lift wing originating from the PZL-104 allows the aircraft to perform tight turns. The full turn time (t) is given by 2ðV t ˆ p g n2 ¡ 1

(26)

where n is the load coefficient. The higher the maximum lift coefficient CLmax , the higher the load coefficient n and the shorter the turn time t. Calculations indicate that the Pelikan PZL-240 has a shortest full turn time of 30 s. By comparison, the Ayres Thrush has a more efficient wing

COMPARISON WITH COMPETITORS

Some performance characteristics of other fire-fighting aircraft have been analysed and included in Table 3. The last two columns provide parameters which identify operational cost: (a) mass of agent divided by the fuel consumption and (b) cost. Both represent the economical efficiency of the fire control aircraft. Mass of agent to the fuel consumption expresses the number of kilograms of Proc Instn Mech Engrs Vol 215 Part G

Fig. 28

Fire-extinguishing system layout: 1, four-chamber hopper of 6 m 3 volume; 2, hopper for additives; 3, electric pump; 4, water gauge; 5, self-refilling unit; 6, valve; 7, refilling hose; 8, probe for in-flight refilling in retracted position; 9, probe for in-flight refilling in refilling position G03200 # IMechE 2001


Table 3

Comparison between selected fire-fighting aircraft

Amount of agent (kg)

Engine and rating (h.p.)

Take-off run (m)

Turbine PT6 A34AG, 750 Turbine M601D, 750

210

1300=160 ˆ 8:1

500

240

1500=180 ˆ 8:3

400

Piston Asz-62, 1000 Piston Asz-62, 1000 Turbine PW-123, 2 3 2300 Turbine PW-120, 2000

245 170 450

1800=210 ˆ 8:6 1200=240 ˆ 5:0 6000=600 ˆ 10:0

500 No longer in production 4.500

270

4000=280 ˆ 14:3

1.300

Aircraft

1

3 200

1 300

3 500

1 500

3 4 5

Schweizer AG-CAT Super-B-Turbine PZL-106-BT Turbo‘Kruk’ M-18 Dromader AN-2 Canadair CL-215T

4 700 5 500 21 000

1 800 1 200 6 000

6

PZL-240 Pelikan

7 500

4 000

and a higher maximum speed but has the shortest full turn time of 60 s.

10

CONCLUSIONS

A number of features relating to the design and operation of a fire-fighting aircraft have been addressed. Areas requiring further research have been identified. The latter include an investigation of the interaction between the air flowing around the aircraft and the water column as it emerges from and leaves the hopper and the hopper design optimization which has to be investigated to ensure that it is capable of releasing a coherent column of water and that it reaches the ground without excessive dissipation. It has been found that shifting the hopper away from the aircraft centre of mass worsens longitudinal stability and increases time response after water release. Design areas include the hopper as the main assembly of the aircraft and its corresponding reinforcements. Special attention has to be given to the hopper control systems and nozzles. Comparison between selected fire-fighting aircraft shows that Pelikan PZL-240 has significant advantages over existing fire-fighting aircraft. Three different codes of panel methods have been used, namely (a) classical steady panel method for thick wings and bodies, (b) the unsteady version applicable for changeable configurations and (c) the VLM used for steady flow analysis around thin wings. All these methods have been used to compute pressure distributions and aerodynamic characteristics. It has been found that increasing the gap decreases the induced drag monotonically (under the assumption that lifting force coefficient is constant). Effects of stagger have been analysed and it has been found that positive stagger increases the induced drag and simultaneously shifts the neutral point of static stability forward (which is undesirable). The important part of this paper has been to compare some computational methods using the equivalent monoplane and giving selected algorithms for computing the equivalent wing. After simulations it has been suggested that the best method for G03200 # IMechE 2001

Mass of agent (kg) to the fuel consumption (kg h/l)

Maximum take-off mass (kg)

Number

2

145

Cost (approximate) (31000 US$)

designing the equivalent wing is that based on pressure distribution by means of the panel method. Although the method of analysis (i.e. panel methods) is dated, and the biplane configuration is regarded by many as obsolete, the application of these methods to biplanes is, as far as the authors are aware, original and innovatory. It seems that the biplane configuration provides a good solution for long endurance patrol with high capacity hopper and firefighting flights, mainly because of its compact design arrangement and short time, small turn radius.

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14 Dupont, J. Premier vol de’lOutrider. Air Cosmos-Aviat. Int., 28 March 1997, (1606), 20. 15 Wood, D. Jane’s World Aircraft Recognition Handbook, 4th edition, 1989 (Butler and Tanner, London). 16 Bednarek, K., Danilecki, S., Frydrychewicz, A., Goraj, Z., et al. Conceptual design of high altitude reconnaissance unmanned aerial vehicle—HARVE. Warsaw University of Technology, Institute of Aeronautics and Applied Mechanics, Warsaw, 1995 (unpublished). 17 Nenadovitch, M. Recherches sur les Cellules Biplane Rigides d’Envergure Infine, 1936 (Publications Scientifiques et Techniques du Ministere de L’Air, Institut Aerotechnique de SaintCyr, Paris). 18 Norton, F. H. Effect of staggering a biplane. NACA TN 70, 1972. 19 Goraj, Z., Frydrychewicz, A. and Winiecki, J. Design concept of a high altitude long endurance unmanned aerial vehicle. Aircr. Des. Int. J., 1999, 2(1), 19 –44. 20 Anderson Jr, J. D. Computational Fluid Dynamics, 1995 (McGraw-Hill, New York). 21 Jones, B. Elements of Practical Aerodynamics, 1939 (John Wiley, New York). 22 Glauert, H. The Elements of Aerofoil and Airscrew Theory, 2nd edition, 1947 (Cambridge University Press, Cambridge). 23 Hess, J. L. and Smith, A. M. O. Calculation of potential flow


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about arbitrary bodies. In Progress in Aeronautical Sciences, Vol. 8 (Ed. D. Ku¨chemann), 1967, pp. 1–138 (Pergamon, Oxford). Katz, J. and Plotkin, A. Low-Speed Aerodynamics—From Wing Theory to Panel Methods, 1991 (McGraw-Hill, New York). Olson, E. C. and Selberg, B. P. Experimental determination of improved aerodynamic characteristics utilising biplane wing configurations. J. Aircr., 1976, 13(4), 256–261. Bertin, J. J. and Smith, M. L. Aerodynamics for Engineers, 1989 (Prentice-Hall, London). ESDU data sheets, sub-series Aerodynamics, No. 66031, October 1996 (Engineering Sciences Data Unit, London). Blaszczyk, P., Grabowski-Goetzendorf, T., Goraj, Z. and Sznajder, J. Modelling of unsteady flow about a fire fighting aircraft dropping the water bomb. In 36th Aerospace Meeting and Exhibit, Reno, Nevada, 1998, AIAA paper 98-0763. Frydrychewicz, A. PZL-240 ‘PELIKAN’—proposal of design. Report, PZL-Warszawa-Okecie, Warsaw, 1996 (unpublished). Frydrychewicz, A., Goraj, Z., Ransom, E. C. P. and Wagstaff, P. Firefighting aircraft of medium capacity—from concept to design layout. In Proceedings of the Third Seminar on Recent Research and Design Progress in Aeronautical Engineering and Its Influence on Education, Institute of Aeronautics and Applied Mechanics Bulletin 9, 1999, pp. 181–194.

G03200 # IMechE 2001