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

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