Determination of geometrical parameters for airplane units
Determination of wing parameters.
1.7.1.1 The wing surface area of the projected plane is.
S = m0.g/p0,
where g = 9.8 m/c2, p0 = 5809.5 N/m2
S = 190000x 9.8/ 5809.5
S =320.5 m2
1.7.1.2 Wing span
L = 
,
where the value of λ is taken from table. A,
L = (7.8)(320.5)
L = 50 m.
1.7.1.3 Wing chord
Root (on axis of airplane symmetry) b0 and tip bK wing chords are determined proceeding from the values S, η, L;
b0 = 
. 
,
bo =(320.5/50)(2x2.6/2.6+1)
b0 = 9.23 m.
bK = 
bk =9.23/2.6
bK = 3.55 m.
where value η = 4.2 taken from table A
1.7.1.4 Wing mean aerodynamic chord (MAC) is calculated by the formula
bA = 
.b0 
,
bA =(2x9.23/3)[ (2.6)+2.6+1/2.6(2.6+1)]
bA = 6.81 m
Coordinate of MAC along a wing span is determined by the relation
zA = 
. 
.
zA =(50/6)(2.6+2/2.6+1)
zA = 10.637 m
Coordinate of MAC nose along an axis 0x
xA = zA.tanχLE
where χLE (sweep angle of the wing’s leading edge) = 32.37O
tanχLE = tanχ + 
.
tanχLE = 0.577+[2.6-1/7.8(2.6+1)]
= 0.6339
where χ is the sweep angle of the wing’s quarter-chord line and is taken from table A.
thus xA = 10.637 x 0.6339
xA = 6.742 m.
Moreover the geometrical verification of the wing MAC has also been done.
Fig1.20 Geometrical determination of MAC.
Fuselage Parameters
The size and shape of subsonic aircraft are generally determined by the size and mass of the cargo. Arrangement of the cargo may vary depending size and range.
The main parameters needed to start constructing the fuselage section are as follows:
LF (Overall fuselage length) = λF .DF = 9 x 5.8 = 52.2 m;
DF (Diameter of the fuselage) = 5.8 m;
LN (Fuselage nose length) = λN .DF = 1.9 x 5.8 = 11.02;
LT (Fuselage rear length) = λT .DF = 3x 5.8 = 17.4 .
where λF, λN, λT are the aspect ratio of fuselage, nose part, rear part respectively.
Tail-unit parameters
The tail-unit of an airplane consists of the horizontal stabilizer and the vertical stabilizer. The geometrical calculations that involves for constructing the tail-unit are the same as that of the wing.
Distance from the airplane centre of mass up to the horizontal tail-unit centre of pressure LHS for normal configuration is determined according to recommendation which are based on statistical data.
Distance from the vertical tail-unit centre of pressure up to the airplane centre of mass LVS in zero approximation may be considered such which is equal to the distance of horizontal tail-unit, where LVS = LHS.
Horizontal tail unit
1.7.3.1.1 The surface area the horizontal stabilizer is:
SHS = ̅SHS . S
SHS = 0.23x 320.5
SHS = 73.715 m2;
where ̅SHS is the ratio of surface area of the horizontal stabilizer to the wing’s area.
1.7.3.1.2 The length of the Horizontal stabilizer is equal to
LHS = (λHS.SHS)-2
Lhs = (2.1)(76.92)
LHS = 12.7 m;
where the value of λHS is taken from table. A
1.7.3.1..3 Chord of HS
Root (on axis of airplane symmetry) b0 HS and tip bK HS chords are determined proceeding from the values SHS, ηHS, LHS;
b0 HS = (SHS/LHS).(2ηHS/ηHS+1)
bohs =(73.715/12.7)(2x2.4/2.4+1)
b0 HS = 8.55 m
bK HS =bK HS/ηHS
bkhs =8.55/2.4
bK HS = 3.56m
where value η = 2 and is takem from table A.
1.7.3.1.4 Mean aerodynamic chord (MAC) of HS is calculated by the formula
bA HS = { .b0 HS(ηHS2 + ηHS +1)}/ {ηHS (ηHS + 1)} ,
bAHS ={[2x(2.4)+2.4+1]/3}/2.4(2.4+1)
bA HS= 6.4 m
Coordinate of MAC along a Horizontal stabilizer is determined by the relation
zA HS =(LHS/6) .( ηHS+2)/( ηHS+1)
zAHS =(12.7/6) (2.4+2/2.4+1)
zA HS= 2.73 m
Coordinate of MAC nose along an axis 0x
xA HS = zA HS.tanχLE(HS)
where χLE (sweep angle of the horizontal stabilizer’s leading edge) is taken from table A,
tanχLE(HS) = tanχHS + ( ηHS-1)/{𝝀HS( ηHS+1)} 
.
tanχLE(HS) = (tan 340) + (2.4-1)/{ 2.1(2.4+1)}
tanχLE(HS) = 0.868
thus, xA HS = 2.73 x 0.868 = 2.37 m
Vertical tail unit
1.7.3.2.1 The surface area the vertical stabilizer is:
SVS = ̅SVS . S
SVS = 0.16 x 320.5 m
SVS = 51.28 m2;
where ̅SVS is the ratio of surface area of the vertical stabilizer to the wing’s area.
1.7.3.2.2 Length of Vertical stabilizer is equal to
LVS = (λVS.SVS)-2
Lvs = (1.4)(51.28)
LVS = 8.47 m;
where the value of λVS is taken from table. A’
1.7.3.2.3 Chords of VS
Root (on axis of airplane symmetry) b0 VS and tip bK VS chords are determined proceeding from the values SVS, ηVS, LVS;
b0 VS = (SVS/LVS).(2ηVS/ηVS+1)
bovs = (51.28/8.47) (2x2.6/2.1+1)
b0 VS = 8.742 m
bK VS =bK VS/ηVS
bkvs =8.742/2.6
bK VS = 3.362 m
where value ηVS = 1.5 and is taken from table A,
1.7.3.2.4 The mean aerodynamic chord (MAC) is calculated by the formula
bA VS = { .b0 VS(ηVS2 + ηVS +1)}/ {ηVS (ηVS + 1)} ,
bAVS =(2x8.742/3)[(2.6 +2.6+1)/(2.6)(2.6+1)]
bAVS = 6.451 m
Coordinate of MAC along a Vertical stabilizer is determined by the relation
zA VS = (LVS/3) .( ηVS+2)/( ηVS+1)
zAvs =(8.47/3).(2.6+2/2.6+1)zAVS = 3.609 m
Co-ordinate of MAC nose along an axis 0x
xA VS= zA VS.tanχLE(VS)
where χLE(VS) sweep angle of the Vertical stabilizer’s leading edge is taken from table A,
tanχLE(VS) = tanχVS + ( ηVS-1)/{𝝀VS( ηVS+1)}
tanχLE(VS) = (tan 390) {(2.6-1)/{1.4(2.6+1)}
tanχLE(VS) = 1.062
thus, xA VS = 1.86 x 1.062 = 1.96 m