Physical Parameters of Existing Helicopters
The following information has been graciously provided by the manufacturer of each helicopter.
Weights (kg) 
Engines 

Empty 
4,420 
Type 
Turbomeca Makila lAl 
Maximum takeoff 
8,600 
Number 
2 
Fuel capacity 
1,620 
Maximum T. O. rating 
2,712 kW 
Maximum usable power 
2,133 kW 

Rotor Parameters 
Main Rotor 
Tail Rotor 

Radius (m) 
R 
7.79 
1.525 
Chord (m) 
c 
0.6 
0.2 
Solidity 
О 
0.098 
0.209 
No. of blades 
b 
4 
5 
Tip speed (m/sec) 
CLR 
217 
204 
Twist (deg) 
Ox 
12.06 
15.71 
Hinge offset ratio 
e/R 
0.037 
0.072 
Airfoil 
HAS 13112,13109,13106 
NACA 2301223010 

Collective range (deg) 
25 
41 

Longitudinal cyclic range (deg) 
30 
— 

Lateral cyclic range (deg) 
13.5 
— 

Polar moment of inertia (m2kg) 
J 
7,035 
13.1 
Weights (kg) 
Engines 

Empty Maximum takeoff Fuel capacity 
1900 4000 892 
Type Number Maximum T. O. rating Maximum usable power 
Turbomeca Arriel 1C 2 984 kW 899 kW 
Rotor Parameters 
Main Rotor 
Fenestron 

Radius (m) 
R 
5.965 
.45 
Chord (m) 
c 
.385 
0.0435 
Solidity 
a 
0.082 
— 
No. of blades 
b 
4 
13 
Tip speed (m/sec) 
ClR 
218 
227 
Twist (deg) 
0! 
10.2 
8° 
Hinge offset ratio 
e/R 
0.038 
0 
Airfoil 
OA 212, OA 209 OA 207 
NACA 63A312 NACA 63A309 

Collective range (deg) 
13.75 
67 

Longitudinal cyclic range (deg) 
26 
— 

Lateral cyclic range (deg) 
13 
— 

Polar moment of inertia (m2kg) 
J 
2,090 
0.35 
Weights (kg) 
Engines 

Empty Maximum takeoff Fuel capacity 
1,051 1,950 405 
Type Number Maximum T. O. rating Maximum usable power 
Turbomeca Arriel IB 1 478 kW 441 kW 
Rotor Parameters 
Main Rotor 
Tail Rotor 

Radius (m) 
R 
5.345 
0.930 
Chord (m) 
c 
0.3 
0.185 
Solidity 
a 
0.0536 
0.127 
No. of blades 
b 
3 
2 
Tip speed (m/sec) 
CLR 
213 
199 
Twist (deg) 
Єї 
12.275* 
0 
Hinge offset ratio 
e/R 
0.038 
0 
Airfoil 
NACA 0012 
NACA 0012 

Collective range (deg) 
15 
27 

Longitudinal cyclic range (deg) 
26 
— 

Lateral cyclic range (deg) 
12 
— 

Polar moment of inertia (m2 kg) 
J 
995 
1.06 
♦Linear twist from tip to rotor center. 
AGUSTA A1Q9
Weights (lb) 
Engines 

Empty 
3300 
Type 
Allison 250C20B 
Maximum takeoff 
5,730 
Number 
2 
Fuel capacity 
1,202 
Maximum T. O. rating 
840 
Maximum usable power 
740 

Rotor Parameters 
Main Rotor 
Tail Rotor 

Radius (ft) 
R 
18.04 
3.33 
Chord (ft) 
c 
1.10 
0.66 
Solidity 
a 
0.0775 
0.1256 
No. of blades 
h 
4 
2 
Tip speed (ft/sec) 
ClR 
727 
727 
Twist (deg) 
0! 
6 
0 
Hinge offset ratio 
e/R 
0.027 
— 
Airfoil 
NACA 23011.3/13006 
NACA 0016/0009 

Collective range (deg) 
0/+16 
7/+21 

Longitudinal cyclic range (deg) 
—10.5/+12.5 
– 

Lateral cyclic range (deg) 
±6.25 
— 

Tolar moment of inertia (slug ft2) 
J 
2,000 
2 
Weights (lb) 
Engines 

Empty 
6,598 
Type 
Lycoming 703 
Maximum takeoff 
10,000 
Number 
1 
Fuel capacity 
1,684 
Maximum T. O. rating 
1,800 
Maximum usable power 
1,290 

Rotor Parameters 
Main Rotor 
Tail Rotor 

Radius (ft) 
R 
22 
4.25 
Chord (ft) 
c 
2.25 
.96 
Solidity 
о 
.065 
.1435 
No. of blades 
b 
2 
2 
Tip speed (ft/sec) 
VtR 
746 
739 
Twist (deg) 
0, 
10.0 
0 
Hinge offset ratio 
e/R 
— 
— 
Airfoil 
933% SYM. Sect. (Special) 
NACA 0018 @ x = .25 

Tapering to 0008.27 @ Tip 

Collective range (deg) 
8.5 to 24 
19.85 to 10.15 

Longitudinal cyclic range (deg) 
±10 
— 

Lateral cyclic range (deg) 
±10 
— 

Polar moment of inertia (slug ft2) 
J 
2,770 
5.9 
Weights (lb) 
Engines 

Empty 
4,929 
Type 
Lycoming 750C2 
Maximum takeoff 
8,250 
Number 
2 
Fuel capacity 
1,275 
Maximum T. O. rating 
1,470 
Maximum usable power 
1,088 

Rotor Parameters 
Main Rotor 
Tail Rotor 

Radius (ft) 
R 
21 
3.45 
Chord (ft) 
c 
2.167 
.8 
Solidity 
о 
.0657 
.1476 
No. of blades 
b 
2 
2 
Tip speed (ft/sec) 
CLR 
765 
671 
Twist (deg) 
0! 
10.74 
0 
Hinge offset ratio 
e/R 
— 
— 
Airfoil 
FX 71H080 
BHT 10.9 FC 

Collective range (deg) 
11.7 ± 9 
1.7 to 25 

Longitudinal cyclic range (deg) 
13 to +15 
— 

Lateral cyclic range (deg) 
±9 
— 

Polar moment of inertia (slug ft2) 
J 
1,664 
1.0 
[1] (B2x20)
[2] Constant chord
• Linear twist
[3] 0O, collective pitch that is required to produce enough rotor thrust to balance the weight and to compensate for the inflow.
• 01} blade twist.
• ax and bx , tilt of the tip path plane with respect to the shaft that is required to produce enough moment about the center of gravity to balance or trim the helicopter.
• Ax and Bv cyclic pitch required to compensate for the unsymmetrical velocity pattern and to produce the amount of als and bXs required to trim.
[4] = Іо^Уо + 4h + 2уг + 4уз + 2ул + 4уъ + 2Уб + 4уі + 2у% + 4y* + Уі
where in this casсу is the calculated value of dCT/o/dr/R. The root and tip losses can be handled as trapezoidal corrections to the total integral:
A method for making a rough estimate of the vertical drag penalty in hover was given in Chapter 1. This method will now be refined in order to raise the confidence level in the hover performance calculations. The method consists of the following steps:
• Divide the plan view of the airframe into segments.
• Estimate the drag coefficient of each segment as a function of its shape.
• Determine the distribution of dynamic pressure in the rotor wake.
• Sum the effects of each segment.
[6] Calculate the rotor thrust as the sum of the weight and the vertical drag.
• Correct the rotor power at this thrust for the "ground effect” due to the airframe.
The questions most asked of the helicopter aerodynamicist concerning autorotation are:
• How much does the rotor speed decay following a power failure before the pilot can react?
• What is the minimum steady rate of descent?
[8] How far can the helicopter fly following the power failure while the pilot looks for a – suitable landing spot?
The interest of the helicopter aerodynamicist in airfoils is either for the analysis of an existing rotor or for the design of a new one. In the first case, he may acquire the data he needs either directly, from twodimensional wind tunnel tests or from whirl tower tests of a rotor with the specific airfoil, or indirectly, from test results of similar airfoils modified by empirical or theoretical means. For the design of a new helicopter, he may select one of the many airfoils already investigated or design an entirely new airfoil to incorporate characteristics he considers desirable. Since a blade with a good airfoil costs little or no more to build than one with a poor airfoil, there is strong motivation for improving airfoils even if the expected performance benefits are relatively small. A good airfoil for a rotor has:
• High maximum static and dynamic lift coefficients to allow flight at high tip speed ratios and/or at high maneuver load factors.
[10] A high drag divergence Mach number to allow flight at high advancing tip Mach numbers without prohibitive power losses or noise.
• Low drag at moderate lift coefficients and Mach numbers to minimize the power in normal flight conditions.
[11] +
Since both a0 and vl are direct functions of rotor thrust:
AAx = (n — 1) Л.
Maneuver ‘ 7 1 level flight
[12] Plot pitching moment versus angle of attack with stabilizer off and with stabilizer on at several incidence settings, as in Figure 8.16 (page 502).
• At each intersection the lift of the horizontal stabilizer is zero so the downwash angle must be equal to the geometric angle of attack of the stabilizer:
£fH CLp + Ifl
• Plot Єрн versus aF and determine ePa o and (deF/daF)H Span Efficiency Factor
Since the stabilizer lift distribution is usually strongly affected by flow irregularities coming back from the fuselage and rotor, it is not certain that
[13] far, this study has been limited to the longitudinal mode, but it is evident that the lateral mode in hover could have been treated in the same manner by using the moment of inertia in roll instead of pitch in the equations. For a simple analysis, it is again necessary to assume that the pilot holds heading by adjusting tail rotor