NL2011184C2 - Helicopter. - Google Patents

Description

Helicopter

The invention relates to a helicopter provided with a main rotor and stabilizer wings at its tail, and provided with a control system to convert a pilot- or autopilot input into control actions for the helicopter.

Helicopters having these features are known in the prior art. In the prior art helicopter flight control system, a pitch command signal originating from the pilot or autopilot is converted to a longitudinal cyclic pitch input on the main rotor .

In general, rotorcraft can experience large structural loads on various components whilst performing manoeuvres in forward flight. This is the case for conventional helicopters but also for unconventional designs such as tilt rotor aircraft. The current invention focuses on helicopters with a main rotor and stabilizer wings at its tail.

In addition to the load factor, the structural load with helicopters is typically a function of; (a) the airspeed and the corresponding dynamic pressure and inflow distribution over the rotor disk, (b) the aircraft motion in terms of angular rates and (c) the control inputs in the rotor system (longitudinal cyclic, lateral cyclic and collective pitch). There are various solutions suggested to reduce the structural loads on specific rotorcraft components during manoeuvres. These vary from the physical re-design of the main rotor system to the use of active control systems. These active control systems either send a tactile cue to the pilot when a structural limit will be exceeded (allowing the pilot to override the system) or directly prevent a structural limit to be exceeded. An overview of critical rotorcraft loads encountered during manoeuvres and several technical solutions proposed to reduce or limit these loads is provided here in the form of several short examples.

For the Eurocopter TIGER helicopter, it was decided in the early stages of flight testing to perform load optimization of the main rotor system (physical re-design) to achieve a longer service life of critical components and to expand the flight envelope [1]. Amongst other critical loads, this helicopter, which features a hingeless main rotor system, experienced large pitch link loads during high speed pull ups and high speed (diving, right) turns in the flight testing campaign. The load optimization included a change in the main rotor hub geometry, main rotor tuning and an increase in the maximum allowable hub bending moment limit.

Huff et al. [2] investigated manoeuvring effects on the torque variability and resulting transmission vibrations of the AH-1 Cobra helicopter. A health and usage monitoring system (HUMS) is used to measure the structural loads encountered during flight. A HUMS enables: (a) the detection of faults, (b) increased awareness of the pilots and (c) timely maintenance based on the actual structural loading spectrum [3]. It does not limit or reduce structural loads.

Critical manoeuvres with respect to structural loads for the Bell Boeing V-22 Osprey tilt-rotor were determined by King et al. [4]. It was shown that pull-ups both in helicopter and airplane mode can result in large oscillatory main rotor yoke bending moments. Operations at the extremes of the flight envelope or during rolling pull-outs can potentially lead to flapping limit excursions on the V-22 Osprey. Further high roll rate manoeuvres in airplane mode are the cause of large torque variations in the drivetrain. The issues of the V-22 Osprey tilt-rotor were solved by King et al [4] through nonlinear control laws that make active use of all helicopter and airplane mode controls. Another load alleviating control law for the V-22 Osprey tilt-rotor is presented by Kimball [5].

This control law makes use of a model following control method that protects the drive system from over torque and reduces the pilot workload throughout the flight envelope.

Detailed investigations were performed into the steady and dynamic loads of the UH-60A Black Hawk helicopter [6, 7]. These investigations included extensive flight testing and high fidelity simulations. It was revealed that severe loads occur in the main rotor system during high speed manoeuvres.

The loads under investigation were the pitch link load, flap-wise bending load and chordwise bending load.

Horn and Sahani [8] demonstrated that the hub moments of the UH-60A helicopter can be protected with a carefree handling system that uses software algorithms to predict hub moment limit excursions and provides tactile cueing to the pilot. This strategy is highly effective to limit structural loads but can also directly affect the handling qualities such as attitude quickness.

In another study, Horn et al. [9] present a carefree handling system that prevents the XV-15 tilt-rotor from exceeding angle of attack limits and load factor limits and which provides tactile cueing during manoeuvres to limit longitudinal rotor flapping. This system makes use of neural networks to predict the flight envelope limits. A physical re-design of the main rotor blades (composite blades instead of metal blades) of the Sikorsky S-61 helicopter and its impact on the oscillatory pitch link loads was described by Curtiss et all. [10].

Modern control techniques were applied to a nonlinear simulation model of the Eurocopter Eurotilt tilt-rotor concept [11]. A Linear Quadratic Gaussian formulation was used to minimize rotor flapping and in-plane bending of the rotor in airplane mode. Furthermore, the torque split on the interconnect drive shaft in asymmetric manoeuvres was reduced. This last issue was also tackled with an active control law for the European tilt-rotor configuration ERICA [12]. This control law made use of a combination of differential collective pitch and airplane mode controls (rudder and aileron). It was demonstrated that handling qualities could be improved whilst achieving a significant load reduction.

In the design of the RAH-66 Comanche helicopter flight control laws, an angular acceleration limit was included in the command model to limit the hub moments for aggressive pitch manoeuvres [13].

According to Howitt [14], limiting the hub moments becomes a larger issue in general with the introduction of bearingless main rotor systems for conventional helicopters. He proposes to make use of rotor state (flap angle) feedback within the control laws to limit the hub moments and thereby to provide intrinsic carefree handling. Two approaches are suggested by Howitt to limit the hub moments. In the first approach, the flap angle command is limited directly by the active control law. In the second approach, a tactile cue is sent to the pilot when the flap angle demand is expected to result in exceedance of hub moment limits.

In a study by Johnson et al. [15], rotor state feed- back is also used to regulate longitudinal and lateral flapping on a high fidelity model of the AH-64D Apache helicopter for flight envelope limit protection. This approach is considered similar to that of Howitt.

In accordance with the efforts in the prior art, an object of the invention is to limit the loads (in particular hub moments and pitch link loads) on the main rotor system of the prior art helicopter that features a horizontal stabilizer. In prior art flight control systems, the pitch command signal originating from the pilot or autopilot is converted to a longitudinal cyclic pitch input on the main rotor.

Conversely, the helicopter of the invention is provided with the features of one or more of the appended claims.

Essentially according to the invention, the control system is arranged to convert a pitch command signal derived from either the pilot or the autopilot input into simultaneous first and second control signals for the main rotor and stabilizer wings respectively. An advantage of this is that this manner of control may improve at the same time both handling qualities of the helicopter and the reduction of its structural loads without costly redesign. The derivation of the first and second control signals for the main rotor and stabilizer wings respectively, can be implemented in either software or in hardware .

Appropriately the first control signal for the main rotor and the second control signal for the stabilizer wings are derived from the pitch command signal in a fixed ratio.

Further preferably the fixed ratio is dependent on the helicopter's forward speed. A beneficial feature is further that it includes a turn coordination system providing pitch and yaw rate command signals based on the helicopter's actual roll angle.

The invention is also embodied in a method of controlling a helicopter. Features of this method are provided in the appended claims 5-8.

The invention will hereinafter be further elucidated with reference to the drawing of an exemplary embodiment.

In the drawing: -figure 1 relates to the control system of the helicopter of the invention; -figure 2 concerns the part of the control system of the invention that relates to lateral directional control; -figure 3 shows aircraft attitude response for moderate amplitude doublet input at 130 kn forward flight; -figure 4 shows hub forces and moments for moderate amplitude doublet input at 130 kn forward flight; -figure 5 shows longitudinal control activity for moderate amplitude doublet input at 130 kn forward flight; -figure 6 shows aircraft velocity, altitude and load factor for moderate amplitude doublet input at 130 kn forward flight; and -figure 7 shows main rotor load quickness, defined as the peak load during a maneuver divided by the peak pitch attitude change, for hub forces and hub moments.

The architecture of the longitudinal control system of the invention is depicted in Figure 1. The functions of the outer loop of the control system can also be provided by a human pilot. When using an auto pilot (outer loop), the input from a stick displacement by the pilot is converted into a load factor demand signal ncommand. Based on the forward airspeed in the body axis system (u) and the gravitational constant (g), this demand signal is converted into a pitch command signal qcommand. The pitch command signal is filtered through a low pass filter and summed with the pitch command signal from the turn coordination system qTC. This turn coordination system will be discussed hereinafter with reference to figure 2.

Next, the measured pitch signal (qmeasured) is subtracted from the total pitch command signal in order to determine a pitch error signal. This error signal is minimized by a control law, usually a proportional + integral (PI) control law. The output signal of this control law is sent to a speed dependent gearing ratio (marked as 'SLA control allocation') which divides the signal into a longitudinal cyclic pitch command signal for the main rotor and a horizontal stabilizer incidence angle command signal for the stabilizer wings at the helicopter's tail. The function of the box SLA control allocation can be implemented in either software or in hardware. A human pilot can also directly provide a pitch control input to the box 'SLA control allocation' without making use of the outer loop of the control system.

The previously mentioned box 'SLA control allocation' comes in addition to the prior art system. In a prior art control system, the link between the output of the (PI) control law and the longitudinal cyclic pitch for the main rotor is direct without intermediate functionality.

Another aspect of the invention relates to the issue that helicopters can have a large off-axis response (e.g. roll due to pitch). In order to be able to execute pure pitch manoeuvres, a feature of the invention is a lateral directional control system as shown in Figure 2. It is designed to provide an attitude command attitude hold (ACAH) response type signal for the roll axis and a rate command (RC) response type signal for the yaw axis. For this purpose preferably a separate turn coordination system (TC) is present which creates a pitch and yaw rate command signal additional to the pilot command signal, and which is based on the current roll angle of the helicopter. It is remarked that the this lateral directional control law functions can also be provided by a human pilot.

When using the lateral directional autopilot, the pilot lateral stick input is converted into a roll angle command signal (phicommand) and subseguently filtered with a low pass filter. The pedal input provided by the pilot is converted into a yaw rate command signal (rcommand). Again control laws, usually PI control laws, are used to control the roll angle and yaw rate based on feedback of the measured roll angle, roll rate and yaw rate. Due to the presence of the turn coordination system as shown in figure 2, the pilot does not have to use the helicopter pedals to execute coordinated turns.

The results achieved with the helicopter of the invention are exemplified with reference to a set of manoeuvres of varying aggressiveness that are performed to evaluate the control system of the helicopter of the invention.

The main manoeuvre under investigation is a doublet (pull-up / pushover) based on the description of the Utility Tactical Transport Aircraft System (UTTAS) manoeuvre. The UT-TAS manoeuvre is specified as follows [16]:

From a level unaccelerated flight condition at 150 knots equivalent airspeed (KEAS), it shall be possible to attain, within 1.0 second from the initial control input, a sus tained load factor of 1.75 In a symmetrical pull-up. Following this load factor buildup, it shall be possible to maintain a minimum load factor of 1.75 for 3.0 seconds after the initial attainment of 1.75. Airspeed at the end of the 1.7 5g, 3.0-second duration segment of the maneuver shall not be less than 130 KEAS. At no time during this manoeuvre shall it be necessary to change the main rotor collective control from that required for the initial level unaccelerated flight condition. Also, from a level unaccelerated flight condition at 150 KEAS, it shall be possible to attain, within 1.0 second from the initial control inputr a sustained load factor of 0.0 in a pushover. Following the attainment of this load factorr it shall be possible to maintain a load factor of 0.0 for 2.0 seconds.

At no time during this maneuver shall it be necessary to change the main rotor collective control from that required for the initial level unaccelerated flight condition. At no time during either the pull-up or pushover maneuvers described above shall angular deviations in roll and yawr greater than +5 degrees from the initial unaccelerated level-flight conditions, be permitted.

The doublet lasts for 4 seconds (2 seconds pull-up and 2 seconds pushover). The aircraft attitude response for a moderate amplitude doublet is presented in Figure 3, followed by Figure 4 - Figure 6 that present the corresponding hub moments and forces, the helicopter motion in terms of speed altitude and load factor and finally the required control activity.

Figure 3 - Figure 6 show that both control systems provide an almost identical aircraft motion. A load factor of 1.75 'g' is achieved for the moderate amplitude manoeuvre, which starts at a forward airspeed of 130 knots. The airspeed reduces by about 15 knots and the main rotor collective pitch is unused. The angular deviations in roll and yaw remain within plus or minus 5 degrees due to the activity of the lateral directional control law. Hence, the manoeuvre complies with the UTTAS specifications. The only difference is a lower initial speed and a slightly shorter manoeuvre time. However, it should be noted that also more aggressive manoeuvres are conducted. It can be observed that the method of control according to the invention results in a very effective reduction of the peak hub moments and hub forces.

The push-pull manoeuvre described in the previous paragraph was performed at three levels of aggressiveness by varying the amplitude of the pilot input. In order to assess the handling qualities, both the maximum achieved load factor and pitch attitude quickness were determined as shown in Table 1, below.

Table 1.

From the experiments it can be concluded that the handling qualities of both the prior art longitudinal control system and the structural load alleviation (SLA) control system of the invention are similar in terms of bandwidth, pitch attitude quickness and off-axis response. However, the resulting structural loads on the main rotor hub in the system of the invention is significantly better.

The loads acting on the main rotor system and hub can be distinguished in vibratory loads (high number of cycles) and the maximum loads encountered during a manoeuvre. The maximum loads are expressed in terms of load quickness [see 17] in order to make a fair comparison between manoeuvres of different amplitude and duration. The load quickness is defined as: the maximum force or moment during a specific manoeuvre divided by the maximum attitude change. In this example, only longitudinal manoeuvres are considered. The load quickness parameter can also be applied to assess asymmetric manoeuvres.

Three doublet inputs were performed at 130 knots forward flight at three levels of aggressiveness. The doublet input had a total duration of 4 seconds (2 seconds positive in put, 2 seconds negative input). The resulting load quickness for the hub forces and hub moments are presented in Figure 7.

It can be observed that the load alleviation is highly effective especially for the hub moments and forces in the X and Y direction. The hub force and moment in the Z direction is hardly affected by the new means of control. This is in agreement with analytical hub loads and hub moment equations [18] since the motion of the helicopter is mainly in pitch, with only small off-axis deviations. Furthermore, the control input on the main rotor is longitudinal cyclic pitch combined with lateral cyclic pitch. For the prior art control method, the magnitude of the longitudinal cyclic pitch input is about three times larger than the lateral cyclic pitch input. When in accordance with the invention the horizontal stabilizer is used as well, longitudinal and lateral cyclic pitch have a similar magnitude.

The average load alleviation for the different control methods is summarized in Table 2. The largest load reduction is the roll moment (56.9%) . When all hub forces and moments are combined, an average load reduction of 25.3% is achieved for nearly identical manoeuvres in terms of aircraft motion and handling qualities.

Table 2.

Literature: 1. Gotzfried, K., "Survey of tiger main rotor loads from design to flight test," 23rd European Rotorcraft Forum, Dresden, Germany, September 1997. 2. Huff, E. M., Turner, I. Y., Barszcz, E., Dzwonczyk, M., McNames, J., " An analysis of maneuvering effects on transmission vibrations in an AH-1 Cobra Helicopter," Journal of the American Helicopter Society, Vol. 47, No. 1, January 2002 . 3. Anonymous, "Health and Usage Monitoring Systems Toolkit, US Joint Helicopter Safety Implementation team" Technical report, International Helicopter Safety Team, 2013. 4. King, D. W., Dabundo, C., Kisor, R. L., Agnihotri, A., "V-22 load limiting control law development," 49th annual forum of the American Helicopter Society, St. Louis, Missouri, May 1993. 5. Kimball, D. F., "Recent tilt rotor flight control law innovations," Journal of the American Helicopter Society, Vol. 32, No. 3, July 1987. 6. Datta, A., Chopra, I., "Validation and Understanding of UH-60A vibration Loads in Steady Flight", 58th Annual Forum of the American Helicopter Society, 11-14 June 2002, Montreal,

Canada 7. Kufeld, R. M., Bousman, W. G., "High load conditions measured on a UH-60A in maneuvering flight," 51st annual forum of the American Helicopter Society, Fort Worth, Texas, May 1995. 8. Horn, J. F. and Sahani, N., "Detection and avoidance of main rotor hub moment limits on rotorcraft," Journal of Aircraft, Vol. 41, No. 2, March-April 2004. 9. Horn, J., Calise, A. J., Prasad, J. V. R., "Flight envelope limit detection and avoidance for rotorcraft," Journal of the American Helicopter Society, Vol. 47, No. 4, October 2002, pp. 253-262. 10. Curtiss, H. C., Carson, F., Hill, J., Quackenbush, T., "Technical Note - Performance of a Sikorsky S-61 with a new main rotor," Journal of the American Helicopter Society, Vol. 48, No. 3, July 2003, pp. 211-215. 11. Manimala, B., Padfield, G. D. and Walker, D. J., "Load alleviation for a tilt-rotor aircraft in airplane mode," Journal of Aircraft, No. 1, January-February 2006. 12. Voskuijl, M., Pavel, M. D. and van der Vorst, J., "Active control technology for tiltrotor structural load alleviation," 30th European Rotorcraft Forum, Marseille, France, 2004, pp. 295-309. 13. Kothmann, B. D., Armbrust, J., "RAH-66 Comanche Core AFCS Control law development: DEMVAL to EMD," 58th annual forum of the American Helicopter Society, Montreal, Canada, 2002. 14. Howitt, J., "Application of nonlinear dynamic inversion to rotorcraft flight control," 61st annual forum of the American Helicopter Society, Grapevine, Texas, June 2005. 15. Johnson, C. K., Mittleider, D., Tischler, Μ. B., Cheung, K. K., "Analysis, design, & optimization of the helicopter active control technology (HACT) Flight control system," 58th annual forum of the American Helicopter Society, Montreal, Canada, June 2002. 16. Yamakawa, G. M., Broadhurst, D. G., Smith, J. R., "Utility tactical transport aircraft system (UTTAS) maneuver criteria," Technical Report, US Army Aviation Systems Command, AVSCOM, St Louis, Missouri, April 1972. 17. Pavel, M. D., Padfield, G. D., "The extension of ADS-33-metrics for agility enhancement and structural load alleviation," Journal of the American Helicopter Society, Vol. 51, No. 4, 2006, pp. 319-330. 18. Padfield, G. D., "Helicopter Flight Dynamics," Blackwell Science, Oxford, 1996.

Claims (8)

Helicopter provided with a main rotor and stabilizer wings at its take-off and provided with a control system for converting a pilot or autopilot input into control actions for the helicopter, characterized in that the control system is adapted for converting a pilot speed control signal which is derived from the pilot or autopilot input in simultaneous first and second control signals for the main rotor and the stabilization wings, respectively. Helicopter according to claim 1, characterized in that the first control signal for the main rotor and the second control signal for the stabilization wings are derived in a fixed ratio from the speed control signal. Helicopter according to claim 2, characterized in that the fixed ratio is dependent on the forward speed of the helicopter. A helicopter according to any one of the preceding claims 1-3, characterized in that it has a turning coordination system, which provides speed and yaw angle speed control signals based on the current rolling angle of the helicopter. 5. Method for controlling a helicopter provided with a main rotor and stabilizing wings at its take-off, and provided with a control system for converting a pilot or autopilot input into control actions for the helicopter, characterized in that a speed control signal, which is derived from the pilot or autopilot input, is converted into simultaneous first and second control signals for the main rotor and the stabilization wings, respectively. Method for controlling a helicopter according to claim 5, characterized in that the first control signal for the main rotor and the second control signal for the stabilization wings are derived in a fixed ratio from the speed control signal. Method for controlling a helicopter according to claim 6, characterized in that the fixed ratio is determined with reference to the forward speed of the helicopter. A method for controlling a helicopter according to claim one of the preceding claims 5-7, characterized in that, for coordinating rotations, speed and yaw angle speed, control signals are determined based on the current rolling angle of the helicopter.