Design and application of an Electric Tail Rotor Drive Control(ETRDC)for helicopters with performance tests

Yuwen ZHANG,Chen JIANG,Yunjie WANG,Fan SUN,Haowen WANG

School of Aerospace,Tsinghua University,Beijing 100084,China

KEYWORDS Electric tail rotor;Helicopter;More Electric Aircraft(MEA);Propulsion optimization;Tail rotor

Abstract With the development of electric helicopters’motor technology and the widespread use of electric drive rotors,more aircraft use electric rotors to provide thrust and directional control.For a helicopter tail rotor,the wake of the main rotor influences the tail rotor’s in flow and wake.In the procedure of controlling,crosswind will also cause changes to the tail disk load.This paper describes requirements and design principles of an electric motor drive and variable pitch tail rotor system.A particular spoke-type architecture of the motor is designed,and the performance of blades is analyzed by the CFD method.The demand for simplicity of moving parts and strict constraints on the weight of a helicopter makes the design of electrical and mechanical components challenging.Different solutions have been investigated to propose an effective alternative to the mechanical actuation system.A test platform is constructed which can collect the dynamic response of the thrust control.The enhancement of the response speed due to an individual motor speed control and variable-pitch system is validated.

1.Introduction

In analogy with the aircraft electrification trend referred to as the More Electric Aircraft(MEA)approach,helicopters propulsion and their flight control devices are also experiencing a system optimization by the adoption of electric actuators.1–2Envisioned benefits for the development of electrically powered alternatives to hydraulically,pneumatically,or mechanically powered systems are:optimization of the power distribution,3,4easier and reduced maintenance due to the unified and simplified integration,5,6standardization of components,7,8new solutions and potential architectures not available before,7,9reduced noise generated by the tail rotor;reduced take-off weight.10,11

In this paper,one challenge is the development of a directdrive Electrical Tail Rotor Drive Control(ETRDC)system12in replacement of the traditional actuation system that receives power by means of torque tubes and gearboxes.

The main task for a tail rotor is to counteract the torque generated by the main rotor,and thus strong reliability and safety requirements for a tail rotor is in the first place of its design.Moreover,a direct-drive actuation system shall be installed in the tail with strict constraints in terms of size and weight.

Permanent Magnet(PM)motors13represent a promising solution to achieve high power density,and they exhibit well-known features such as high efficiency and good operating performance.

Most of the motor-driven lift components are used on quadrotors.The electric propulsion system of a typical UAV includes the following components:(A)blades,(B)a gear-box(optional),(C)a cooling system(optional),(D)an electric motor,(E)a driver,(F)an energy source,(G)wiring,plugs,and connectors.Most of the systems do not own blades of a large size.14Traditional methods for blade design are based on the well-known work by Betz and Prandtl from1919,15and such an optimized design was equipped in Rutan’s Voyager.16,17

Bouabdallah and Siegwart18has described a method for iteratively designing a vehicle with a maximum limitation of mass and length to achieve a desired thrust-to-weight ratio.The method requires a database of actuators,batteries,and airframe components to calculate the total mass.Bershadsky et al.4has presented a database which parameterizes drive components to meet the need of design and optimization.

Generally,the control of thrust of small electric aircraft is realized by adjustment of the rotational speed of motors.19The aerodynamic interference between a helicopter’s main rotor and tail is complex,and the alternating load and flight direction of its fuselage makes its aerodynamic condition rigorous.

This paper presents a new contracted structure application on ETRDC with the purpose of accelerating the aerodynamic control response.A comparison with the traditional variable pitch tail structure is carried out,and a detailed design process and parameters are proposed.

Under the constraints of weight and volume,a direct pitch control actuator is developed in order to eliminate the number of transmission mechanisms.A test was conducted to verify the speed of control response,and the time consumption from one thrust level to another is about 80%less.

2.Parameterization of drive,control,and lift components

The main dimension of a helicopter is shown in Fig.1.l is the distance between the main shaft and the center of the tail rotor,RHBis the radius of the main rotor,RTRis the radius of the tail rotor,and εRTRis the distance between the main rotor tip and the center of the tail rotor.The maximum take-off weight is 230 kg,and the subsequent examples are calculated based on this aircraft.

Fig.1 Nomenclature of fuselage.

2.1.Tail rotor power

Considering the loss caused by the fuselage and the helicopter maneuver reserve factor,the moment balance equation is described by

where TTRis the thrust of the tail rotor,MHBis the torque of the main rotor,and ζynpis the maneuver reserve factor.The induced power of the main rotor PINDcan be described by

where THBis the main rotor thrust,and the relationship between the induced power and required power of the rotor can be described by

where ηOHBis the efficiency of blades,and P is the required power of the main rotor.The relationship20between the induced velocity viand the main rotor thrust THBcan be described by

Therefore,the relationship between the main rotor thrust and the induced power can be calculated by

where A is the rotor disk area,ρ is the air density,1.225 kg/m3,Δ = ρ0/ρ is the relative density of altitude,D is the diameter of the main rotor disk,and ηOHB=0.7-0.8 is the efficiency of blades.21,22

The relationship between THBand the gravity of the whole fuselage G can be described by

where ζOδαis the blowing loss of the main rotor.

The relationship between the torque of the main rotor MHBand the disk load of the main rotor wDHBcan be described by

Substituting Eq.(1)into Eq.(7),we obtain

where ωHBis the rotational speed of the main rotor(rad/s).

Substituting Eq.(5)into Eq.(8),we obtain

where ηOTR=0.6–0.7 is the efficiency of the tail rotor.

Substituting Eq.(5)into Eq.(6),we obtain

The relative required power consumption of the tail rotor can be obtained from

Knowing the power requirement(27 kW)of the main rotor and the diameter ratio of the tail rotor to the main rotor being generally around 15%–20%,a ratio assumption,considering the center of gravity of the fuselage,can be made.In this paper,the ratio is 0.16,and P-TR=0.12,PTR=3.24 kW.The relationship between the radius ratio and the power consumption ratio is drawn in Fig.2.VHBtipis the tip velocity of the main rotor.Hence,the power consumption of the tail rotor is determined if the radius ratio is fixed.

2.2.Tail rotor design

Tail rotor blades are made of various composite materials including nylon-plastic,carbon fiber,wood,and other plastic.The mass of those blades can be found by4

where p1is 0.08884,and p2is 0 for wooden blades,0.05555 and 0.2216 for plastic,0.1178 and-0.3387 for nylon-reinforced plastic,and 0.1207 and-0.5122 for carbon fiber.In this paper,the blades are made of carbon fiber.

Blade performance is calculated by the CFD method.The mesh scale is of two million grids,and the minimum grid in the outer flow field is 0.01 mm while the maximum grid is 0.05 mm.The size of minimum grids in the inner flow field is 0.003 mm,and the maximum size is 0.01 mm.The rotational speed is from 1800 r/min to 5400 r/min.

Fig.2 Relationship between power ratio and radius ratio(VHBtip=200 m/s).

Fig.3 Distribution of twist angles.

Fig.4 Airfoil at different radii.

The diameter of the rotor disk is 720 mm,the distribution of twist angles is shown in Fig.3(in which r is the radial position of section,and R is radius of blade),and the airfoil at different radii is shown in Fig.4.

The flow field analysis result is shown in Fig.5.Comparisons of thrust,torque,and power are shown in Fig.6.The maximum thrust error between the CFD method and the experiment is 6.5 N.The maximum torque between the CFD method and the experiment is 0.5 N·m,and the maximum power between them is 0.1 kW.The CFD method has a great agreement with the experiment.

2.3.Motor design

Brushless DC motors are generally preferred over traditional brushed designs for their greater efficiency in converting electric energy to mechanical energy.Although In Runner(IR)motors allow them tighter due to the body of the motor being static,the Out Runner(OR)configuration allows motors to produce more torque than IR counterparts.13IR is a common choice for many small(＜100 g)multirotor builds,and OR is a good choice for larger structures.The motor efficiency is described as

Fig.5 Flow field of blades at 5000 r/min.

Fig.6 Comparisons of thrust,torque,and power by CFD method and experiment.

where T is the torque(N·m),n is the rotational speed(r/min),U is the voltage,and I is the current.The finite element mesh of the motor and analysis result are shown in Fig.7.A comparison between the Finite Element Method(FEM)and the experiment is shown in Fig.8,in which the maximum error is 1.5 N·m,and thus the results are of certain accuracy.

As shown in Fig.8,there is an error,nearly 14%,between the theoretical calculation of the motor and the actual measurement.The poles of the motor are made of N42SH,and the coercivity and remanence are selected according to the data from sheet.23The two properties affect the results obviously.Meanwhile,the materials of the poles between different motor manufacturers are different according to their own manufacturing processes.

Fig.7 Motor finite element analysis.

Fig.8 Comparison of motor torque.

The main data of the motor is shown in Table 1.

2.4.Actuation system

The screw nut mechanism has great advantages in transmission efficiency and accuracy.The reliability will be improvedif the screw nut mechanism is used for transmission,and the loss caused by gearing will be reduced.Details of the actuator is shown in Fig.9.Under rated operating conditions,the relationship between the screw’s thrust output and the torque input is described in Eq.(14),and the power which drives the screw Psis in Eq.(15).

Table 1 Main data of motor.

Fig.9 Structure of actuator.

Table 2 Main measured data of actuator.

where Tsis the input torque(N·m),F is the maximum load in duty load(N),Phis the screw lead(mm),nsis the rotational speed(r/min),ηp=0.9[ 1/((1+πd0/Ph)μ)]，μ =0.006，and d0is the nominal diameter(mm).

A multi-turn photoelectric encoder is applied to measure the rotation angle of the screw.The signal of the rotation angle is utilized by a control card to control the displacement of the screw.The main data of the actuator is listed in Table 2.

3.Comparisons of change pitch structure and weight

The propulsion system of a small UAV(batteries,motors,blades,etc.)accounts for as much as 60%of the vehicle’s weight.14Therefore,optimization of the propulsion system is extremely crucial.For most pitch variable structures,either motor-or transmission-driven tail rotors require two or three control pushrods,which are named as A,B,and C respectively in Fig.10,to convert the actuator control input to a total linear movement.Actuators need an additional bracket structure to be connected to the tail rotor system.Meanwhile,the structure must be designed and manufactured with high stiffness and strength.Otherwise,the deformation of the bracket will affect the position accuracy when the actuator is operating,and thus will further seriously affect the entire control system of the helicopter’s yaw precision.

Fig.10 General structure of a tail rotor.

Fig.11 Whole scale of electric tail rotor drive and pitch control system.

The best way to improve the reliability of a mechanical system is to reduce it.7In order to overcome the problem of the large number of parts brought by the traditional pitch control structure,the design proposal proposed in this paper will cancel control pushrods and use an actuator with a control pushrod coaxial with the motor centerline.This solution,in principle,avoids the use of control pushrods and eliminates the need for additional mounting brackets for the actuator.Due to the use of a coaxial actuator to change the pitch of blades,the linearity problem of the steering actuator will be avoided.Only one bracket is required for the connection of the entire tail rotor system to the helicopter fuselage,and hence a decoupled design of the tail rotor system and the fuselage structure can be achieved.The hub and cuff as well as the pitch control mechanism are enclosed in a cowling,which protects those rotating parts against dust and sand.Meanwhile,the smooth-shaped cowling that covers all the parts of the rotating mechanism together can reduce aerodynamic losses.

A comparison between the numbers of parts used by systems is demonstrated in Fig.11 and Fig.10,and the weights are listed in Table 3.The system weight reduction is 28.7%.

4.Tests and results

An experimental platform,as shown in Fig.12,was built in order to measure the steady-state,dynamic performance of the tail rotor thrust and the control output of the actuator.The height of the platform from the ground is 1.2 m,and the diameter of the bracket surface is nearly the same as that ofthe motor.A six-axis force/torque transducer is fixed to this bracket so that three directions of force and three torques can be measured.

Table 3 List of main parts of structures.

Fig.12 Experimental platform.

Experiments were carried out according to the following steps.Firstly,the six components of the transducer at different rotational speeds were measured,and followed by different pitches,at a fixed rotational speed. After the static test, the thrust response due to a step input of pitch and the thrust response due to a step input of rotational speed were recorded.The power consumption of the tail rotor in actual flight was recorded.

4.1.Thrust and power test results

The relationships between the thrust,output power of the motor,and rotational speed are shown in Fig.13,in which α is the angle of attack of section.The ratio of the output power of the motor to the input power of the motor drive is shown in Fig.14.Time of thrust history is listed in Table 4,which indicates that the time spent is greatly reduced.

Fig.13 Relationship between thrust,output power and rotational speed.

Fig.14 Efficiency of output power and drive input.

Table 4 Time of thrust change.

As shown in Fig.13,the thrust of blades as well as the output power increases with the rotational speed.The maximum thrust is near 200 N,which indicates that the results of the CFD method are correct.The input power of the tail rotor is the product of the input voltage and current of the motor controller.The output power of the tail rotor is the product of the motor torque and the rotational speed.

4.2.Transient time

A comparison of the consumption of time due to different structures is shown in Fig.15.The new design has improved the response speed by 86.7%.

Fig.15 Comparison of consumption of time due to different structures.

Fig.16 An XV-2 helicopter equipped with ETRDC.24

Fig.17 Power of tail rotor measured in one typical forward flight.

4.3.Flight tests

Flight tests were conducted with a helicopter equipped with the unique tail rotor proposed by the present study,shown in Fig.16.24The tests demonstrate that the present solution of tail rotor design meets the requirement of the power demanded,shown in Fig.17.In addition,the performance of the tail rotor is rather satisfactory.

5.Conclusions

(1)A design method is proposed for the helicopter electric tail rotor,which can determine the power requirement of the tail rotor system.A finite element analysis of the motor and a CFD method are used to accurately predict the torque of the tail rotor motor and the aerodynamic performance of the blades.Computation results are in good agreement with experimental data.

(2)A fast responding mechanism is proposed,aimed at overcoming the disadvantages of a single motor-driven rotor,which has a long adjustment period before the aerodynamic force starts to respond.Test data shows that the transient time is reduced by 86.7%.

(3)A compact pitch control mechanism is proposed,in which the number of components is significantly reduced and the weight reduction is 28.7%.

(4)The proposed mechanism is of practical value and has been equipped and tested on a helicopter.