Propeller Design Requirements Specification

Typical UAV propeller design requirements focus on maximising the UAV mission performance. The UAV performance requirements define the required propulsive power over the various phases of its mission. In this case satisfying the requirement to fly over a large range of altitudes requires a thorough knowledge of the relevant atmospheric characteristics from sea level up to 15 000 m.

Other constraints on the propeller design solution space are imposed by the geometry of the airframe. In this particular case a twin-boom configuration limits the maximum diameter of the propeller. There are additional requirements of high Mach number and low Reynolds number performance demanded of the blade sections. These require an initial investigation into the feasibility before the detail design can be initiated.

As the focus of this work is on the capabilities of the propeller design and analysis methods little effort is applied to the structural, aero-elastic or mass properties of the propeller blade except that required for safe testing of a scale wind tunnel model.

4.1.1 UAV Mission Description

The UAV mission description is summarised in Figure 4.1 into four flight phases. These are:

1. Assisted acceleration from standstill at sea level to an airspeed of 27 m/s. This is 1.3 times the predicted stall speed of 21 m/s – a typical speed for the climb phase. 2. Climb to a cruise altitude of 15 000 m at maximum climb rate to minimise flight time in commercial airspace. A minimum climb rate of 5m/s at sea level and 1 m/s at 15 000 m must be attained.

3. Cruise at 15 000 m altitude for maximum range on an “out-and-return” course 4. Descend, approach and land

Figure 4.1 HALE UAV mission profile

Note that the cruise flight portion of the mission is an “out and return” course at a constant altitude typically the flight velocity is varied to maximise the airframe efficiency with the reduction in weight over the mission.

The descent and landing phases use an insignificant portion of the fuel as the engine is typically at a low throttle setting for the descent due to the high airframe efficiency. The total fuel usage on the descent and landing phase was estimated at 4 % of total fuel. This amount would not vary significantly with propeller design due to the low power settings required and indeed the propeller would typically not be providing any thrust. As such the descent, approach and land phase was not modelled in the mission analysis except through the assumption of an additional 4% fuel fraction remaining over the normal reserves.

It is assumed that there are no wind or gust effects at any altitude in the mission profile to simplify the design requirements. Maximising range requires flying at airspeeds that are dependent on the local wind strength and direction that may not be optimal for the airframe. The additional power expended flying into the wind is typically more than that saved on the return leg due to the non-linear relationship between the airframe drag and airspeed.

4.1.2 UAV Performance

The HALE UAV airframe chosen for this study was based on a study previously carried out by the author (Monk, 1995) to determine the optimal configuration and sizing of an 18 m span HALE UAV. The work included parametric sizing of the engine and propeller through the application of various empirical methods. The overall propulsive efficiency was at that time estimated with no detailed analysis of the actual design of the propeller. As the original work was classified confidential so as to protect the original airframe’s performance requirements, an approximate set of specifications has been used in this work.

A maximum take off mass of 750 kg was assumed including 100 kg of usable fuel before reserves. This includes the 4% provision for the descent and landing phases. The chosen power plant was a Rotax 914 UL four stroke aircraft engine turbocharged to maintain a constant power of 59.6 kW (80 hp) from sea level to 15 000 m as implemented on the Perseus B UAV, Figure 4.2. (Goebel, 2008). The fuel consumption values of 0.419 litres per kW per hour were obtained from the Rotax 914 technical specifications (Rotax, 2009).

The UAV would be assisted in the take-off phase to a safe climb speed of approximately 27 m/s, 30% above its predicted stall speed. This is the minimum airspeed at which steady rate of climb is to be measured according to the Federal Aviation Regulation on En-route Climb or Descent (Federal Aviation Regulations, 1996). The flight speed range within which the propeller had to operate was thus defined to be from 27 m/s to maximum cruise speed.

The choice of gear ratio between the engine and propeller was not fixed at this point to allow more flexibility in determining the final design solution.

The specifications for the UAV are summarised in Table 4.1. The aspect ratio, maximum lift coefficient, Oswald efficiency factor and Cd0 were obtained from (Monk, 1995).

Table 4.1 UAV specifications

Airframe Mass 750 kg Available Power 59.600 kW Wingspan 18.0 M Wing Area 16.2 m2 Aspect Ratio 20.0 - Clmax 1.65 - Cd0 0.0200 - Oswald Efficiency 0.80 -

Specific Fuel Consumption 0.419 l/kW/hr

4.1.3 Propeller Design Constraints

When designing a propeller for a required mission performance for a given airframe, the airframe geometry and power plant choice also place constraints on the design of the propeller.

The UAV configuration consisted of a twin-boom, 18 m span wing and rear mounted engine driving a “pusher” propeller between the booms. The separation distance of three metres between the booms limited the diameter of the propeller to a practical limit of 2.8 m. The ground clearance requirements of the airframe at take-off attitude dictated a two- blade or folding propeller design to meet these requirements. A non-folding two-blade

propeller was chosen. The propeller would be restrained horizontally until the UAV had been towed into the air at which point in time the propeller would be allowed to rotate, a method used with success by Aurora Flight Sciences on their Perseus B UAV (Goebel, 2008).

Figure 4.3 The HALE UAV concept layout

The UAV was assumed to have a variable pitch hub capable of allowing rotation of the propeller through the large range of pitch angles required to match the true airspeed over the large change in altitudes.

Large propeller diameters typically result in relatively high blade root structural loads. The structural requirements of the propeller blades normally calculated as part of the design process, do not form a formal part of this study whose focus is on the investigation into the use of Larrabee’s methods although mention will be made later of the propeller’s structural feasibility.

The formal mass and performance trade-off studies typically carried out to evaluate the degradation in performance of the aircraft due to the additional mass of the pitch change mechanism compared with the lower mass but lower potential performance of the fixed pitch propeller similarly fall outside of the scope of this work. The structural stress predictions and vibration modal analysis are likewise not covered here.

The UAV’s mission profile requires flight from sea level into the tropopause. Inherent in the design process is an understanding and modelling of the atmospheric conditions from sea level up to these altitudes. The atmospheric conditions used in all the calculations were based on the standard atmosphere equations adopted by the United States Committee on Extension to the Standard Atmosphere (COESA, 1976).