Aeronautical Engineer’s Data Book - part 5 ppt

102 Aeronautical Engineer’s Data Book the principle of moments the following expres- sion can be derived for k CP : x AC C M AC k CP = ᎏ – ᎏᎏᎏ c C L cos ␣ + C D sin ␣ Assuming that cos ␣ Х1 and C D sin ␣ Х 0 gives: x AC C M AC k CP Х ᎏ – ᎏ c C L 6.6 Supersonic conditions As an aircraft is accelerated to approach super- sonic speed the equations of motion which describe the flow change in character. In order to predict the behaviour of airfoil sections in upper subsonic and supersonic regions, compressible flow equations are required. 6.6.1 Basic definitions M Mach number M ∞ Free stream Mach number M c Critical Mach number, i.e. the value of which results in flow of M ∞ = 1 at some location on the airfoil surface. Figure 6.6 shows approximate forms of the pressure distribution on a two-dimensional airfoil around the critical region. Owing to the complex non-linear form of the equations of motion which describe high speed flow, two popular simplifica- tions are used: the small perturbation approxima- tion and the so-called exact approximation. 6.6.2 Supersonic effects on drag In the supersonic region, induced drag (due to lift) increases in relation to the parameter ͙ M 2 – 1 ෆ function of the plan form geometry of the wing. 6.6.3 Supersonic effects on aerodynamic centre Figure 6.7 shows the location of wing aerody- namic centre for several values of tip chord/root chord ratio ( ␥ ). These are empirically based results which can be used as a ‘rule of thumb’. 103 0 Basic aerodynamics M1 (local) M ∞ > M crit M ∞ > M crit –1.2 –0.8 –0.4 –C –C p p 0.4 0.8 1.2 0 0.2 0.4 0.6 0.8 1.0 0 0.2 0.4 0.6 0.8 1.0 x/c x/c M ∞ >> M crit Supersonic regions –0.8 –0.4 –1.2 –C p 0 0.4 0.8 1.2 0 0.2 0.4 0.6 0.8 1.0 x/c Fig. 6.6 Variation of pressure deterioration (2-D airfoil) 6.7 Wing loading: semi-ellipse assumption The simplest general loading condition assump- tion for symmetric flight is that of the semi- ellipse. The equivalent equations for lift, downwash and induced drag become: For lift: VK 0 πs L = ␳ ᎏ 2 1 replacing L by C L / 2 ␳ V 2 S gives: C L VS K 0 = ᎏ πs 104 Aeronautical Engineer’s Data Book 2.0 1.8 1.6 1.4 X a.c. 1.2 C r 1.0 0.8 0.6 0.4 0.2 1.4 1.2 1.0 X a.c. 0.8 C r 0.6 0.4 0.2 0 1.2 1.0 X a.c. 0.8 C r 0.6 0.4 0.2 0 0 1 0 1 0 tanΛ LE β β tanΛ LE λ = C t / C R = 1.0 C t / C R = 0.5 C t / C R = 0.25 AR tanΛ LE AR tanΛ LE 6 5 4 3 2 1 6 5 4 3 2 1 AR tanΛ LE 6 5 4 3 2 1 Subsonic Supersonic Subsonic Supersonic Subsonic Supersonic Unswept T.E. Unswept T.E. Sonic T.E. Sonic T.E. Taper ratio β tanΛ LE tanΛ LE β Fig. 6.7 Wing aerodynamic centre location: subsonic/ supersonic flight. Originally published in The AIAA Aerospace Engineers Design Guide, 4th Edition. Copyright © 1998 by The American Institute of Aeronautics and Astronautics Inc. Reprinted with permission. ᎏ ᎏ 105 Basic aerodynamics For downwash velocity (w): w = K 0 ᎏ 4S , i.e. it is constant along the span. For induced drag (vortex): C L 2 πAR D D = V where aspect ratio (AR) = 2 span 4s 2 = area ᎏ S Hence, C D V falls (theoretically) to zero as aspect ratio increases. At zero lift in symmetric flight, C D = 0. V Section 7 Principles of flight dynamics 7.1 Flight dynamics – conceptual breakdown Flight dynamics is a multi-disciplinary subject consisting of a framework of fundamental mathematical and physical relationships. Figure 7.1 shows a conceptual breakdown of the subject relationships. A central tenet of the framework are the equations of motion, which provide a mathematical description of the physical response of an aircraft to its controls. 7.2 Axes notation Motions can only be properly described in relation to a chosen system of axes. Two of the most common systems are earth axes and aircraft body axes. The equations of motion and handling properties Aerodynamic characteristics Common aerodynamic parameters Stability and control derivatives Stability and control parameters Aircraft flying of the airframe Fig. 7.1 Flight dynamics – the conceptual breakdown 107 Principles of flight dynamics Conventional earth axes are used as a reference frame for ‘short-term’ aircraft motion. S N y 0 z 0 o 0 x 0 y E z E x E o E • The horizontal plane o E , x E , y E , lies parallel to the plane o 0 , x 0 , y 0 , on the earth’s surface. • The axis o E , z E , points vertically downwards. Fig. 7.2 Conventional earth axes 7.2.1 Earth axes Aircraft motion is measured with reference to a fixed earth framework (see Figure 7.2). The system assumes that the earth is flat, an assump- tion which is adequate for short distance flights. 7.2.2 Aircraft body axes Aircraft motion is measured with reference to an orthogonal axes system (Ox b , y b , z b ) fixed on the aircraft, i.e. the axes move as the aircraft moves (see Figure 7.3). 7.2.3 Wind or ‘stability’ axes This is similar to section 7.2.2 in that the axes system is fixed in the aircraft, but with the Ox- axis orientated parallel to the velocity vector V 0 (see Figure 7.3). 7.2.4 Motion variables The important motion and ‘perturbation’ variables are force, moment, linear velocity, angular velocity and attitude. Figure 7.4 and Table 7.1 show the common notation used. 7.2.5 Axes transformation It is possible to connect between axes refer- ences: e.g. if Ox 0 , y 0 , z 0 are wind axes and components in body axes and ␾ , ␪ , ␺ are the angles with respect to each other in roll, pitch and yaw, it can be shown that for linear quanti- ties in matrix format:  = D  Ox 0 Oy 0 Oz 0 Ox 3 Oy 3 Oz 3 108 Aeronautical Engineer’s Data Book x b x w z b z w y b, .y w V 0 Conventional body axis system. O x b is parallel to the ‘fuselage horizontal’ datum O z b is ‘vertically downwards’ O Conventional wind (or‘stability’) axis system: O x w is parallel to the velocity vector V o Roll L,p, φ Pitch M,q, θ Yaw N,r, ψ X,U e ,U,u Z,W e ,W,w Y ,V e ,V,v Fig. 7.3 Aircraft body axes Fig. 7.4 Motion variables: common notation   и  и и и и и и и и и ␪ ␺ 109 Principles of flight dynamics Table 7.1 Motion and perturbation notation Perturbations Aircraft axis Ox Force X Moment L Linear velocity U Angular velocity p Attitude ␾ Oy Oz Y Z M N V W q r Motions X Axial ‘drag’ force Y Side force Z Normal ‘lift’ force L Rolling moment M Pitching moment N Yawing moment p Roll rate q Pitch rate r Yaw rate U Axial velocity V Lateral velocity W Normal velocity Where the direction cosine matrix D is given by: cos ␪ cos ␺ cos ␪ cos ␺ – sin ␪ sin ␾ sin ␪ cos ␺ sin ␾ sin ␪ sin ␺ sin ␾ cos ␪ D =  – cos ␾ sin ␺ + cos ␾ sin ␺ cos ␾ sin ␪ cos ␺ cos ␾ sin ␪ cos ␺ cos ␾ cos ␪ + sin ␾ sin ␺ – sin ␾ cos ␺ Angular velocity transformations can be expressed as: p 1 0 –sin ␪ ␾ q =  0 cos ␾ sin ␾ cos ␪  ␪ r 0 –sin ␾ cos ␾ cos ␪ ␺ where p, q, r are angular body rates: Roll rate p = ␾ – ␺ sin ␪ ␾ , ␪ , ␺  where и Pitch rate q = ␪ cos ␾ are attitude + ␺ sin ␾ cos ␪ rates with и respect to Yaw rate r = ␺ cos ␾ cos ␪ datum axes – ␪ sin ␾ и и и   110 Aeronautical Engineer’s Data Book Inverting gives: ␾ 1 sin ␾ tan ␪ cos ␾ tan ␪ p ␪ =  0 cos ␾ –sin ␾ q    ␺ 0 sin ␾ sec ␪ cos ␾ sec ␪ r 7.3 The generalized force equations The equations of motions for a rigid aircraft are derived from Newton’s second law (F = ma) expressed for six degrees of freedom. 7.3.1 Inertial acceleration components To apply F = ma, it is first necessary to define acceleration components with respect to earth (‘inertial’) axes. The equations are: 1 a x = U – rV + qW – x(q 2 + r 2 ) + y(pq – r) + z(pr + q) 1 a y = V – pW + rU + x(pq + r) – y(p 2 + r 2 ) + x(qr – p) 1 a z = W – qU + pV = x(pr – q) + y(qr + p) – z(p 2 + q 2 ) 1 1 where: a 1 x , a y , a are vertical acceleration z components of a point p(x, y, z) in the rigid aircraft. U, V, W are components of velocity along the axes Ox, Oy, Oz. p, q, r are components of angular velocity. 7.3.2 Generalized force equations The generalized force equations of a rigid body (describing the motion of its centre of gravity) are: m(U – rV + qW) = X where m is m(V – pW + rU) = Y the total mass m(W – qU + pV) = Z of the body 7.4 The generalized moment equations A consideration of moments of forces acting at a point p(x, y, z) in a rigid body can be expressed as follows: и и и 111 Principles of flight dynamics Moments of inertia I x = ∑ ␦ m(y 2 + z 2 ) Moment of inertia about Ox axis I I = ∑ ␦ m(x 2 + z 2 ) Moment of inertia about Oy axis z = ∑ ␦ m(x 2 + y 2 ) Moment of inertia about Oz axis I y I = ∑ ␦ m xy Product of inertia about Ox and Oy axes xz = ∑ ␦ m xz Product of inertia about Ox and Oz axes I = ∑ ␦ m yz Product of inertia about xy yz Oy and Oz axes The simplified moment equations become I x p и – (I y – I z ) qr – I xz (pq + r и ) = L 2 I y q и – (I x – I z ) pr – I xz (p – r 2 ) = M  I z r и – (I x – I y ) pq – I xz (qr + p и ) = N 7.5 Non-linear equations of motion The generalized motion of an aircraft can be expressed by the following set of non-linear equations of motion: m(U – rV + qW) = X a + X g + X c + X + X dp m(V – pW + rU) = Y a + Y + Y c + Y p + Y dg m(W – qU + pV) = Z a + Z g + Z c + Z + Z dp I x p и – (I y – I x ) qr – I xz (pq + r и ) = L a + L g + L c + L p + L d  2 I M I y q и + (I x – I z ) pr + I xz (p – r 2 ) = M a + M g + c + M p + M d z r и – (I x – I y ) pq + I xz (qr – p и ) = N a + N + g N c + N p + N d 7.6 The linearized equations of motion In order to use them for practical analysis, the equations of motions are expressed in their linearized form by using the assumption that all perturbations of an aircraft are small, and about the ‘steady trim’ condition. Hence the equations become: [...]... 0.414 0.6 45 0 .54 5 Dimensions Length (m) 1.62 Fan diameter (m) 1.272 Basic eng 13 85 weight (lb) 2 .51 4 1.219 13 25 2.616 1.9 45 5700 2 1+5LP +1CF 2HP 3LP Layout Number of shafts 2 Compressor various Turbine 2HP 2LP 35 000 0.83 0 .56 2 0 .54 5 35 000 0.8 51 85 0.2 ISA+10 0 .57 4 35 000 35 000 0.8 0.8 57 25 0.174 ISA+10 0 .57 4 35 000 0.8 35 000 0.8 35 000 0.83 35 000 0.82 1 150 0 0.162 ISA+10 0 .56 5 35 000 0.8 255 0 0.184... B777­ 200/300 19 95 67 50 0 30 90 000 30 32.4 1926 0.33 39.3 3037 18 000 Pratt & Witney Rolls-Royce V 252 2 A5 V 253 3 A5 PW4 052 PW4 056 PW4168 PW4084 MD90­ 10/30 A319 1993 22 000 30 5 24.9 738 0.34 A321200 B76 7-2 00 &200ER B74 7-4 00 A330 76 7-3 00ER 1994 33 000 30 4.6 33.4 848 0.37 1986 52 200 33.3 4. 85 27 .5 17 05 0. 351 1987 56 750 33.3 4. 85 29.7 17 05 0. 359 55 50 ISA+10 62 25 ISA+10 1993 68 000 30 5. 1 32 1934 ZMKB TRENT... 0.184 0.69 35 000 0. 85 11813 0.1 95 ISA+10 0 .57 36 089 0. 75 3307 0.196 2.616 1.2 45 1670 4.343 2.794 10 726 5. 181 3.404 16 644 3.204 1.681 52 52 3.204 1.681 52 30 3.879 2.477 9400 3.879 2.477 9400 4.143 2 .53 5 14 350 4.869 2.8 45 13 700 3.912 2.474 10 55 0 2 .59 1 .52 2 951 3.1 75 2.192 9670 1.373 3197 2 1+4LP 9HP 2 1F +14cHP 2 1+4LP 14HP 2 1+3LP 10HP 2 1+4LP 10HP 2 1+4LP 10HP 2 1+4LP 11HP 2 1+4LP 11HP 2 1+5LP 11HP... TAY 611 RB-21 152 4H D-436T1 B777 A330 F100.70 B74 7-4 00 Gulfst V B76 7-3 00 Tu-33 4-1 An 72,74 1994 84 000 30 6.41 34.2 255 0 19 95 71 100 30 4.89 36.84 1978 1988 13 850 30 3.04 15. 8 410 0.43 1989 60 600 30 4.3 33 16 05 0 .56 3 1996 16 8 65 30 4. 95 25. 2 15 386 ISA+10 3400 ISA +5 12 726 ISA+10 Cruise Altitude (ft) Mach number Thrust (lb) Thrust lapse rate Flat rating (°C) SFC (lb/hr/lb) 40 000 0.8 1310 35 000 0.8... Anotov An-70 transport have been designed with propfans 8.2 .5 Turboshafts Turboshaft engines are used predominantly for helicopters A typical example such as the Rolls-Royce Turbomeca RTM 32201 has a three-stage axial compressor direct-coupled to a two-stage compressor turbine, and a two-stage Table 8.2 Aircraft engines – basic data Company Allied Signal CFE Engine type/Model LF507 CFE738 CFM 56 CF34 5C2... BA14 6-3 00 Falcon Avro RJ 2000 A340 Canadair RJ In service date Thrust (lb) Flat rating (°C) Bypass ratio Pressure ratio Mass flow (lb/s) SFC (lb/hr/lb) 1991 7000 23 5. 6 13.8 256 0.406 1994 31 200 30 6.4 31 .5 10 65 0.32 1996 9220 Climb Max thrust (lb) Flat rating (°C) 1992 59 18 30 5. 3 23 240 0.369 CFMI 758 0 General Electric (GE) 21 0. 35 CF6 80E1A2 IAE (PW, RR, MTU, JAE) GE 90 85B A330 B777­ 200/300 19 95. .. 1+4LP 10HP 2 1+4LP 11HP 2 1+4LP 11HP 2 1+5LP 11HP 2 1+6LP 11HP 3 1LP 8IP 6HP 2 1+3LP 12HP 3 1LP 7IP 6HP 3 1+1L 6I 7HP 1HP 5LP 2HP 4LP 2HP 5LP 2HP 6LP 2HP 5LP 2HP 5LP 2HP 4LP 2HP 4LP 2HP 5LP 2HP 7LP 1HP 1IP 4LP 2HP 3LP 1HP 1IP 3LP 1HP 1IP 3LP 0.61 126 Aeronautical Engineer’s Data Book power turbine Drive is taken off the power turbine shaft, through a gearbox, to drive the main and tail rotor blades... performances of propellers having the same geometry: F = ␳n2d4pCF Q = ␳n2d5pCQ P = ␳n3d5pCp CF, CQ and CP are termed the thrust, torque, and power coefficients These are normally expressed in USCS units, i.e.: F Thrust coefficient CF = ᎏ ␳n2d4 Q Torque coefficient CQ = ᎏ ␳n2d5 Power coefficient CP P =ᎏ ␳n3d4 118 Aeronautical Engineer’s Data Book where d n Q F P ␳ = propeller diameter (ft) = speed in revs per... Turbine-driven propeller Gas thrust Propeller thrust Turboshaft Shaft power Output (e.g to drive helicopter rotor) Fig 8.3 Gas turbine engine types 120 Aeronautical Engineer’s Data Book 100 90 Engine efficiency (%) Turboprop 80 70 Turbofan 60 Turbojet 50 0 .5 0.6 0.7 0.8 0.9 Mach No (cruise) Fig 8.4 ‘Order of magnitude’ engine efficiencies the main types and Figure 8.4 an indication of engine efficiency... f ( P3 / P2 ) for α = 5 η0 = f (P3/P2) for α = 5 0.3 0.8 0.2 0.4 η0 = f (α) for P3/P2 = 10 1 0.1 Overall efficiency η0 Dimensionless specific thrust parameter λ Principles of propulsion 3 7 11 15 Compressor pressure ratio P3/P2 2 4 6 8 Cycle temperature ratio α = T4/t0 Fig 8 .5 Turbojet performance indicative design points Table 8.1 Turbojet performance parameter groupings Non-dimensional group Flight . Turbine-driven Turboshaft Fig. 8.3 Gas turbine engine types 120 Aeronautical Engineer’s Data Book Engine efficiency (%) 100 90 Turboprop 80 70 Turbofan 60 Turbojet 50 0 .5 0.6 0.7. tanΛ LE λ = C t / C R = 1.0 C t / C R = 0 .5 C t / C R = 0. 25 AR tanΛ LE AR tanΛ LE 6 5 4 3 2 1 6 5 4 3 2 1 AR tanΛ LE 6 5 4 3 2 1 Subsonic Supersonic Subsonic Supersonic. pitch and yaw, it can be shown that for linear quanti- ties in matrix format:  = D  Ox 0 Oy 0 Oz 0 Ox 3 Oy 3 Oz 3 108 Aeronautical Engineer’s Data Book x b x w z b z w y b, .y w V 0 Conventional