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 23012-23010
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 250-C20B
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		9-33% 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 750C-2
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 71-H-080	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] (B2-x20)

[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 two-dimensional 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