The Pennsylvania State University
The Graduate School
Department of Aerospace Engineering
INCORPORATING FLIGHT HANDLING QUALITIES REQUIREMENTS INTO THE
DESIGN OF FUTURE HIGH-SPEED COMPOUND ROTORCRAFT
A Thesis in
Aerospace Engineering
by
Angelina Conti
2018 Angelina Conti
Submitted in Partial Fulfillment of the Requirements for the Degree of
Master of Science
May 2018
The thesis of Angelina Conti was reviewed and approved* by the following:
Joseph F. Horn Professor of Aerospace Engineering Thesis Advisor
Edward C. Smith Professor of Aerospace Engineering
Amy R. Pritchett Department Head of Aerospace Engineering
*Signatures are on file in the Graduate School
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ABSTRACT
The next fleet of rotorcraft will need to achieve higher forward flight speeds to meet the demanding requirements of military operation. One potential solution is to offload the main rotor by compounding with a lifting wing and auxiliary propulsion system. Compound rotorcraft also feature additional control effectors (flaperons, elevators, propeller thrust, etc.), which can be used to augment control response of the aircraft. To leverage the benefits of fly-by-wire flight controls and meet flight handling qualities requirements, it is essential that controller design be incorporated early in the development process.
Modern handling qualities design standards, such as ADS-33E-PRF, provide specifications that can be implemented in early design phases. Predictive strategies relate signal response characteristics to desired flight handling qualities specifications, and piloted simulation verifies the results for various maneuvers called Mission Task Elements. This thesis incorporates predictive and experimental handling qualities analysis into the design of a future high-speed compound rotorcraft and its fly-by-wire flight control system. The simulated rotorcraft features a compounded wing and pusher propeller to offload the main rotor, various drag clean-ups and larger installed power to reach high forward flight speeds, and a full-authority dynamic inversion controller to optimize handling qualities.
This study focuses on the combined effect of controller and design parameter variations on handling qualities. Initial results revealed that the controller command filter parameter impacts the bandwidth of the aircraft, where increasing this parameter generally improved handling qualities for all cyclic pitch deflections. The addition of redundant controls improved handling qualities by relieving control saturation, and trends from variation of flapping hinge offset and wing/tail size were generally unclear. Future research studies should apply alternative predictive and piloted simulation methods in order to identify and verify trends.
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TABLE OF CONTENTS
List of Figures ...... vi
List of Tables ...... x
Acknowledgements ...... xi
Chapter 1 Introduction ...... 1
Flight Dynamics of Rotorcraft ...... 1 Traditional Rotorcraft ...... 1 Compound Rotorcraft ...... 7 Motivation for Research ...... 11 Involvement of Flight Handling Qualities in the Design Process ...... 13 Research Goals ...... 15
Chapter 2 Rotorcraft Handling Qualities Requirements and Proposed Updates for Future Configurations ...... 17
Metrics for Quantitative/Predictive Handling Qualities in Forward Flight ...... 18 Small Amplitude/Moderate to High Frequency Metrics ...... 18 Moderate to Large Amplitude/Low to Moderate Frequency Metrics ...... 22 Mission Task Elements (MTEs) for Piloted Simulation ...... 25 Combat Break Turn MTE ...... 27 Sum of Sines (SOS) Tracking MTE ...... 29 Attitude Capture and Hold (BACH/PACH) MTE ...... 32 High Speed Acceleration-Deceleration (Accel/Decel) MTE ...... 34 Decelerated Approach MTE ...... 35
Chapter 3 Modeling and Simulation of a Generic Compound Rotorcraft ...... 38
Flight Dynamics Model ...... 38 Flight Controller ...... 40 Trim Analysis ...... 45 Penn State Flight Simulator ...... 47
Chapter 4 Handling Qualities Analysis of Baseline Configuration ...... 50
Command Filters ...... 50 Roll Command Filter ...... 51 Pitch Command Filter ...... 64 Flaperon/Stabilator Gearing ...... 73
Chapter 5 Vehicle Design Modifications ...... 78
Flapping Hinge Offset ...... 78 Wing Span and Tail Area ...... 84
v
Chapter 6 Concluding Remarks ...... 94
Conclusions ...... 95 Command Filter Predictive Analysis ...... 95 Command Filter Piloted Simulation Analysis ...... 96 Flaperon/Stabilator Gearing Predictive Analysis ...... 98 Flapping Hinge Offset Predictive Analysis ...... 99 Wing Span and Tail Area Predictive Analysis ...... 100 Future Work ...... 101 References ...... 104
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LIST OF FIGURES
Figure 1-1. Tilting the rotor plane enables forward flight (Glauert’s flow model from Ref. [1])...... 2
Figure 1-2. The differences in relative speed due to forward flight (from Ref. [1])...... 3
Figure 1-3. An articulated rotor system allows each blade to reach equilibrium though flapping, lead-lag, and feathering (from Ref. [1])...... 4
Figure 1-4. Blades respond to variation of dynamic pressure by flapping up or down to decrease or increase the lift produced as a function of azimuth angle (from Ref. [1]). ... 5
Figure 1-5. 1949 Fairey ‘Jet Gyrodyne’ (from Ref. [8])...... 7
Figure 1-6. Piasecki X-49A demonstrator aircraft (from Ref. [2])...... 8
Figure 1-7. The differences in power required for conventional and compound rotorcraft (from Ref. [2])...... 10
Figure 1-8. Bell Helicopter’s entry to the JMR flight demonstration, the V-280 Valor (from Ref. [14])...... 12
Figure 1-9. Sikorsky-Boeing’s SK>1 Defiant (from Ref. [16])...... 13
Figure 1-10. Generic compound rotorcraft used in flight handling qualities research at PSU (from Ref. 21)...... 16
Figure 2-1. ADS-33E-PRF definitions from the frequency response of the linearized model (from Ref. [20]...... 19
Figure 2-2. ADS-33E-PRF handling qualities criteria for pitch-axis (left) and roll-axis (right) Bandwidth/Phase Delay (from Ref. [20])...... 20
Figure 2-3. ADS-33E-PRF handling qualities criteria for roll and pitch cross-coupling (from Ref. [20])...... 21
Figure 2-4. Time history of the aircraft response to a step-input command in roll or pitch (from Ref. [20])...... 22
Figure 2-5. ADS-33E-PRF handling qualities criteria for attitude quickness (from Ref. [20])...... 23
Figure 2-6. Time history response of a system to a 5-second stick input (from Ref. [25]). .... 24
Figure 2-7. Military standard for pitch attitude dropback in response to a 5-second stick input (from Ref. [25])...... 25
Figure 2-8. Cooper-Harper Handling Qualities Rating (HQR) Scale (from Ref. [23])...... 26
vii
Figure 2-9. Perpendicular runways at Mid-state Airport in State College, PA (left) and explanation for its application to the Break Turn MTE (right) (from References [27, 26])...... 28
Figure 2-10. Compensatory Tracking display (from Ref. [28])...... 29
Figure 2-11. SOS Tracking “bowtie” display (from Ref. [28])...... 30
Figure 2-12. Examples of the PSOS (top) and RSOS (bottom) command signals (from Ref. [28])...... 31
Figure 2-13. Example of the pitch command signal for ACH MTE (from Ref. [29])...... 32
Figure 2-14. Overview of the PAPI cue (from Ref. [31])...... 36
Figure 3-1. High-level block diagram of the controller’s response to pilot lateral cyclic control deflection...... 43
Figure 3-2. Bode plot of the system response...... 44
Figure 3-3. Collective trim allocation of the baseline model in trimmed forward flight...... 45
Figure 3-4. Power required as a function of propeller pitch for various airspeeds...... 46
Figure 3-5. Pitch attitude, lateral cyclic, longitudinal cyclic, and collective required for minimum power required at various airspeeds...... 47
Figure 3-6. Penn State flight simulator facility cab and visual display...... 48
Figure 3-7. HUD provided to pilots in the Penn State flight simulator facility...... 48
Figure 4-1.Bandwidth analysis for configurations of various roll command filter
parameters 휔푛훷...... 53
Figure 4-2. Cross-coupling analysis for configurations of various 휔푛훷...... 54
Figure 4-3. Explanation of Cross-Coupling Analysis results...... 55
Figure 4-4. Roll Attitude Quickness analysis for configurations of various 휔푛훷...... 57
Figure 4-5. Lateral cyclic command during the RAQ test (signal length of 1.0 sec) for
various 휔푛훷...... 58
Figure 4-6. Roll Attitude Quickness results for configurations of various airspeeds and
휔푛훷...... 59
Figure 4-7. Break Turn MTE results for configurations of various airspeeds and 휔푛훷...... 60
Figure 4-8. RSOS MTE results for configurations of various 휔푛훷...... 61
viii
Figure 4-9. BACH MTE results for configurations of various 휔푛훷...... 63
Figure 4-10. Bandwidth analysis for configurations of various pitch command filter
parameter 휔푛훳...... 65
Figure 4-11. Cross-coupling analysis for configurations of various 휔푛훳...... 66
Figure 4-12. Pitch Attitude Dropback analysis for configurations of various 휔푛훳...... 67
Figure 4-13. Pitch Attitude Dropback analysis for configurations of various airspeeds and
ω푛훳...... 68
Figure 4-14. Results from the PSOS (left) and PACH (right) MTEs for configurations of
various ω푛훳...... 69
Figure 4-15. High Speed Acceleration MTE results for configurations of various 휔푛훳...... 72
Figure 4-16. Bandwidth/Phase Delay analysis for configurations of various flaperon allocations 휔푑푓푙푎푝 (top) and configurations of various stabilator allocations 휔푑푠푡푎푏 (bottom)...... 74
Figure 4-17. Cross-coupling analysis for configurations of various flaperon allocations 휔푑푓푙푎푝 (left) and configurations of various stabilator allocations 휔푑푠푡푎푏 (right)...... 75
Figure 4-18. Roll Attitude Quickness analysis for configurations of various flaperon allocations 휔푑푓푙푎푝...... 76
Figure 5-1. Bandwidth/Phase Delay analysis for configurations of various flapping hinge offsets using three flaperon allocations휔푑푓푙푎푝 (top) and three stabilator allocations 휔푑푠푡푎푏 (bottom)...... 80
Figure 5-2. Cross-Coupling Analysis for configurations of various flapping hinge offsets using three flaperon allocations 휔푑푓푙푎푝 (left) and three stabilator allocations 휔푑푠푡푎푏 (right)...... 81
Figure 5-3. Roll Attitude Quickness analysis for configurations of various hinge offset and flaperon allocation 휔푑푓푙푎푝 (top row and bottom left) and Pitch Attitude Dropback analysis for configurations of various hinge offset and stabilator allocation 휔푑푠푡푎푏 (bottom right)...... 83
Figure 5-4. Bandwidth/Phase Delay analysis for configurations of various wing sizes using three flaperon allocations 휔푑푓푙푎푝 (top) and configurations of various tail sizes using three stabilator allocations 휔푑푠푡푎푏 (bottom)...... 88
Figure 5-5. Cross-Coupling Analysis for configurations of various wing sizes using three flaperon allocations 휔푑푓푙푎푝 (left) and configurations of various tail sizes using three stabilator allocations 휔푑푠푡푎푏 (right)...... 90
ix
Figure 5-6. Roll Attitude Quickness analysis for configurations of various wing sizes and flaperon allocation 휔푑푓푙푎푝 (top row and bottom left) and Pitch Dropback analysis for configurations of tail sizes and stabilator allocation 휔푑푠푡푎푏 (bottom right)...... 92
x
LIST OF TABLES
Table 2-1. Handling qualities criteria for the Combat Break Turn MTE (from Ref. [26])...... 27
Table 2-2. Handling qualities criteria for the SOS Tracking MTE (from Ref. [28])...... 31
Table 2-3. Handling qualities criteria for the BACH MTE (from Ref. [29])...... 33
Table 2-4. Handling qualities criteria for the PACH MTE (from Ref. [29])...... 34
Table 2-5. Handling qualities criteria for the Accel/Decel MTE (from Ref. [30])...... 35
Table 2-6. Handling qualities criteria for the Decelerated Approach MTE (from Ref. [21])...... 37
Table 4-1. Configurations involved in the study of the command filter...... 51
Table 4-2. Configurations involved in the study of the roll-axis command filter...... 52
Table 4-3. Configurations involved in the study of the pitch-axis command filter...... 64
Table 4-4. Configurations involved in the study of redundant control gearing...... 73
Table 5-1. Configurations involved in the study of flapping hinge offset...... 79
Table 5-2. Configurations involved in the study of wing/tail size...... 85
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ACKNOWLEDGEMENTS
I would like to thank the U.S. Army’s National Rotorcraft Technology Center program for partially funding this research between Vertical Lift Consortium, Inc. and the U.S. Government under Other Transaction number W15QKN-10-9-0003. I would also like to thank my research advisor Dr. Joseph F. Horn for his guidance through this project, as well as Dr. Ed Smith and Tom
Berger for reviewing my thesis. Finally, I would like to thank my family and friends for their support through graduate school and beyond. The past few years have challenged me in more ways than one, and I would not have made it this far without their support. Thank you.
1
Chapter 1
Introduction
Flight Dynamics of Rotorcraft
Traditional Rotorcraft
The history of the helicopter spans over several millennia, beginning with the creation of a child’s toy, the Chinese top, in 400 B.C. The concept of vertical flight intrigued the world’s brightest minds, and although many inventors developed their ideas into prototypes, they were unsuccessful due to technological barriers. Leishman notes in Ref. [1] that there are 7 fundamental technical problems that hindered the development of the helicopter:
1. Understanding aerodynamics of vertical flight
2. Engine technology
3. Weight
4. Balancing rotor-torque
5. Stability and control
6. Vibrations
7. Surviving engine failure.
It was not until the 1940s that Igor Sikorsky resolved these problems with the VS-300A. Most modern helicopters follow a similar design to that of the VS-300A, which consisted of a single main rotor, tail rotor, and variable pitch blades.
The main rotor is the primary source of lift, propulsion, and control. While a fixed-wing aircraft must move forward to generate lift, a helicopter produces lift in stationary flight through
2 rotating several wings about a centralized shaft. The main rotor produces a nearly symmetric lift distribution in hovering flight, and the rotor-torque is balanced through a tail rotor. To obtain forward flight, the rotor plane must tilt to produce a propulsive component, shown in Figure 1-1.
Figure 1-1. Tilting the rotor plane enables forward flight (Glauert’s flow model from Ref. [1]).
The pilot tilts the rotor plane by changing the blade pitch angles as a function of azimuth angle using cyclic control, and the aircraft accelerates (produces more thrust from the main rotor) by collectively changing the blade pitch angles through collective control. As the helicopter moves forward (increasing advance ratio, µ), the relative wind impacting each blades changes as a function of azimuth angle (휓), shown in Figure 1-2.
3
Figure 1-2. The differences in relative speed due to forward flight (from Ref. [1]).
The relative airspeed (given as tip Mach number 푀푡푖푝 in Figure 1-2) impacting the advancing blade increases to include the rotational speed (훺푅) of the main rotor and also the forward flight speed
(푉∞). Conversely, the relative airspeed impacting the retreating blade decreases since the rotational speed of the main rotor and the forward flight speed of the helicopter are in opposite directions. For a fixed blade pitch angle (no cyclic control), the advancing side of the rotor plane generates more lift than the retreating side due to the larger relative airspeed, creating an imbalance. A rigid rotor prevents out-of-plane motion, and therefore the imbalance creates a large rolling moment which degrades controllability of the rotorcraft. Rigid blades also result in very high bending stress on the blades. Articulation of the main rotor allows the blades to independently change orientation in response to the aerodynamic properties throughout its rotation to relieve the lift imbalance and bending stresses. This can be accomplished through hinges that enable flapping (up and down
4 motion) and lead-lag (in-plane motion) as well as a bearing that enables feathering (twisting motion), shown in Figure 1-3.
Figure 1-3. An articulated rotor system allows each blade to reach equilibrium though flapping, lead-lag, and feathering (from Ref. [1]).
In hover, the symmetric lift distribution results in blades that flap up and lag back at all azimuth angles. The blades flap up to balance aerodynamic and centrifugal forces (creates a coning angle), and the blades lag back due to aerodynamic drag. The blades respond to the lift imbalance in forward flight by varying flapping and lead-lag motion as a function of azimuth angle based on the relative airspeed, shown in Figure 1-4.
5
Figure 1-4. Blades respond to variation of dynamic pressure by flapping up or down to decrease or increase the lift produced as a function of azimuth angle (from Ref. [1]).
The excess lift of the advancing side causes the blade to rise (flap up), which decreases the local blade element angle of attack. Lift continues to decrease until the blade passes the front of the rotor disk, where it reaches its maximum displacement. Dynamic pressure continues to decrease on the retreating side and causes the blade to fall (flap down), thereby increasing the local blade element angle of attack. Lift continues to increase until the blade passes the back of the rotor disk, where it reaches its minimum displacement. Therefore, it can be clearly seen that the maximum
6 aerodynamic force and maximum flapping displacement have a 90° phase difference, which creates a natural tendency for the rotor plane to tilt backward. Additionally, the presence of a coning angle results in lateral flapping where the local blade element angle of attack increases at the front of the rotor disk and decreases at the back of the rotor disk. Lateral flapping is 90° out of phase with longitudinal flapping (discussed above), and therefore the aircraft also has a tendency to tilt to the right. The pilot can control and trim these flapping responses using cyclic pitch as discussed below.
The primary axes of control are identical to those of fixed-wing aircraft and include pitch, roll, yaw and thrust as mentioned in Ref. [2]. They are controlled using four main effectors: main rotor collective, longitudinal cyclic, lateral cyclic, and tail rotor collective. The pilot uses main rotor collective to increase thrust and accelerate the helicopter forward, aft, up or down.
Longitudinal and lateral cyclic are used to vary blade pitch as a function of azimuth angle, and therefore it enables the pilot to control the orientation of the rotor disk in roll and pitch. Due to the
90° phase lag inherent to the main rotor, cyclic pitch changes also create flapping motion 90° out of phase to the cyclic input. Lastly, the pilot uses the tail rotor collective to control yaw by increasing or decreasing the blade pitch of the tail rotor.
The flight envelope of a helicopter is limited primarily by the aerodynamics of the main rotor as discussed in References [3, 4], as well as parasite drag and installed engine power as explained in Ref. [5]. As advance ratio increases, the relative airspeed impacting the retreating blade decreases until the onset of stall, which leads to large variations in blade load, high vibrations, and possibly even loss of control. On the other side, the relative airspeed impacting the advancing blade increases until the onset of compressibility, which causes drag divergence that drastically degrades performance. One way to delay the onset of compressibility and stall to higher forward flight speeds is to decrease the speed of the main rotor as shown in References [3, 4]. Without some type of lift or propulsion augmentation, this might not be possible as the stall problem becomes worse. However, the reduced thrust of the main rotor can be compensated through the addition of
7 a wing and propeller, which supplement lift and thrust, respectively. This concept is formally called a compound rotorcraft.
Compound Rotorcraft
Interest in compound rotorcraft stems from the late 1940s due to increasing demand for higher forward flight speeds as discussed in References [3, 4, 6, 7, 8]. An early attempt was the
1949 Fairey Gyrodyne, shown in Figure 1-5, which demonstrated speeds in excess of 200 kts, but was not commercially successful.
Figure 1-5. 1949 Fairey ‘Jet Gyrodyne’ (from Ref. [8]).
Through the 1950s and 1960s, companies such as McDonnell, Bell, Kaman, Lockheed, and
Sikorsky developed their own compound rotorcraft, but they too did not reach mass production due to political issues, funding, and continued design changes (discussed in Ref. [6]). Modern rotorcraft of the 21st century still cannot reach the demanding forward flight speeds required for many operations, so again researchers are investigating compound rotorcraft as a potential solution. Such projects include the X-49A (shown in Figure 1-6), X2 and X3 demonstrator aircraft.
8
Figure 1-6. Piasecki X-49A demonstrator aircraft (from Ref. [2]).
Compound rotorcraft attempt to compromise between traditional rotorcraft and fixed-wing aircraft, and the result is a hybrid aircraft capable of two distinct modes of flight. The addition of a wing and supplementary propulsion system offloads the main rotor, enabling lower rotational speed of the main rotor and therefore higher advance ratios before the onset of compressibility and stall
(further explained in References [7, 9]). In fact, studies have shown that compounding enables rotorcraft to reach forward flight speeds as high as 220 kts and increases maneuverability to 4g at
200 kts (as discussed in Ref. [4]). This opens the opportunity for new missions and increased performance in existing high-speed rotorcraft missions, including air ambulance, sea and mountain rescue, police surveillance, corporate services, oil rig servicing, troop transport and antitank gunships (listed in Ref. [1]).
Offloading the main rotor also has the advantage of creating a more level fuselage attitude
(discussed in Ref. [6]). Thrust of a compound helicopter is primarily supplied by the auxiliary propulsion system in forward flight, thus decreasing the tilt required by the main rotor. One benefit of a more level fuselage attitude is that it decreases parasite drag, which is important for minimizing power required during high-speed flight (explained in Ref. [4]). Leveling the rotor plane also
9 increases passenger comfort and could possibly reduce vibrations associated with wake interactions
(explained in References [4, 10]).
The control system of compound rotorcraft typically consists of primary and redundant controls, where the primary controls mimic those of a traditional helicopter. The wing and auxiliary propulsion system provide two additional control effectors (wing flaperons and propeller pitch, described in Ref. [2]). Increasing the number of control effectors naturally has the potential to increase the pilot’s workload (there is no unique trim condition), and therefore future compound rotorcraft will most likely use an automatic flight control system. However, the benefit of redundant controls is that control allocation can be modified to reach some sort of goal, such as maximizing performance and minimizing power. The flight control system should optimize the lift and thrust allocation between the control effectors to enhance response characteristics while avoiding control saturation. Control saturation occurs when a controller reaches a physical or structural limit, where control input from the pilot results in an undesired response from the vehicle (ex. limited physical rotation and rotational rate of wing flaperons and tail stabilator). Wing flaperons are used on compound rotorcraft as a redundant control of the longitudinal and lateral vehicle dynamics.
Deflecting flaperons differentially contributes to the rolling moment of the vehicle, while deflecting them symmetrically increases the lift produced by the wing for a given forward flight speed
(influences lift sharing between the wing and the main rotor). Some compound rotorcraft also have a stabilator on the tail, where deflection results in a pitching moment of the aircraft.
Several problems result from compounding rotorcraft, particularly in regard to hover and low-speed flight performance. Compounding increases the weight of the vehicle, ultimately leading to worse hover performance. Additionally, compounding hardly contributes to the required loads during low-speed flight and the wing produces a download effect (explained in Ref. [11]). As a result, the main rotor must cover the required loads and compensate for inefficiencies, which leads
10 to higher power required in hover and low-speed flight than that of conventional rotorcraft, shown in Figure 1-7.
Figure 1-7. The differences in power required for conventional and compound rotorcraft (from Ref. [2]).
Although the vehicle performs better than conventional rotorcraft at high forward flight speeds, the power required is significantly higher than that of comparable fixed-wing aircraft due to higher parasite drag (explained in Ref. [11]). Minimizing drag is essential to maximizing performance of the compound rotorcraft at high speeds. Finally, the compound rotorcraft is generally more complex than conventional rotorcraft and fixed-wing aircraft, which could lead to a more expensive and risky design (explained in References [10, 11]).
Autorotation is an important capability of a rotorcraft because it allows it to gradually descend in the case of engine failure, explained in Ref. [1]. Compounding complicates autorotation because the wing provides additional lift that lowers the aircraft’s sink rate, which in turn lowers the flow up through the rotor that drives autorotation (shown in Ref. [6]). Spoilers are often used in
11 this case to drastically decrease the lift supplied by the wing, thus supporting the sink rate and autorotation. They are also used in hover to reduce the download created by the wing.
The design of a compound helicopter is unique because it combines methodologies from rotorcraft and fixed-wing aircraft (further explained in Ref. [6]). Two important dimensions that affect wing performance are wing area and aspect ratio (defined as span squared over wing area).
Increasing wing area ultimately leads to increased lift at the cost of drag. However, it is also desirable to minimize wing area to decrease the download produced by the wing. A balance must be made so that the wing effectively supports the compound through a particular velocity range.
Aspect ratio generally affects the efficiency of the wing, where the increase in aspect ratio of a fixed-wing aircraft increases the wing’s efficiency by reducing the induced drag in cruising flight.
However, a large aspect ratio on a compound rotorcraft can actually degrade its performance since the wing tip interacts with the high-velocity rotor wake at the edge of the rotor plane. Therefore, designers again must compromise, normally settling at an aspect ratio of 6 for compound vehicles.
Motivation for Research
Traditional rotorcraft are reaching the limits of their performance potential in forward flight. The demanding environment of military operation suggests that a new all-purpose rotorcraft capable of controlled hover and forward flight, such as the compound rotorcraft, is required for future missions. To support this endeavor, the United States Army is leading the Future Vertical
Lift (FVL) initiative, which plans to replace the U.S. Department of Defense’s current fleet of rotorcraft with next-generation rotorcraft, beginning with UH-60 Black Hawk transports and AH-
64 Apache gunships (stated in Ref. [12]). The requirements will push the limits of the traditional helicopter, likely resulting in a family of multi-role configurations that utilize various degrees of
12 compounding and control redundancy, along with common systems and open architecture as suggested in Ref. [13].
The Joint Multi-Role Technology Demonstration is perhaps the most well-known part of the FVL initiative. The goal of the program is to demonstrate various technical approaches industry could take when designing a family of system-compatible multi-role next-generation rotorcraft, specifically one capable of at least 230 kts in forward flight with a design gross weight of approximately 30,000 lbs (medium-lift capability) as discussed in Ref. [13]. Initially, four teams were selected to conduct configuration trade studies, which included a preliminary assessment of the risk involved in each approach. Two of these teams were chosen to proceed to flight demonstration, each using drastically different designs to achieve hover and high-speed flight. The
Bell V-280 Valor, displayed in Figure 1-8, is an advanced tiltrotor that transitions between hover and high-speed modes of operation through tilting rotors mounted on a main wing as discussed in
Ref. [14].
Figure 1-8. Bell Helicopter’s entry to the JMR flight demonstration, the V-280 Valor (from Ref. [14]).
In hover, the rotors are horizontal and perform similarly to a traditional rotorcraft. In forward flight, the rotors tilt forward so that they are vertical and produce forward thrust similarly to a propeller-
13 driven aircraft. The aircraft is very efficient at its high cruise speed of 280 kts, but the transition between flight modes takes a significant amount of time, which hinders its performance in some missions (explained in Ref. [15]). The competing rotorcraft, the Sikorsky-Boeing SB>1 Defiant
(shown in Figure 1-9), which uses the Collier Trophy-winning X2 demonstrator technology, takes a different approach by compounding a coaxial rotorcraft with a pusher propeller to provide forward thrust (explained in References [16, 17]).
Figure 1-9. Sikorsky-Boeing’s SK>1 Defiant (from Ref. [16]).
Although the coaxial rotors solve the problem of retreating blade stall, this configuration still experiences harsh vibrations due to advancing blade compressibility (further explained in Ref.
[15]). Severity is reduced through integrated active vibration control technology. Although both aircraft are awaiting flight demonstration, there seems to be an emerging need for research into compounded aircraft in order to meet the requirements of next-generation rotorcraft.
Involvement of Flight Handling Qualities in the Design Process
Preliminary design of rotorcraft involves optimizing performance to meet specific requirements. Traditionally, performance is initially evaluated by cruise speed, range, hover
14 altitude, weight, power requirements, and fuel-efficiency. Although these attributes help predict the rotorcraft’s physical capability to complete missions, they do little to predict the pilot workload and associated control systems required to make it happen. Such considerations are typically left out of the design process until flight test, where they may become a problem and require fixes. Research suggests that 25-50% of flight testing time in an aircraft development program might be spent on fixing handling qualities problems (stated in Ref. [18]). Fixes are often expensive and could have been avoided if handling qualities assessment were more integrated into the initial design process.
Incorporating handling qualities into the initial design process is complicated due to the low fidelity of models at this stage. Preliminary design typically involves an optimization algorithm that sweeps through design variables (typically configuration geometry) to meet requirements while maximizing particular performance characteristics (mentioned in References [11, 18]). These analyses are typically of a static bare-airframe (without a control system) and are therefore not conducive for a flight dynamics analysis. Research suggests that conceptual design tools should be redesigned to calculate flight handling qualities characteristics and compare results against industry standards (argued in Ref. [18]).
ADS-33E-PRF is a government specification of requirements for desirable and adequate flight handling qualities in terms of rotorcraft characteristics typically available in the initial design phase. According to the specification, ADS-33E-PRF (Ref. [19]) is “intended to assure that no limitations on flight safety or on the capability to perform intended missions will result from deficiencies in flying qualities.” It provides requirements for flight handling qualities throughout the flight profile and in simple representative missions (Mission Task Elements, further explained in Ref. [20]). However, future rotorcraft are destined to push the limits of the current ADS-33E-
PRF specification, and therefore ADS-33E-PRF flight handling qualities requirements must be modified (suggested by Ref. [21]). Research at the Pennsylvania State University is part of a larger
15 effort to study the appropriateness and effectiveness of current flight handling qualities requirements for assessing next-generation rotorcraft.
Research Goals
To support the development of flight handling qualities standards, the Pennsylvania State
University (PSU) is studying the impact of design changes on the flight handling qualities of a next- generation compound rotorcraft (further explained in Ref. [21]). The goal of this research is to develop the ADS-33E-PRF specification so that it effectively assesses the handling qualities of high-speed rotorcraft in aggressive maneuvers. This might involve modifying existing requirements and/or creating new ones, which could support the FVL initiative. ADS-33E-PRF uses two general methods of assessing handling qualities. The first method uses quantitative characteristics of flight behavior generated from computer-automated simulation to predict the handling qualities of the aircraft. The second involves pilot evaluation of rotorcraft simulation in various maneuvers called Mission Task Elements (MTEs). If a clear relationship is identified between quantitative characteristics of the aircraft and its actual flight handling qualities, then perhaps these metrics could be incorporated in the initial design process.
The results discussed in this research provides supporting data that could serve to update
ADS-33E-PRF so that it is relevant for next-generation rotorcraft, specifically compound rotorcraft in high-speed flight. Existing methods are improved, new methods are developed, and all methods are tested in computer-automated simulation as well as in piloted simulation. To represent a possible next-generation aircraft, this research assesses the flight handling qualities of a UH-60
Black Hawk rotorcraft with a conventional main rotor and tail rotor compounded with a propeller for thrust and a wing for lift, shown in Figure 1-10.
16
Figure 1-10. Generic compound rotorcraft used in flight handling qualities research at PSU (from Ref. 21).
The flight dynamics model of this compound rotorcraft was created to study flight control design of next-generation rotorcraft and the impact of redundant control effectors on handling qualities as discussed in Ref. [21]. It is generic in the sense that it does not represent an actual experimental or production rotorcraft, thus it is appropriate for preliminary design analysis. The model is teamed with a nonlinear dynamic inversion (NLDI) controller developed at PSU and is piloted using the
PSU Vertical Lift Research Center of Excellence (VLRCOE) Rotorcraft Flight Simulator. Aside from analyzing the handling qualities of the base model, trade studies are conducted to study the impact of controller settings, main rotor hinge offset, redundant control gearing, and wing/stabilator size.
17
Chapter 2
Rotorcraft Handling Qualities Requirements and Proposed Updates for Future Configurations
ADS-33E-PRF (References [19, 20]) is the current industry-standard military rotorcraft handling qualities specification. It uses quantitative metrics to predict rotorcraft handling qualities and piloted simulation to verify the results. Although some methods were specifically designed for forward flight, there is a chance that future high-speed rotorcraft will push the specification to its limits. Therefore, it is important to review the current specification and assess whether it should be updated to include metrics that are more relevant to rotorcraft capable of high-speed precision and aggressive maneuvers.
Research at the Pennsylvania State University (PSU) focuses on the evaluation of handling qualities of a high-speed rotorcraft compounded with a lifting wing and auxiliary propulsion system. Analysis methods include those in ADS-33E-PRF as well as new requirements proposed by various members in industry (discussed in References [25, 26, 28, 29, 30, 31]). The compound is generic in the sense that it does not model any existing rotorcraft, but it includes features that are likely to be seen in future rotorcraft, such as a wing and an auxiliary propulsion system for compounded lift and thrust. The airframe is based on a FORTRAN model of a UH-60 Black Hawk
(GENHEL) that has been used in rotorcraft research for decades as seen in Ref. [22]. The model was updated through research at PSU (see Ref. [2]) to create the generic compound rotorcraft used today. A nonlinear dynamic inversion flight controller designed for the generic compound rotorcraft is modelled in MATLAB/Simulink and links directly to GENHEL to simulate the flight dynamics of the aircraft.
18 Metrics for Quantitative/Predictive Handling Qualities in Forward Flight
The modern approach to handling qualities is to define quantitative criteria that characterize desirable handling qualities based on actual flight test or pilot simulation data, described in Ref. [20]. Bounds are defined to identify configurations with desirable (Level 1), adequate (Level 2), or undesirable (Level 3) handling qualities. The goal of predictive assessment is to define criteria such that the handling qualities can be determined in the early stages of design when only low- or moderate-fidelity models are available. As a result, these criteria typically involve parameters obtained from frequency response or simple step-input commands.
Small Amplitude/Moderate to High Frequency Metrics
When the maneuver amplitude range is small and the frequency is moderate to high, such as in precision tracking, an aircraft’s closed-loop handling qualities can be quantitatively studied through the phase and gain of the linearized system’s frequency response of attitude to pilot’s cyclic control, explained in References [19, 21]. In this case, the parameters of interest are bandwidth frequency and phase delay.
In practice, bandwidth represents the range of frequencies where the pilot can effectively control the aircraft without excessive compensation, such as through lead/lag, as explained in Ref.
[24]. ADS-33E-PRF distinguishes between gain-limited and phase-limited bandwidth, shown in
Figure 2-1, where the bandwidth parameter used in handling qualities criteria (ωBW) is the lesser of the two values as defined in Ref. [20].
19
Figure 2-1. ADS-33E-PRF definitions from the frequency response of the linearized model (from Ref. [20].
Numerically, gain-limited bandwidth is defined in References [20, 23] as the frequency where the gain margin is 6 dB, while phase-limited bandwidth is the frequency where the phase margin is 45°
(where the attitude lags behind the controller by 135°). Although gain-limited bandwidth is important for determining pilot-induced oscillation (PIO) tendencies, attitude command/attitude hold response-type (ACAH) systems are primarily concerned with the phase-limited bandwidth.
Ref. [23] explains that the bandwidth parameter does not give a complete description of a system’s handling qualities, and therefore a second term must be introduced. As frequency increases past the phase-limited bandwidth frequency, the phase of an ideal system (no time delay
20 or higher-order dynamics) approaches -180°. However, the phase for a non-ideal system passes -
180° at ω180 and continues to decrease past 2 ∗ ω180. Research shows that there is a unique relationship between the phase delay (shape of the curve between ω180 and 2 ∗ ω180) and the overall handling qualities of the rotorcraft as seen in Ref. [23]. As a result, the phase delay parameter is included with the bandwidth parameter in the ADS-33E-PRF handling qualities criteria for small amplitude/moderate to high frequency response characteristics.
Numerically, phase delay (τp) is defined as the slope of the phase curve between ω180 and
2 ∗ ω180 for a linear curve, or the slope of a linear least-squares curve fit for a nonlinear curve.
Equation 2-1 describes the phase delay as seen in References [20, 23].
훥훷2휔180 (2-1) 휏푝 = 57.3(2 ∗ 휔180)
In practice, phase delay represents the sensitivity of the system’s resonance to change in pilot gain, but it can also be considered as the initial delay to pilot step inputs, shown in Ref. [24].
Phase delay and bandwidth frequency are compared to criteria from ADS-33E-PRF (Ref.
[20]) for pitch and roll axes, shown in Figure 2-2 for Target Acquisition and Tracking MTEs.
Figure 2-2. ADS-33E-PRF handling qualities criteria for pitch-axis (left) and roll-axis (right) Bandwidth/Phase Delay (from Ref. [20]).
21 Level 1 indicates desired handling qualities, which is represented by high bandwidth and low phase delay. It is important to note that the boundaries of the criteria are dependent on the particular mission, where a precision tracking MTE has stricter criteria than general MTEs as explained in
References [20, 23].
The frequency response of rate to pilot’s cyclic control is used to describe cross-coupling behavior of the aircraft between the roll and pitch axes for aggressive maneuvers. The ADS-33E-
PRF requirements are based on the relationship between the response ratios, 푝/푞 and 푞/푝, derived from the amplitudes of the frequency response functions averaged between ωBW and ω180 as explained in Ref. [23]. Pitch due to roll (푞/푝) and roll due to pitch (푝/푞) are compared to criteria from ADS-33E-PRF (Ref. [21]), given in Figure 2-3, where low 푞/푝 and low 푝/푞 is desired (Level
1).
Figure 2-3. ADS-33E-PRF handling qualities criteria for roll and pitch cross-coupling (from Ref. [20]).
22 Moderate to Large Amplitude/Low to Moderate Frequency Metrics
While bandwidth frequency, phase delay, and cross-coupling are sufficient parameters for small amplitude/high frequency maneuvers (e.g., precision tracking), they are inappropriate for evaluating moderate to large amplitude/low to moderate frequency response characteristics (e.g., maximum maneuvering) as seen in Ref. [23]. Instead, parameters for available control power and attitude quickness are used to evaluate control effectiveness in these maneuvers.
For the attitude quickness test, the aircraft begins in steady-level flight and then responds to a step command in roll (which might be achieved by a large pulse or step input on the pilot inceptor depending on the rotorcraft response type). The size of the input directly impacts the response, and therefore the tests are completed for multiple step-sizes to establish a trend with increasing changes in attitude. ADS-33E-PRF uses three terms in its evaluation of handling qualities within this amplitude and frequency range. Two of these terms (ΔΦpk and ΔΦmin) represent the overshoot of the aircraft attitude in response to the command and are measured from the time history, shown in Figure 2-4.
Figure 2-4. Time history of the aircraft response to a step-input command in roll or pitch (from Ref. [20]).
The last term is attitude quickness, which is defined in ADS-33E-PRF as the ratio of the peak rate of change to the peak attitude change. Attitude quickness (ppk⁄ΔΦpk) versus ΔΦmin is compared
23 to criteria from ADS-33E-PRF, shown in Figure 2-5, where Level 1 corresponds to desirable handling qualities. Note that the boundary becomes more relaxed for larger amplitude maneuvers.
Figure 2-5. ADS-33E-PRF handling qualities criteria for attitude quickness (from Ref. [20]).
Again, the boundaries of the criteria are dependent on the mission. This method can be evaluated numerically or through piloted simulations.
The quickness method in ADS-33E-PRF is appropriate for describing roll control effectiveness in the entire flight regime, but is only appropriate for pitch control effectiveness in low-speed and hover MTEs. As a result, a separate method must be used to evaluate moderate amplitude pitch control characteristics in forward flight. Since the methods described in ADS-33E-
PRF for this case are mostly qualitative, an alternative method is used that analyzes the system’s response to a 5-second stick input with varied input size. An example of aircraft response to this input is shown in Figure 2-6 from Ref. [25].
24
Figure 2-6. Time history response of a system to a 5-second stick input (from Ref. [25]).
Pitch attitude dropback represents the overshoot of the response to stick input and is defined in Ref. [25] as the difference between maximum pitch attitude and the steady-state pitch attitude. The Pitch Attitude Dropback specification evaluates the relationship between the peak rate of change and the dropback, both non-dimensionalized by the steady state rate of change. Results are compared to specific performance criteria in MIL-STD-1797B. An example is shown in Figure
2-7, where low dropback and pitch rate overshoot are desired for Level 1 handling qualities.
25
Figure 2-7. Military standard for pitch attitude dropback in response to a 5-second stick input (from Ref. [25]).
Mission Task Elements (MTEs) for Piloted Simulation
Predictive handling quality metrics are useful during the design phase, but results have minimal meaning unless they are verified with actual flight test or piloted simulation data.
Environmental factors including visual degradation, weather variance, and even instrumentation display contribute to an aircraft’s operational handling qualities, but are difficult to quantify using predictive handling quality metrics as explained in Ref. [23]. As a result, it is important to supplement the evaluation with flight test or piloted simulation of a higher-fidelity model to identify safety concerns, performance limitations, and undesired handling qualities, shown in Ref. [20].
The current industry standard, ADS-33E-PRF, outlines several Mission Task Elements
(MTEs) that focus on the hover to low-speed flight regime but only includes a few MTEs for high- speed forward flight. Furthermore, the high-speed maneuvers included in ADS-33E-PRF were designed when rotorcraft fly-by-wire technology was still in its early stages. To update the specification, several new mission-oriented MTEs were developed through the Rotorcraft Handling
26 Qualities Requirements for Future Configurations and Missions project. This project was led by
Sikorsky Aircraft Corporation (SAC) and includes piloted simulation studies completed at SAC,
Bell Helicopter, The Boeing Company, Systems Technology, Inc. (STI), and PSU. The MTEs that were developed are discussed individually in this chapter and are unique by specified start and end conditions and spatial and temporal constraints as explained in Ref. [23]. Pilots compare results to specific performance criteria and then score the aircraft configuration using the Cooper-Harper
Handling Qualities Rating (HQR) Scale, given in Figure 2-8. Note that a score of 1 means that the aircraft handling qualities are highly desirable. Results are used to validate predictive metrics by evaluating the aircraft’s actual handling qualities in realistic missions.
Figure 2-8. Cooper-Harper Handling Qualities Rating (HQR) Scale (from Ref. [23]).
27 Combat Break Turn MTE
The Combat Break Turn maneuver evaluates handling qualities for non-precision, aggressive tasks such as those seen in evasive air combat maneuvers. Ref. [26] explains that it is used to assess aircraft handling qualities problems with respect to agility, including undesirable cross-coupling between pitch, roll, and yaw. The aircraft begins the maneuver by flying straight in steady-level flight at an altitude of 300 ft and a speed of 0.8 VH (160 kts for the PSU generic compound simulation), where VH is the maximum speed at maximum continuous power (described in Ref. [26]). The pilot quickly banks the aircraft right or left and then stabilizes again in straight, steady-level flight with a total heading change (훥휓) of at least 85°. Pilots complete this maneuver as quickly as possible within acceptable airspeed and altitude tolerances as described in Table 2-1, where the ideal time is defined in Equation 2-2.
휋푉(45 − 훥휓푡푟푎푛푠푖푒푛푡) (2-2) 푇푖푑푒푎푙 = + 2푇푡푟푎푛푠푖푒푛푡 2 90푔√푛푧푙푖푚 − 1
Table 2-1. Handling qualities criteria for the Combat Break Turn MTE (from Ref. [26]). Desired Adequate
Complete maneuver within time T < ΔT + Tideal. ΔT = 3.5 sec ΔT = 7.0 sec
Final change in directional flight path shall be at 95° 105° least 85° and no more than X°.
When rolling out to wings-level attitude, the 5° 10° overshoot in roll attitude shall not exceed X°.
Final airspeed loss shall be no more than X% of 10% 20% initial airspeed.
Maintain altitude within ±X feet 75 ft 150 ft
Any oscillations or inter-axis coupling shall not be: Undesirable Objectionable
28
In Equation 2-2 and Table 2-1, 푔 is the gravitational acceleration, 푉 is the flight speed, 푛푧lim is the normal load factor, 푇푡푟푎푛푠푖푒푛푡 is the time needed for one transient phase (roll-in or roll-out),
훥휓푡푟푎푛푠푖푒푛푡 is the heading change during a transient phase, Tideal is the ideal completion time, 푇 is the actual completion time, ΔT is the tolerance of completion time, and X is defined by desired and adequate criteria (two columns to the right). In piloted simulations at PSU, Ref. [26] shows that pilots use two perpendicular runways (X-Plane visual representation of Mid-State Airport in State
College, PA) as reference for heading during the maneuver, displayed in Figure 2-9.
Figure 2-9. Perpendicular runways at Mid-state Airport in State College, PA (left) and explanation for its application to the Break Turn MTE (right) (from References [27, 26]).
At the end of the MTE, parameters of interest are compared against performance criteria, and then the pilot scores the aircraft configuration using the Cooper-Harper Handling Qualities
Rating (HQR) Scale. Following flight simulation tests, researchers quantitatively assess the aircraft configuration’s handling qualities by analyzing the time histories of various control parameters, such as the aircraft rotation about the principal axes (phi, theta, and psi), roll rate, lateral stick input, and altitude.
29 Sum of Sines (SOS) Tracking MTE
The Sum of Sines (SOS) Tracking MTE evaluates handling qualities for precision, aggressive and precision, non-aggressive tasks, although only the aggressive version of this MTE is used in this paper. It is used to evaluate handling qualities in tight target tracking conditions, such as in air-to-air tracking and nap-of-the-earth tracking, and identify any undesirable characteristics of the aircraft, such as bobbling or Pilot Induced Oscillation (PIO) as explained in Ref. [28]. The pilot begins the task in steady-level flight at 160 kts and then tracks on-screen cues to change the aircraft’s pitch and/or roll for 60 seconds. The forcing function is based on a pilot-vehicle system crossover model, which states that the pilot-vehicle system can be represented by the crossover frequency and the effective system time delay at frequencies near the gain crossover frequency.
The crossover frequency can be used to describe the closed-loop system’s bandwidth, while the effective system time delay incorporates pilot-vehicle interaction, such as aircraft dynamics and pilot compensation. Since the forcing function is known, both the crossover frequency and the effective system time delay can be computed directly.
The displays used in this task provide the pilot with feedback regarding the aircraft’s attitude compared to the commanded attitude. The difference between the two is computed as the error, shown in Figure 2-10.
Figure 2-10. Compensatory Tracking display (from Ref. [28]).
30 The SOS Tracking MTE uses a “bowtie” display, shown in Figure 2-11, where the objective of roll
SOS (RSOS) is to keep the green line between the magenta boundaries, and the objective of pitch
SOS (PSOS) is to keep the green dot inside of the magenta circle.
Figure 2-11. SOS Tracking “bowtie” display (from Ref. [28]).
The command signal is generated using a Fibonacci series to roughly space frequencies evenly on a log scale across the frequency range of interest over a period of 60 seconds. The exact sequence can be modified using phase delays with the superposition of sinewaves by Equation 2-
3, where 푋푐 is displacement, 퐴푖 is amplitude, 휔푖 is frequency, 푡 is a variable, 휙푖 is phase, and 𝑖 is the number of sinewaves in the series.
7 (2-3) 푋푐 = ∑ 퐴푖푠𝑖푛 (휔푖푡 + 휙푖) 푖=1
An example of the RSOS and PSOS Tracking command signals are shown in Figure 2-12.
31
Figure 2-12. Examples of the PSOS (top) and RSOS (bottom) command signals (from Ref. [28]).
Pilots must stay within a tolerance of the command signal, which is defined by the “bowtie” from
Figure 2-11. Tolerances for desired and adequate handling qualities performance for roll and pitch tracking tasks are given in Table 2-2.
Table 2-2. Handling qualities criteria for the SOS Tracking MTE (from Ref. [28]). Desired Adequate Pitch: at least X% of the scoring time within 50% 75% pitch attitude error tolerance: ± 1° ± 2°
Roll: at least X% of the scoring time within 50% 75% roll attitude error tolerance: ± 5° ± 10°
PIO Considerations: No PIO tendencies No divergent PIO tendencies
Inter-axis coupling shall not be Undesirable Objectionable
32 Following completion of the task, pilots give the aircraft configuration a score using the Cooper-
Harper Handling Qualities Rating (HQR) Scale. Researchers then review the time histories and analyze the variability of the aircraft attitude from the command signal.
Attitude Capture and Hold (BACH/PACH) MTE
The Attitude Capture and Hold MTE concerns precision, non-aggressive handling qualities for high-speed flight. The pilot begins the maneuver in trimmed, steady-level flight at 0.8 VH (160 kts for the PSU generic compound simulation) and follows on-screen cues to change the aircraft attitude. An example of the command signals for pitch are shown in Figure 2-13.
Figure 2-13. Example of the pitch command signal for ACH MTE (from Ref. [29]).
A same “bowtie” display of the SOS Tracking MTE is used to show the aircraft’s attitude relative to the cued attitude, shown in Figure 2-10. This MTE is conducted for roll attitude changes (BACH) and pitch attitude changes (PACH) separately. The green dot must be within the magenta circle to capture pitch attitude changes, while the green line must be within the “bowtie” to capture roll
33 attitude changes. Once the attitude is achieved, the pilot holds the attitude for several seconds and then captures the next attitude. BACH requires the pilot to capture bank angle changes of +/-30° within 3 seconds and then hold the attitude for 5 seconds, given in Ref. [29]. Each test also requires the pilot to capture one bank angle change of 60° within 5 seconds and then hold for 5 seconds.
Handling qualities are evaluated using quantitative criteria, shown in Table 2-3, as well as pilot evaluation using the Cooper-Harper Handling Qualities Rating Scale.
Table 2-3. Handling qualities criteria for the BACH MTE (from Ref. [29]). Desired Adequate Bank angle error (from command) ± 5° ± 10° tolerance:
Airspeed deviation tolerance: ± 5 kts ± 10 kts
No more than one bank angle overshoot on 5° 10° the initial capture of each attitude. Magnitude of overshoot is less than:
PIO considerations: No PIO tendencies No divergent PIO tendencies
Inter-axis coupling shall not be Undesirable Objectionable
PACH requires the pilot to capture pitch attitude changes of +/- 5° within 2 seconds and then hold the attitude for 5 seconds. Although the BACH and PACH MTEs were designed to represent precise, non-aggressive handling qualities, the specifications can be modified to increase aggressiveness. Similar to the BACH MTE, handling qualities are evaluated using quantitative criteria, shown in Table 2-4, as well as pilot evaluation using the Cooper-Harper Handling Qualities
Rating Scale.
34 Table 2-4. Handling qualities criteria for the PACH MTE (from Ref. [29]). Desired Adequate Pitch angle error (from command) ± 1° ± 2° tolerance:
Airspeed deviation tolerance: ± 5 kts ± 10 kts
No more than one pitch attitude overshoot 1° 2° on the initial capture of each attitude. Magnitude of overshoot is less than:
PIO considerations: No PIO tendencies No divergent PIO tendencies
Inter-axis coupling shall not be Undesirable Objectionable
High Speed Acceleration-Deceleration (Accel/Decel) MTE
The High Speed Acceleration-Deceleration (Accel/Decel) MTE evaluates non-precision, aggressive tasks throughout the flight regime. It checks for undesirable coupling between the pitch and roll axes and ensures seamless transition from low-speed helicopter-mode to high-speed airplane-mode as explained in Ref. [30]. The aircraft begins the acceleration portion of the maneuver at 50 kts in forward flight at a specified altitude. The pilot then commands the aircraft to decelerate as aggressively as possible until a forward flight speed of VH-10 kts is reached. The pilot then completes the deceleration maneuver, where the aircraft begins at VH-10 kts in forward flight.
The pilot then commands the aircraft to decelerate as aggressively as possible until a forward flight speed of 50 kts is reached. The results are compared to performance criteria, shown in Table 2-5, and then the acceleration and deceleration maneuvers are scored separately using the Cooper-
Harper Handling Qualities Rating Scale.
35 Table 2-5. Handling qualities criteria for the Accel/Decel MTE (from Ref. [30]). Desired Adequate Acceleration Phase ΔT = 8 sec ΔT = 15 sec Complete acceleration within time T < ΔT + Tminimum
Deceleration Phase ΔT = 10 sec ΔT = 20 sec Complete deceleration and capture 50 kts within time T < ΔT + Tminimum
Both Phases ± 2 kts 10° Final airspeed to be captured within
Maintain altitude within X ft of initial altitude ± 100 ft ± 150 ft
Maintain heading within X° of initial heading ± 5° ± 10°
Maintain bank angle (from trim) within: ± 5° ± 10°
Any oscillations or inter-axis coupling shall not be: Undesirable Objectionable
Control harmony between axes shall not be: Undesirable Objectionable
Rotor RPM shall remain within the limits of X OFE SFE without undue pilot compensations:
Altitude and heading tolerances serve as guidelines so that they are not exchanged for performance benefits. Tminimum represents the acceleration and deceleration times at maximum performance and can be determined analytically (ex. performance charts) or empirically (ex. piloted simulation) prior to the maneuver.
Decelerated Approach MTE
The Decelerated Approach MTE is similar to the High Speed Deceleration MTE, but extends its application to precision, non-aggressive tasks. It is used to assess the workload required by the pilot to maintain a constant glideslope, to control airspeed, and to transition from high-speed mode to low-speed mode as explained in Ref. [21]. The pilot begins the maneuver in steady-level
36 flight at a distance 2 nmi from the target and at the airspeed for best range (160 kts for the PSU generic compound simulation). The pilot then descends at a constant glide slope of 4° towards a landing pad. To assist the pilot, the glide slope is cued using the lights of a Precision Approach
Path Indicator (PAPI), shown in Figure 2-14.
Figure 2-14. Overview of the PAPI cue (from Ref. [31]).
While following the glide slope, the pilot must decelerate the aircraft so that the final flight speed is 50 kts at a distance 0.25 nmi away from the landing pad. After completion of the maneuver, the pilot will give the aircraft configuration a score using the Cooper-Harper Handling Qualities Rating
(HQR) Scale. The tolerances that define desired and adequate handling qualities for this maneuver are described in Table 2-6.
37 Table 2-6. Handling qualities criteria for the Decelerated Approach MTE (from Ref. [21]). Desired Adequate Once pilot enters approach path, Never see four lights of If four lights are same glide slope tolerance is based on the same color (4 red, color, they shall remain in lights of PAPI system or 4 white) that condition for no more than 5 seconds.
Maintain lateral track within ±100 ft ±200 ft
Reach within ±X kts of final target ±5 kts ±10 kts airspeed (50 knots) at end of the maneuver (0.25 nm from runway)
Pilots are allowed to use auxiliary systems such as reverse thrust capability of the propeller to control airspeed in this MTE as explained in Ref. [21]. Following flight simulation tests, researchers analyze the time histories of various flight handling parameters, such as rotation about the principal axes (roll, pitch, and yaw), airspeed, altitude, and approach path, including flight path and top view.
38
Chapter 3
Modeling and Simulation of a Generic Compound Rotorcraft
Flight Dynamics Model
The goal of this research is to investigate the handling qualities of future high-speed compound rotorcraft, which has the potential to support the Future Vertical Lift (FVL) initiative.
One of the main areas of focus of the FVL initiative is the replacement of medium weight class rotorcraft, such as the UH-60A Black Hawk helicopter, which currently account for 75% of today’s operational military rotorcraft as seen in Ref. [13]. It is assumed that a compound rotorcraft in the same category as the Black Hawk would be similar in size, but have significantly less drag along with lift and thrust compounding to enable higher forward flight speeds as explained in Ref. [2].
Since this research aims to study a generic compound (one that does not replicate an existing design), it is assumed that the flight dynamics of a compound share many features to that of a Black
Hawk. In 1981, NASA published a report detailing the design of a 6 degree-of-freedom mathematical model of the Black Hawk based on the Sikorsky General Helicopter Flight Dynamics
Simulation, known as GENHEL as described in Ref. [22]. GENHEL treats the aircraft as a rigid body, but includes blade flapping and lagging in the blade-element model of the main rotor. It was validated with flight-test data and has become a well-established model in the rotorcraft community. Due to its reputation and its relevance to this project, GENHEL was selected as the core of the flight dynamics model used at the Pennsylvania State University (PSU) to support the development of new handling-qualities requirements and design methods.
Since GENHEL models a traditional rotorcraft, the software must be modified to include qualities of future rotorcraft configurations. Research at PSU led to the development of the
39 GENHEL-PSU compound model, which uses GENHEL at the core, but also includes an upgraded main rotor model along with modules for lift and thrust compounding. The GENHEL-PSU compound model uses the same airframe as the original Black Hawk, but reduces the flat plates drag by 5.93 ft2 to account for retractable landing gear and various drag clean-ups as explained in
Ref. [2]. The main rotor is similar to the Black Hawk’s, but has a smaller twist to reduce drag in forward flight. This of course has the penalty of increased power in hover, but enables reduced power in high-speed flight. Additionally, compounding alleviates the load of the main rotor, which allows the main rotor and the tail rotor to operate at lower rotational speeds. As discussed in Chapter
1, a lower rotational speed enables the aircraft to reach higher advance ratios before the onset of advancing blade compressibility and retreating blade stall. The main rotor operates under the ideal engine assumption of constant speed with an installed power of 4500 HP, which is larger than that of the baseline Black Hawk (3244 HP) as explained in Ref. [2].
The GENHEL-PSU compound rotorcraft model includes a module for lift compounding, which is achieved through the addition of a wing mounted in a high position. The wing was designed to compromise between high forward-flight speed and maneuverability as shown in Ref.
[2]. The result was a design point of 10% wing area to rotor disk area, which led to an equivalent usable— does not include the wing carry-through section— span of 45 ft and wing area of 226 ft2.
The wing has a linear taper, constant airfoil, and no dihedral, twist, or sweep. Two-dimensional aerodynamics are modeled for a NACA 63-412 airfoil, which was selected due to its appropriate lift-to-drag ratio, using tables of aerodynamic coefficients generated from XFOIL and corrections for finite wing effects at an Oswald efficiency factor of 0.85. The module also includes the effects of stall at large angles of attack, such as around hover. The wing features flaperons as a redundant control effector for roll in high-speed flight, which span about 90% of the wing span. Aerodynamic properties of the flaperons are modeled using tables generated from wind tunnel data for a NACA
23012 plain trailing-edge flap of 20% wing chord. Finally, aerodynamic interactions between the
40 wing and other components of the aircraft, specifically the main rotor, are modeled using look up tables of induced velocity components that were derived by Sikorsky’s Wake Analysis Program when the GENHEL code was developed.
Thrust compounding is accomplished through an auxiliary propulsion system mounted at the back of the aircraft (pusher propeller). It features variable pitch and constant speed, and can be used for forward and reverse thrust. The propeller is modeled for a Clark-Y airfoil using BEMVT, which is a combination of blade-element momentum theory and vortex theory, while high angles of attack are modeled using quasi-steady lift and drag coefficients as seen in Ref. [2]. These models were validated using wind tunnel results for an isolated propeller (NACA Technical Report 594 in
Ref. [33]). Propeller pitch serves as a redundant control for thrust.
Flight Controller
The controller used in this experiment was developed at PSU for research in the area of rotorcraft stability and control. It is based on a nonlinear dynamic inversion (NLDI) architecture, where a nonlinear system is modified via feedback linearization so that control design can be treated with linear methods as explained in Ref. [33]. The control design model represents the body-axis forces (푋, 푌, 푍, or tensor 퐹⃑) and body-axis moments (퐿, 푀, 푁, or tensor 푇⃑⃑) experienced by the aircraft using the exact nonlinear 6 degree-of-freedom rigid body equations of motion, shown in
Equations 3-1 to 3-4 from References [2, 34].
1 퐵 (3-1) 퐹⃑ = 푣⃑̇ + 휔⃑⃑⃑ 푥 푣⃑ 푚 푐 푐
̇ (3-2) 1 푋 푈̇ 0 −푅 푄 푈 푈 − 푄푊 − 푅푉 [푌] = [ 푉̇ ] + [ 푅 0 −푃] [ 푉 ] = [푉̇ + 푅푈 − 푃푊] 푚 푍 푊̇ −푄 푃 0 푊 푊̇ + 푃푉 − 푄푈
41
푇⃑⃑ = 퐻⃑⃑⃑̇ 퐵 + 휔⃑⃑⃑ 푥 퐻⃑⃑⃑ (3-3)
̇ ̇ (3-4) 퐿 퐼푥푥푃 + 퐼푧푧푅 0 −푅 푄 퐼푥푥 0 퐼푥푧 푃 [푀] = [ 퐼푦푦푄̇ ] + [ 푅 0 −푃] [ 0 퐼푦푦 0 ] [푄] 푁 −푄 푃 0 퐼 0 퐼 푅 퐼푧푧푅̇ + 퐼푥푧푃̇ 푥푧 푧푧
퐼푥푥푃̇ + 퐼푧푧푅̇ + 푄푅(퐼푧푧 − 퐼푦푦) + 푃푄퐼푥푧 2 2 = [퐼푦푦푄̇ + 푃푅(퐼푥푥 − 퐼푧푧) + (푅 − 푃 )퐼푥푧 ]
퐼푧푧푅̇ + 퐼푥푧푃̇ + 푃푄(퐼푦푦 − 퐼푥푥) − 푄푅퐼푥푧
푈, 푉, and 푊 (or tensor 푣⃑푐) are the body-axis velocities (longitudinal, lateral, and vertical), 푃, 푄, and 푅 (or tensor 휔⃑⃑⃑ ) are the body angular rates (roll, pitch, and yaw), 퐼푥푥, 퐼푦푦, and 퐼푧푧 are moments of inertia with respect to the principal axes, 푚 is mass, and 퐻⃑⃑⃑ is angular momentum tensor.
Scripts internal to GENHEL define the aircraft trim for a range of velocities and then perturb the aircraft forces and moments to define stability and control derivatives as explained in References
[2, 34]. These derivatives are used as coefficients to create linear models of the forces and moments of the aircraft, which are used in the controller model for design.
The state variables are the rigid body states, defined as the velocity vector v⃑⃑c and body axis angular rate vector ω⃑⃑⃑, along with Euler angles 훳 and Φ. The primary control vector includes the deflection in lateral cyclic control (훿푙푎푡), longitudinal cyclic control (훿푙표푛푔), and tail rotor collective control (훿푝푒푑). The state vector (푥⃑) and primary control vector (푢⃑⃑) are shown in Equations 3-5 and
3-6, respectively.
푥⃑ = [푈 푉 푊 푃 푄 푅 훳 훷]푇 (3-5)
훿푙푎푡 (3-6) 푢⃑⃑ = [훿푙표푛푔] 훿푝푒푑
Linear models derived in GENHEL are used to approximate the forces (푋, 푌, 푍) and moments (퐿,
푀, 푁) of the aircraft, which are applied to the nonlinear equations of motion to produce the nonlinear state derivative vector, shown in Equation 3-7, where 훳 and 훷 are the aircraft Euler
42 angles corresponding to aircraft pitch attitude and roll attitude, respectively. The output equation is shown in Equation 3-8 and definitions are shown in Equation 3-9 to Equation 3-16 as seen in
References [2, 34], where 푔 is acceleration due to gravity.
푇 푥⃑̇ = 푓⃑(푥⃑) + 푔⃑(푥⃑)푢⃑⃑ = [푈̇ 푉̇ 푊̇ 푃̇ 푄̇ 푅̇ 훳̇ 훷̇ ] (3-7)
훷̇ (3-8) 푦⃑ = ℎ⃑⃑(푥⃑) = [훳̇ ] 푅
푋 (3-9) 푈̇ = − 푔푠𝑖푛훳 − 푄푊 + 푅푉 푚
푌 (3-10) 푉̇ = + 푔푐표푠훳푠𝑖푛훷 − 푅푈 + 푃푊 푚
푍 (3-11) 푊̇ = + 푔푐표푠훳푐표푠훷 − 푃푉 − 푄푈 푚
1 ̇ 2 2 (3-12) 푃 = 2 [퐼푧퐿 + 퐼푥푧푁 + 퐼푥푧(퐼푥 − 퐼푦 + 퐼푧)푃푄 − (퐼푧 − 퐼푧퐼푦 + 퐼푥푧)푄푅] 퐼푥퐼푧 − 퐼푥푧
1 2 2 (3-13) 푄̇ = [푀 − (퐼푥 − 퐼푧)푅푃 − 퐼푥푧(푃 − 푅 )] 퐼푦
1 ̇ 2 2 (3-14) 푅 = 2 [퐼푥푁 + 퐼푥푧퐿 + 퐼푥푧(퐼푥 − 퐼푦 + 퐼푧)푄푅 − (퐼푥 − 퐼푥퐼푦 + 퐼푥푧)푃푄] 퐼푥퐼푧 − 퐼푥푧
훷̇ = 푃 + 푄푠𝑖푛훷푡푎푛훳 + 푅푐표푠훷푡푎푛훳 (3-15)
훳̇ = 푄푐표푠훷 − 푅푠𝑖푛훷 (3-16)
NLDI requires one controlled variable for each pilot input, which uses a full state feedback system as explained in References [2, 33]. When a pilot commands the control inceptors, the measured deflection is multiplied by a gain to give the commanded control rate. This value then passes through a transfer function (푡푓), shown in Equation 3-17, that defines the ideal response of the aircraft to the input, where ωn is referred to as the command filter parameter as explained in
References [2, 34].
43 휔 푡푓 = 푛 (3-17) 푠 + 휔푛
In an ideal system, the actual rate of the aircraft equals the commanded control rate as seen in Ref.
[2]. To account for errors due to a non-ideal system, the command filtered control rate error passes through a feedback compensator. Finally, feedback linearization removes non-linear effects and decouples the three control axes. The high-level block diagram of the closed-loop response to pilot lateral cyclic control deflection is shown below as explained in References [2, 34].
Figure 3-1. High-level block diagram of the controller’s response to pilot lateral cyclic control deflection.
The closed-loop response of the non-ideal system can be approximated by multiplying the ideal response by a phase delay, shown in Equation 3-18.
훷 휔 (3-18) ≈ 푛 푒−푇푠 푃푐푚푑 푠(푠 + 휔푛)
At this point it is helpful to look at a bode plot, given in Figure 3-2, for further explanation.
44
Figure 3-2. Bode plot of the system response.
Phase margin bandwidth is defined in References [20, 23] as the frequency where the phase is equal to -135°. There is no time delay for an ideal system (푇 = 0). In this case the bandwidth equals the command filter frequency (ωn) and the phase approaches -180° at high frequencies. For a non- ideal system, T is some nonzero number that results in a bandwidth less than that of the ideal system. The phase drops off at the non-ideal bandwidth frequency, causing the phase to fall below
-180° at high frequencies. Since bandwidth is commonly used to describe the response behavior of systems, manipulation of the command filter is expected to impact the rotorcraft’s flight handling qualities as seen in Ref. [23].
The compound rotorcraft features redundant control effectors in addition to traditional helicopter control effectors. Roll control is provided by a combination of cyclic pitch and wing flaperons (deployed differentially). Flaperons are deployed symmetrically as a high-lift or spoiler device. Pitch control is provided by cyclic pitch and a stabilator. Finally, forward propulsion is provided by collective pitch (main rotor) and propeller pitch (auxiliary propulsion system). A weighting parameter optimizes the allocation between traditional and redundant controls to minimize power required. Details of this allocation scheme are presented in Ref. [21].
45 Trim Analysis
As mentioned previously, the controller design is based on linear models of the rotorcraft flight dynamics (GENHEL-PSU) at various operating points in trim. The rotorcraft forces and moments are perturbed from trim in order to define stability and control derivatives for the linear models. Compound rotorcraft do not have a unique trim solution because they have more controls available than body axes. The designer can assist the control selection by manually scheduling propeller pitch as a function of airspeed. In general, rotorcraft collective trim follows a bucket shape, where increasing airspeed initially requires less collective trim due to the decrease in induced inflow and therefore power required. Increasing airspeed past the point of minimum power requires more collective trim because of the increase in profile drag. The baseline compound rotorcraft used in this study produces a similar shape, shown in Figure 3-3.
Figure 3-3. Collective trim allocation of the baseline model in trimmed forward flight.
The rotorcraft uses about 78% collective trim to maintain hover. As airspeed increases, the collective trim decreases to approximately 43% at 80 kts and then increases to approximately 62% at 200 kts. These results are consistent with trim theory as seen in Ref. [23].
46 The generic compound is designed to reduce power in forward flight, and as a result the auxiliary propulsion system is equipped with variable propeller pitch to efficiently produce propulsion force. Since power required is a function of airspeed, propeller pitch is scheduled within the model at various operating points in trim. Figure 3-4 shows the relationship between power required and propeller pitch at various airspeeds.
Figure 3-4. Power required as a function of propeller pitch for various airspeeds.
Increasing airspeed decreases the magnitude of propeller pitch required to minimize power required. The resulting curves for lateral cyclic, longitudinal cyclic, and collective are shown in
Figure 3-5. This analysis provides the standard schedule for the rotorcraft controls in trimmed forward flight. The linear models are extracted around these equlibirum conditions for use in the
NDLI controller design discussed above.
47
Figure 3-5. Pitch attitude, lateral cyclic, longitudinal cyclic, and collective required for minimum power required at various airspeeds.
Penn State Flight Simulator
Experimental tests were conducted by a government pilot and an industry pilot in 2017 at the PSU Vertical Lift Research Center of Excellence (VLRCOE) flight simulator facility. The
NLDI controller is compiled and linked to the GENHEL compound rotorcraft model to model the flight dynamics of a generic compound (described earlier in this chapter and explained in Ref. [2]).
GENHEL links directly to X-Plane visual software, which is projected onto a 170° field of view
(15 ft diameter) cylindrical screen by three ceiling-mounted projectors. A Heads-Up-Display
48 (HUD) is also projected onto the screen to provide pilots with added cues (torque, load factor, airspeed, and altitude) in place of the full visual and motion cueing of an actual aircraft. A visual of the screen and the cab is provided in Figure 3-6, while the HUD is provided in Figure 3-7.
Figure 3-6. Penn State flight simulator facility cab and visual display.
Throttle Torque G meter
VSI
DME
Airspeed Altitude Tape Tape
Figure 3-7. HUD provided to pilots in the Penn State flight simulator facility.
49 The simulator cab is built around the cockpit of a Bell XV-15 tiltrotor and is equipped with custom instrumentations and displays, along with a center cyclic stick, collective lever, and pedals. A throttle inceptor is also included on the center console so that the pilot can switch between the main rotor and the auxiliary propeller to provide thrust (forward and reverse). The control inceptors are linked to a control loading system to provide representative force-feel of a UH-60 Black Hawk.
50
Chapter 4
Handling Qualities Analysis of Baseline Configuration
Predictive and experimental flight handling qualities analysis has the potential to significantly improve rotorcraft design by providing a realistic perspective. It enables researchers to develop practical relationships between controller design, configuration design, and the resulting aircraft stability and control. This chapter studies the impact of controller design and redundant control allocation on the handling qualities of a generic compound rotorcraft. A total of 45 configurations are tested for roll-axis and pitch-axis handling qualities using predictive methods described in Chapter 2. Several configurations are also evaluated in piloted simulations at the
Pennsylvania State University (PSU) VLRCOE flight simulator facility. These configurations will be explained in the following sections.
Command Filters
The command filter (Equation 3-17) of the NLDI controller defines the ideal response of the aircraft to pilot input, where the command filter parameter 휔푛 is related to the bandwidth frequency of the system and is equivalent to the attitude damping. As a result, variation of 휔푛 is predicted to significantly impact the overall flight handling qualities of the rotorcraft. Previous studies at PSU found that optimal command filter parameters for the roll-axis and pitch-axis of the
GENHEL-PSU compound rotorcraft are:
휔푛훷 = 3.75 푟푎푑/푠푒푐
휔푛훳 = 2.50 푟푎푑/푠푒푐.
51 The configuration with these command filter parameters is referred to as the baseline. Decreasing the parameter theoretically leads to degradation in bandwidth, while increasing the parameter theoretically leads to a ‘better than ideal’ controller. However, realistic problems such as control saturation might lead even the ‘improved’ controller to experience poor performance in handling qualities. This section explores the effect of the command filter parameter on the roll-axis and pitch- axis handling qualities of numerical simulation, along with validation using piloted simulation. In total, 9 variations of the command filter parameter are tested, shown in Table 4-1.
Table 4-1. Configurations involved in the study of the command filter. Configuration Type Axis 휔푛훷 휔푛훳 Baseline Roll & Pitch 3.75 2.50 Very low BW (roll) Roll 1.25 2.50 Low BW (roll) Roll 2.50 2.50 High BW (roll) Roll 5.00 2.50 Very high BW (roll) Roll 6.25 2.50 Very low BW (pitch) Pitch 3.75 0.50 Low BW (pitch) Pitch 3.75 1.25 High BW (pitch) Pitch 3.75 3.75 Very High BW (pitch) Pitch 3.75 5.00
Roll Command Filter
The roll-axis is the primary focus in this rotorcraft high-speed handling qualities analysis due to its simplicity. Predictive methods include bandwidth/phase delay, cross-coupling, and attitude quickness, while piloted simulation methods include Break Turn, Roll SOS Tracking
(RSOS), and Bank Attitude Capture and Hold (BACH), as discussed in Chapter 2. Together, these methods are used to define desired (Level 1), adequate (Level 2) and degraded (Level 3) handling qualities for configurations at various stages in the design process.
The configurations used in this study involve variation of the roll-axis command filter
parameter 휔푛훷. In total, fives configurations are investigated, each using the baseline aircraft
52 configuration and optimal pitch-axis command filter parameter (휔푛훳 = 2.5 rad/sec). The baseline
configuration also uses the optimal roll-axis command filter parameter (휔푛훷 = 3.75 rad/sec). Two
configurations are considered to have very low and low bandwidth (휔푛훷 = 1.25, 2.50 rad/sec),
while the final two configurations have high and very high bandwidth (휔푛훷 = 5.00, 6.25 rad/sec).
The configurations are given in Table 4-2.
Table 4-2. Configurations involved in the study of the roll-axis command filter. Configuration Type Axis 휔푛훷 휔푛훳 Baseline Roll & Pitch 3.75 2.50 Very low BW (roll) Roll 1.25 2.50 Low BW (roll) Roll 2.50 2.50 High BW (roll) Roll 5.00 2.50 Very high BW (roll) Roll 6.25 2.50
Chapter 2 explains that flight handling qualities can be divided into two categories: (1) small-amplitude/moderate to high frequency response characteristics and (2) moderate to large ampltude/low to moderate frequency response characteristics. ADS-33E-PRF recommends that the first category be evaluated by the Bandwidth/Phase Delay and Cross-coupling tests, which involve analsis of the linearized system’s frequency response of attitude to the pilot’s control stick. It is expected that the results from the Bandwidth/Phase Delay test will match the earlier prediction that bandwidth is a function of the command filter, resulting in improved handling qualities for higher bandwidth configurations. The results of this study are shown in Figure 4-1.
53
Figure 4-1.Bandwidth analysis for configurations of various roll command filter parameters 휔푛훷.
The results of the Bandwidth/Phase Delay test show that variation in 휔푛훷 only significantly
influences the roll-axis. The results show that variation of 휔푛훷 drastically impacts the handling qualities, where a roll-axis command filter parameter of 1.25 rad/sec actually puts the configuration
in the degraded Level 3 category. As expected, increasing 휔푛훷 leads to a system with a higher
bandwidth and improved handling qualities. However, increasing 휔푛훷 past the optimal value of
3.75 rad/sec appears to also increase the phase delay slightly. If more tests were conducted, the
resulting trend might show that a system with a very large 휔푛훷 could possibly lead to a phase delay that degrades the handling qualities to Level 2. As a result, the earlier assumption that a higher bandwidth system automatically has better handling qualities is false. Instead, increasing the command filter could lead to problems in controller response, such as in an underdamped system.
Additionally, all configurations are within the boundaries of desired Level 1 handling qualities for
the pitch-axis, proving that the optimal 휔푛훳 that was selected in previous research studies is appropriate for the generic compound used in this study.
Investigation of the cross-coupling characteristics between the roll-axis and pitch-axis for
various 휔푛훷, shown in Figure 4-2, reveals a similar problem at large 휔푛훷.
54
Figure 4-2. Cross-coupling analysis for configurations of various 휔푛훷.
Variation of 휔푛훷 does not appear to impact the pitch-due-to-roll (푞/푝) cross-coupling, but it
significantly impacts the roll-due-to-pitch (푝/푞). Figure 4-2 shows that increasing 휔푛훷 also increases 푝/푞 and decreases the spacing between configurations. At first glance this result is
unexpected; roll-axis command filter parameter 휔푛훷 is expected to only affect responses due to lateral stick input. Equations 4-1 and 4-2 show the equations for roll-due-to-pitch coupling and pitch-due-to-roll coupling.
푝/훿 (4-1) 푝/푞 = 푙표푛 = 푟표푙푙 푑푢푒 푡표 푝𝑖푡푐ℎ 푐표푢푝푙𝑖푛푔 푞/훿푙표푛
푞/훿 (4-2) 푞/푝 = 푙푎푡 = 푝𝑖푡푐ℎ 푑푢푒 푡표 푟표푙푙 푐표푢푝푙𝑖푛푔 푝/훿푙푎푡
Therefore, it is expected that variation of 휔푛훷 will affect pitch-due-to-roll (푞/푝) coupling and not roll-due-to-pitch (푝/푞) coupling. To investigate this further, 푝/훿푙표푛 and 푞/훿푙표푛 are compared for
configurations with various 휔푛훷.
55
Figure 4-3. Explanation of Cross-Coupling Analysis results.
The plots show that the roll command filter parameter 휔푛훷 does not affect 푝/훿푙표푛 or 푞/훿푙표푛, which is consistent the expectation. However, the Cross-Coupling specification in ADS-33E-PRF evaluates 푝/푞 at frequencies between 휔퐵푊 and 휔180 for the roll axis (훷/훿푙푎푡). The solid black
lines in the bottom plot represent 휔퐵푊 and 휔180 for the configuration with 휔푛훷 = 2.50 푟푎푑/푠푒푐,
while the dashed black lines represent these frequencies for the configuration with 휔푛훷 =
5.00 푟푎푑/푠푒푐. Clearly these frequencies increase with higher 휔푛훷. Magnitudes of 푝/훿푙표푛 generally
increase with 휔푛훷 in this frequency range while magnitudes of 푞/훿푙표푛 decrease with a normal roll
56 ( ) ( ) off. Therefore, the ratio of 푝/훿푙표푛 / 푞/훿푙표푛 increases significantly with higher 휔푛훷. This analysis confirms the results from Figure 4-2.
Although all configurations technically fall into the desired Level 1 category, if the criteria were expanded to higher magnitudes of 푝/푞, it would appear that the configuration with the highest
휔푛훷 (6.25 rad/sec) would fall into adequate Level 2. This implies that higher bandwidth systems would most likely have worse handling qualities in high-precision maneuvers such as in target tracking. Unfortunately, ADS-33E-PRF does not define criteria for larger magnitudes of 푝/푞, so it
is unclear whether large 휔푛훷 would fall into Level 1 or worse. Hopefully these criteria will be refined as more flight test data becomes available.
The flight handling qualities of the generic compound in moderate to large amplitude/low to moderate frequency response is analyzed through the Roll Attitude Quickness (RAQ) test. The
baseline configuration along with the two configurations with degraded handling qualities (휔푛훷 =
1.25, 2.50) are evaluated for RAQ at a trimmed airspeed of 160 kts, which is the ideal airspeed for many ADS-33E-PRF Mission Task Elements (MTEs). Each configuration is tested for max Rate-
Command/Attitude Hold (RCAH) input using triangular commands of various lengths (0.4 sec, 0.6 sec, 0.8 sec, and 1.0 sec). The results are shown in Figure 4-4.
57
Figure 4-4. Roll Attitude Quickness analysis for configurations of various 휔푛훷.
As expected, increasing the length of the input command increases the resulting change in bank angle. However, as amplitude increases nonlinear effects begin to impact the system, resulting in a negative slope. The baseline configuration along with the configuration with moderately low
bandwidth (휔푛훷 = 2.50 푟푎푑/푠푒푐) are within the desired Level 1 category, while the configuration
with very low bandwidth (휔푛훷 = 1.25 푟푎푑/푠푒푐) is in the adequate Level 2 category. This shows that the phase delay prevents the very low bandwidth configuration from responding, essentially creating an over-damped system. At larger input lengths, the handling qualities degrade for both the baseline configuration and the moderately low bandwidth configuration, leading to an intersection with the adequate Level 2 boundary. One explanation could be control saturation. To investigate this further, the lateral cyclic commands XA for all configurations are compared in
Figure 4-5.
58
Figure 4-5. Lateral cyclic command during the RAQ test (signal length of 1.0 sec) for various
휔푛훷.
The controllers of the two configurations with the highest 휔푛훷 reach the physical limit of the lateral cyclic command during the RAQ test. As a result, the pilot cannot give the desired command, and the aircraft handling qualities degrade.
Another area of interest is the impact of airspeed on handling qualities. At higher flight speeds, it is expected that the aircraft will become more difficult to control because of limited precision in control actuation (aircraft becomes more sensitive to control effector displacement at higher forward flight speeds). To test this theory, the configurations are also evaluated using RAQ at airspeeds of 140 and 180 kts. Results are shown in Figure 4-6.
59
Figure 4-6. Roll Attitude Quickness results for configurations of various airspeeds and 휔푛훷.
As airspeed increases the curves shift downward. This means that the handling qualities of all configurations degrade as airspeed increases. The results are consistent with the earlier prediction.
Overall, the predictive analysis reveals that the baseline configuration has the best handling qualities. The configurations with lower bandwidth perform poorly, typically falling within Level
2 or even Level 3. Although cross-coupling is not a big problem, it appears that configurations with very large bandwidths might have slightly degraded handling qualities if the criteria were refined.
Experimental tests conducted by two pilots in the PSU VLRCOE flight simulator facility in 2017
are used to verify these results. Four of the configurations (휔푛훷 = 1.25, 2.50, 3.75, 5.00 푟푎푑/푠푒푐) are investigated for flight handling qualities using the Break Turn, RSOS, and BACH MTEs.
As discussed in Chapter 2, the Break Turn MTE requires that a pilot fly in trimmed steady- level flight at 160 kts and then complete a controlled 90° heading change with 60° bank angle, finally ending again in trimmed steady-level flight (explained in Ref. [26]) The goal is to complete the maneuver as fast as possible without significant altitude or airspeed changes. The aircraft flight handling qualities are judged based on desired performance criteria and pilot subjective opinion, resulting in an HQR score between 1 and 9. This MTE is most closely related to the RAQ predictive
60 assessment because it evaluates the aircraft’s handling qualities for aggressive, non-precision situations. Results for left and right Break Turn tests are shown in Figure 4-7.
Figure 4-7. Break Turn MTE results for configurations of various airspeeds and 휔푛훷.
The results show a general trend where increasing 휔푛훷 leads to a lower HQR score (better handling qualities). In general, the pilots noted that the Right Break Turn felt easier, and therefore the configurations received better scores than with the Left Break Turn. The configurations with
degraded bandwidth (휔푛훷 = 1.25, 2.50 푟푎푑/푠푒푐) required considerable pilot compensation to prevent heading overshoot, leading to less aggressive performance and slower completion times.
However, the altitude and airspeed typically stayed within the desired criteria. The pilots gave the
baseline configuration (휔푛훷 = 3.75 푟푎푑/푠푒푐) the lowest HQR (best handling qualities), noting that its performance was much cleaner all around. However, the pilots noted that the aircraft tended to climb and bleed off speed during the turn, which resulted in careful maneuvering and moderate, but tolerable, pilot workload. The MTE also required moderate pilot compensation to prevent heading overshoot, which led to some loss of precision in the initial bank angle capture. The
configuration with high bandwidth (휔푛훷 = 5.00 푟푎푑/푠푒푐) received a higher (worse) HQR score
61 than the baseline aircraft for left Break Turn. Pilot 1 noted more difficulty in capturing the initial bank angle and maintaining airspeed, leading to higher pilot workload than the baseline configuration. However, the pilot still met the criteria for desired performance, and therefore the configuration received a lower (better) HQR than the configurations with degraded handling qualities. These observations are consistent with the results from the predictive analysis, which
generally showed better performance for higher 휔푛훷 and some handling qualities deficiencies for
configurations with very high 휔푛훷.
The next experimental test is the RSOS MTE, which requires the pilot to fly in trimmed steady-level flight at 160 kts and then track on-screen cues by modifying the aircraft’s attitude as quickly and as accurately as possible. The aircraft handling qualities are judged entirely on the pilot’s opinion of whether he made the desired performance criteria. This MTE is most closely related to the Bandwidth/Phase Delay and Cross-coupling predictive assessments since it involves high-precision tracking for small/moderate amplitude changes in bank angle. Results for the RSOS
MTE are shown in Figure 4-8.
Figure 4-8. RSOS MTE results for configurations of various 휔푛훷.
62
The results show a similar trend to the Break Turn maneuver where increasing 휔푛훷 leads to a lower
HQR score (better handling qualities). The pilots reported that pitch-due-to-roll cross-coupling was distracting and led to moderate pilot compensation and control deficiencies. These problems were worse for the configurations of low bandwidth, leading to large tracking errors and unusual control strategies to compensate for slower response from the controller. The pilots reported improved handling qualities for the baseline configuration and noted that the deficiencies were minor, but annoying. The configuration with high bandwidth received the best score. The pilots noted that control felt tight, allowing for a more aggressive strategy with reduced pilot workload. Although deficiencies were still present, Pilot 1 gave the high bandwidth configuration an HQR 3, which corresponds to desired Level 1 handling qualities. These observations are similar to that of the
Bandwidth/Phase Delay predicted analysis, showing that high bandwidth configurations have better handling qualities than low bandwidth configurations. However, the earlier results reported that higher bandwidth configurations might actually have degraded handling qualities due to control saturation. This was not an issue with the piloted simulations of the RSOS MTE.
Additionally, the cross-coupling predicted analysis showed that increasing the roll-axis bandwidth increases the roll-due-to-pitch cross-coupling. The results were the opposite with the piloted simulation, where increasing the bandwidth did not change the roll-due-to-pitch, but decreased the pitch-due-to-roll. Future studies should examine this disparity to conclude if the ADS-33E-PRF small amplitude/moderate to high frequency predictive methods are truly representative of the handling qualities of compounded configurations in high-speed flight.
The final experimental test is the BACH MTE, which is similar to RSOS in that they both require high-precision tracking. The main difference is that the BACH MTE requires the pilot to capture and hold an attitude. Again, this MTE is most closely related to the Bandwidth/Phase Delay and Cross-coupling MTEs since it concerns small amplitude attitude changes. Results for the
BACH MTE are shown in Figure 4-9.
63
Figure 4-9. BACH MTE results for configurations of various 휔푛훷.
The results show a similar trend to the Break Turn and RSOS maneuvers where increasing 휔푛훷 leads to a lower HQR score (better handling qualities). The pilots experienced significant problems with attitude overshoot and pitch-due-to-roll cross-coupling while testing the configuration with the lowest bandwidth. Both pilots reported extensive pilot compensation, adopting the control strategy to lead and then apply a lateral reversal to capture the bank angle. The pilot also noted high
workload due to cross-coupling. Increasing 휔푛훷 to 2.50 rad/sec led to much better HQR scores.
Although one pilot said that the aircraft was crisp in roll capture, both pilots agreed that cross- coupling was still an issue. The baseline configuration received only slightly improved HQR scores.
Pilots noted that while the bank angle capture was easy, maintaining off-axis control led to less precision and moderate pilot compensation. The configuration with high bandwidth also performed well with the bank angle capture. One pilot noted that roll response was crisp, while the other pilot described the response as a little less predictable. Cross-coupling was still an issue with this configuration. Although these results are similar to those of the Bandwidth/Phase Delay, it seems as though the predicted analysis does not provide an accurate representation of the cross-coupling characteristics of the aircraft. The predicted analysis showed that all configurations were within the
64 desired Level 1 criteria, but the pilots reported that cross-coupling was the main problem that limited the HQR scores. This discrepancy might be a result of poorly defined boundaries of the predicted analysis. As mentioned with the RSOS results, it appears that the ADS-33E-PRF criteria might not be appropriate for compound rotorcraft in high-speed flight. Future research should investigate this further.
Pitch Command Filter
Although ADS-33E-PRF includes some tests for the pitch-axis, they focus mainly on handling qualities in low-speed flight and hover. Since bandwidth/phase delay and cross-coupling parameters consider the entire flight regime, they are included in this analysis. However, ADS-
33E-PRF does not contain a test for pitch-attitude quickness in forward flight. As a result, the pitch- axis analysis is supplemented with an attitude quickness test commonly used for fixed-wing aircraft in forward flight. Experimental tests include Acceleration/Deceleration (Accel/Decel), Pitch SOS
Tracking (PSOS), and Pitch Attitude Capture and Hold (PACH), as discussed in Chapter 2.
Five configurations are tested using the baseline roll command filter parameter (3.75
rad/sec) and a variation of the pitch command filter parameter 휔푛훳 (0.50, 1.25, 2.50, 3.75, or 5.00
rad/sec), where the baseline 휔푛훳 is 2.50 rad/sec. These configurations are given in Table 4-3.
Table 4-3. Configurations involved in the study of the pitch-axis command filter. Configuration Type Axis 휔푛훷 휔푛훳 Baseline Roll & Pitch 3.75 2.50 Very low BW (pitch) Pitch 3.75 0.50 Low BW (pitch) Pitch 3.75 1.25 High BW (pitch) Pitch 3.75 3.75 Very High BW (pitch) Pitch 3.75 5.00
65 The stabilator is not used in this simulation, so the only difference between configurations is the pitch command filter. As with the roll command filter analysis, the pitch command filter analysis is divided into two categories: (1) small amplitude/moderate to high frequency response characteristics and (2) moderate to large amplitude/low to moderate frequency response characteristics. The first category is tested using the Bandwidth/Phase Delay and Cross-coupling tests. It is expected that the results from this study will match the earlier prediction that increased bandwidth leads to improved pitch-axis handling qualities. Results from the Bandwidth analysis are given in Figure 4-10.
Figure 4-10. Bandwidth analysis for configurations of various pitch command filter parameter
휔푛훳.
As expected, varying the pitch command filter parameter has no effect on the roll bandwidth analysis, indicating that there is not an identifiable cross-coupling in the bandwidth or phase delay for small amplitude/moderate to high frequency attitude changes, such as in precision tracking. All configurations are within the boundaries of Level 1 for the Roll Bandwidth Analysis,
proving that the ideal roll command filter parameter (휔푛훷 = 3.75 푟푎푑/푠푒푐) produces desired
handling qualities. The Pitch Bandwidth Analysis shows that increasing 휔푛훳 improves the pitch-
66 axis handling qualities from adequate Level 2 to desired Level 1. Additionally, the trend appears to
curve as 휔푛훳 increases, reaching a maximum at the ideal pitch command filter parameter (2.50
rad/sec) and then curving downward as 휔푛훳 increases further. This implies that the phase delay is
actually lower for the configurations with degraded handling qualities (휔푛훳 = 0.50, 1.25) than for
the ideal response. Additionally, increasing 휔푛훳 past the ideal parameter leads to another decrease in phase delay accompanied by an increase in bandwidth. It is expected that extending the trend line would show continued improvement in pitch-axis handling qualities as a result of increasing
휔푛훳. However, ADS-33E-PRF does not define criteria for handling qualities that surpass the
standard for Level 1, so it is unclear if having a very large 휔푛훳 would give the expected Level 1 handling qualities.
Investigation of the cross-coupling characteristics reveals an interesting trend, shown in
Figure 4-11.
Figure 4-11. Cross-coupling analysis for configurations of various 휔푛훳.
The baseline configuration produces the best handling qualities in terms of cross-coupling, while
increasing or decreasing 휔푛훳 degrades the handling qualities, pushing the score from desired Level
67
1 towards the boundary for adequate Level 2. Variation in 휔푛훳 changes pitch-due-to-roll (푞/푝) cross-coupling, but does not impact roll-due-to-pitch (푝/푞) cross-coupling. Unlike the results from
variation of 휔푛훷, decreasing 휔푛훳 actually degrades handling qualities. Another interesting
observation is that cross-coupling handling qualities appear to degrade at high 휔푛훳 although the bandwidth analysis gave no indication of problems.
The last predictive method, the Pitch Attitude Dropback (PAQ) test, studies the moderate to large amplitude/low to moderate frequency response characteristics of the aircraft. This method is not found in ADS-33E-PRF for forward flight. Similar to the RAQ test, the baseline
configuration, along with the two configurations with degraded handling qualities (휔푛훳 =
0.50, 1.25) are tested at a trimmed airspeed of 160 kts for various commanded pitch rates, given by non-dimensional longitudinal input (0.075, 0.1, 0.125, and 0.15) where -1 means full forward and 1 means full aft. The pitch attitude dropback (DB) and the peak rate of change 푞푝푘 are compared in the specification, shown in Figure 4-12.
Figure 4-12. Pitch Attitude Dropback analysis for configurations of various 휔푛훳.
68 Although there appears to be some variation, all configurations are well within the boundary for desired Level 1 handling qualities. This means that the aircraft should respond as desired to large amplitude changes in pitch. PAQ results for the configurations at airspeeds of 140 and 180 kts, shown in Figure 4-13, are also well within Level 1.
Figure 4-13. Pitch Attitude Dropback analysis for configurations of various airspeeds and ωnϴ.
However, these results are suspicious considering the compound rotorcraft will have high profile drag at high airspeeds. Future research should refine this method and develop other PAQ tests specifically to analyze the handling qualities of rotorcraft in forward flight.
Overall, the predictive analysis revealed that the baseline configuration has the best
handling qualities. Similar to the results from variation in 휔푛훷, the handling qualities of the
configurations with low 휔푛훳 were degraded. The results of the Bandwidth/Phase Delay test showed
that increasing 휔푛훳 leads to improved handling qualities. However, increasing 휔푛훳 past that of the baseline configuration actually degrades the Cross-coupling handling qualities. To study the pitch
command filter further, three of the configurations (휔푛훳 = 0.50, 1.25, 2.50 푟푎푑/푠푒푐) were tested in the Penn State flight simulator in 2017 by two experienced government pilots for the PSOS and
69
PACH MTEs. Two of the configurations (휔푛훳 = 0.50, 2.50 푟푎푑/푠푒푐) are also tested for the High
Speed Acceleration MTE.
Similar to the RSOS and BACH MTEs, the PSOS and PACH MTEs require the pilot to fly in trimmed steady-level flight at 160 kts and then follow on-screen cues by modifying the aircraft’s attitude as quickly and as accurately as possible. The pilots rate the handling qualities of each configuration using the Cooper-Harper Handing Qualities Rating (HQR) Scale. These MTEs are most closely related to the Bandwidth/Phase Delay and Cross-coupling predictive assessments since they involve high-precision tracking. Results for the PSOS and PACH MTEs are shown in
Figure 4-14.
Figure 4-14. Results from the PSOS (left) and PACH (right) MTEs for configurations of various
ωnϴ.
The results show similar trends to the RSOS and BACH MTEs, where increasing 휔푛훳 leads to lower HQR scores (better handling qualities). However, there is a large disparity between the pilots’ scores for the PSOS MTE. The trend is more pronounced in the scores from Pilot 1, while Pilot 2 gave all configurations a Level 2 rating, where the baseline configuration performed slightly better
70 than the rest. Future research should incorporate more pilot evaluations to verify that a trend actually exists. On the other hand, the results from the PACH MTE are fairly consistent between the pilots. Although more research is required to draw concrete conclusions, it appears that the
PACH MTE might be better suited to assess the handling qualities for low to moderate amplitude/high frequency maneuvers.
The results from the PSOS MTE are fairly consistent with the Bandwidth/Phase Delay analysis. Both pilots did not meet the criteria of desired performance for the lowest bandwidth configuration. Pilot 1 noted that the aircraft was difficult to control even with maximum pilot compensation, and Pilot 2 simply stated that he knew he could not reach the desired criteria.
Increasing 휔푛훳 to 1.25 rad/sec drastically improved the HQR from Pilot 1, who mentioned that the workload was tolerable and that there were no undesirable motions. The score from Pilot 2 did not improve and neither did his performance; the pilot was within the desired criteria 42% of the time for both configurations. The scores from both pilots improved for the baseline configuration, where
Pilot 1 noted only small problems with overshoot and minimal compensation. Lastly, the high bandwidth configuration was only flown by Pilot 2, who gave it a worse score than that of the baseline configuration due to oscillatory behavior as a result of low damping. Unfortunately, it is difficult to draw conclusions from the results of one pilot, so it is unclear if the handling qualities actually degraded for the high bandwidth configuration. In general, this MTE needs more refinement. The large disparity between pilot scores shows that this MTE might actually test pilot coping ability rather than aircraft handling qualities. Additionally, the pilots agreed that cross- coupling was not a problem in pitch for any configuration. However, results were consistent with
the earlier prediction that increasing 휔푛훳 past 2.50 rad/sec degrades the handling qualities.
Although all configurations were within Level 1 for the Cross-coupling predicted analysis, it is surprising that the pilots did not notice a difference between configurations.
71 Results from the PACH MTE are fairly consistent between Pilot 1 and Pilot 2. There is a
clear trend that increasing 휔푛훳 to 2.50 rad/sec improves handling qualities, but increasing 휔푛훳 past
2.50 rad/sec degrades handling qualities. The pilots noted that the configuration with very low bandwidth was controllable, but oscillatory behavior led to high pilot workload. The oscillatory behavior decreased for the low bandwidth configuration, although one pilot noted that this maneuver still required pilot compensation. The baseline configuration performed much better; both pilots agreed that there were only small overshoots, and overall the aircraft felt crisper. Only one pilot scored the high bandwidth configuration, and although the max error and overshoot were less than that of the baseline, the pilot gave this configuration a worse score. The pilots did not mention problems with cross-coupling for any configuration, although the trend from the HQR scores matches the results from the Cross-coupling predicted analysis. Since the pilots did not notice any cross-coupling characteristics in both the PSOS and PACH MTEs, it is recommended that future research efforts focus on developing MTEs that better test for cross-coupling issues.
The final experimental test is the High Speed Acceleration-Deceleration MTE, which was only tested by the industry pilot. The pilot only completed the acceleration portion of this MTE, which requires the pilot to begin in trimmed steady-level forward flight at 50 kts and then accelerate at a constant altitude until (푉퐻 − 10) 푘푡푠 is reached, where 푉퐻 equals 160 kts for the generic compound. The goal is to complete the maneuver as quickly as possible. This MTE is most closely related to the Bandwidth/Phase Delay and Cross-coupling analyses since small-amplitude attitude commands are required to hold the altitude. This MTE also checks for undesired cross-coupling between the roll-axis and pitch-axis. Results from this test are shown in Figure 4-15.
72
Figure 4-15. High Speed Acceleration MTE results for configurations of various 휔푛훳.
Although only two data points are available, the trend is consistent with the results from the other
tests, where increasing 휔푛훳 improves handling qualities. The pilot noted difficulty in maintaining the speed capture and altitude with the low bandwidth configuration, ultimately leading to increased workload and HQR 5. The handling qualities of the baseline configuration were better; the pilot was able to keep all parameters within desired tolerances with moderate pilot compensation.
However, the pilot noted some undesired behavior in the roll-axis due to large power charges. It is difficult to draw concrete conclusions from this study, but in general the results agree with the trend from earlier analyses; increasing bandwidth improves the handling qualities. An interesting aspect of these results is that cross-coupling seems to impact the pilot workload. It is recommended that future research investigate the relationship between cross-coupling and handling qualities in this
MTE and perhaps try to incorporate it into other MTEs to better assess precision tracking response characteristics. Additionally, an MTE should be developed to evaluate moderate to large amplitude/low to moderate frequency response characteristics.
73 Flaperon/Stabilator Gearing
The next control parameter of interest is the gearing of the redundant controls, specifically flaperons in terms of roll control and a stabilator in terms of pitch control. The gearing is modified using a parameter 휔푑, which represents the fraction of control authority on a scale of 0 (no control authority) to 1 (complete control authority). Thus far, tests have not included flaperons and the stabilator, and it is expected that they will improve handling qualities in forward flight. As with the command filter study, predictive tests are conducted for the 9 variations in command filter parameter 휔푛 (Table 4-1). In addition, each configuration is tested for 5 variations in redundant control allocation 휔푑 (weighting parameter), shown in Table 4-5.
Table 4-4. Configurations involved in the study of redundant control gearing.
Configuration Type Axis 휔푑푓푙푎푝 휔푑푠푡푎푏 Baseline Roll & Pitch 0 0 Some flap allocation Roll 0.375 0 High flap allocation Roll 0.750 0 Some stab allocation Pitch 0 0.375 High stab allocation Pitch 0 0.750
Flaperons primarily influence the roll-axis, while the stabilator primarily influences the pitch-axis. As a result, it is expected that variation in gearing of the redundant controls will only impact the respective control axis. Results for the Bandwidth/Phase Delay test are shown in Figure
4-16.
74
Figure 4-16. Bandwidth/Phase Delay analysis for configurations of various flaperon allocations 휔푑푓푙푎푝 (top) and configurations of various stabilator allocations 휔푑푠푡푎푏 (bottom).
In general, the results are consistent with the earlier expectation that variation in flaperon gearing only impacts the Roll Bandwidth Analysis and that variation in stabilator gearing only impacts the
Pitch Bandwidth Analysis. The baseline is within the boundaries of desired Level 1 handling qualities, showing that the flaperons and stabilator are not necessary for small-amplitude/high frequency changes. However, the results for the Roll Bandwidth Analysis indicate that some flaperon authority improves handling qualities (pushes the score further into the desired Level 1 category), while a large amount of authority might actually degrade handling qualities. If the trend line were to continue, flaperon allocation larger than 0.75 might result in handling qualities that are
75 closer to the adequate Level 2 category. Note, the discontinuity in the plot for results from 휔푛훷 =
5.00, 6.25 푟푎푑/푠푒푐 is a result of using a two point straight line method to measure slope of the phase response which was highly nonlinear for this case. A linear least-squares curve fit is more suitable for nonlinear phase curves. The other trend from the Bandwidth/Phase Delay Analysis
shows that increasing stabilator authority improves pitch-axis handling qualities for 휔푛훳 = 2.50 or higher. For low bandwidth configuration, the slope appears vertical, where variation in stabilator
authority only impacts the phase delay. As 휔푛훳 increases, the slope decreases. It is expected that variation in stabilator authority will have a small impact on phase delay for very high bandwidth configurations. This trend is interesting and seems to indicate that the purpose of a stabilator changes based on the inherent bandwidth of the controller.
The effect of redundant control allocation is further explored using the cross-coupling test; results are shown in Figure 4-17.
Figure 4-17. Cross-coupling analysis for configurations of various flaperon allocations 휔푑푓푙푎푝 (left) and configurations of various stabilator allocations 휔푑푠푡푎푏 (right).
76
Increasing the flaperon allocation 휔푑푓푙푎푝 also increases the roll-due-to-pitch (푝/푞), which might
actually place the configuration with high 휔푛훷 and high 휔푑푓푙푎푝 in the adequate Level 2 category if the boundaries were refined. Stabilator authority has the opposite effect. As stabilator allocation
휔푑푠푡푎푏 increases, handling qualities improve. Although all configurations are within the desired
Level 1 category, the addition of stabilator authority pushes the configuration further away from the boundary. As a result, it is expected that a configuration with low flaperon allocation and high stabilator allocation will have desired cross-coupling characteristics.
Finally, attitude quickness is evaluated for all configurations using RAQ for flaperon allocation and PAQ for stabilator allocation. The PAQ results showed very little variation, and therefore they are not included. The results from RAQ are shown in Figure 4-18.
Figure 4-18. Roll Attitude Quickness analysis for configurations of various flaperon allocations 휔푑푓푙푎푝.
As seen in the command filter analysis, increasing 휔푛훷 improves handling qualities for small attitude changes, while it does not seem to have a large effect on handling qualities for large attitude changes due to rotor control saturation (Figure 4-5). Rotor control saturation is recovered by allocating part of the control to the flaperons. Increasing flaperon allocation shows a similar trend
77 but with larger spacing between configurations. Therefore, the impact of increasing 휔푛훷 is greater with larger flaperon allocations. Additionally, the attitude where handling qualities converge
(where 휔푛훷 does not make a difference) is much higher for the configurations with high flaperon allocation. This might indicate that flaperon allocation actually delays the negative impact of nonlinearities in the system by providing more control authority. It is recommended that future research compare these results to those of piloted simulation using the Break Turn MTE.
In summary, the predictive analysis of redundant control allocation revealed that the addition of some flaperon/stabilator authority will improve the overall handling qualities for moderate to large amplitude/low to moderate frequency responses. However, the addition of flaperon authority might cause problems for small-amplitude/moderate to high frequency responses.
78
Chapter 5
Vehicle Design Modifications
Aircraft geometry is typically in flux at the early stages of design, and therefore it is important to understand the impact of configuration changes on handling qualities. One design parameter of interest is the location of the main rotor flapping hinge relative to the axis of rotation.
This distance is normally determined by physical constraints, and it impacts the moment created from cyclic pitch deflection. Another parameter of interest is the size of the redundant control surfaces and the aerodynamic surfaces associated with them, namely wing span and tail area.
Increasing wing span and tail area directly impact the size and/or moment arm of the redundant control effectors (wing flaperons and stabilator), and therefore it is expected to increase their effectiveness. The impact of design modifications on handling qualities using ADS-33E-PRF and proposed metrics is explored in this chapter.
Flapping Hinge Offset
Articulation of rotor blades allows each blade to respond independently to the dissymmetry of loads created in forward flight. Dynamic pressure differences on the advancing blade and retreating blade causes the blades to periodically flap up and down using a flapping hinge, which drastically reduces the load imbalance and thus makes the rotorcraft easier to control. The flapping hinge is usually located some distance away from the axis of rotation due to physical constraints from the rotor hub. Increasing flapping hinge offset (푒) increases the moment generated by the main rotor for a given amount of flapping (thus increases control effectiveness of cyclic pitch) and
79 increases the natural frequency (휈훽) of the flapping response given in References. [1, 23] by
Equation 5-1.
(5-1) 3푒 3 휈 = √1 + ≈ √1 + 푒 훽 2(1 − 푒) 2
Increasing 휈훽 tends to decrease the phase lag from the rotor, and as a result it tends to increase cross-coupling since the response will occur before the 90° lag between the roll axis and pitch axis, but also decreases the effective phase lag of the main rotor cyclic controls. Increasing 휈훽 tends to decrease the amount of flapping induced by cyclic pitch, but in terms of control effectiveness this is more than overcome by the larger moments produced by the main rotor with higher hinge offset.
The configurations in this study use a main rotor radius of 26.85 ft with various values of flapping hinge offset. The baseline configuration uses a flapping hinge offset of 1.80 ft, which is
6.7% of the main rotor radius. The last two configurations increase the hinge offset to 2.60 ft (9.7%) and 3.40 ft (12.7%). The configurations are given in Table 5-2. Only the predictive metrics are used to assess handling qualities for these configurations.
Table 5-1. Configurations involved in the study of flapping hinge offset. Configuration Type Hinge Offset (H.O., ft) Baseline 1.80 High Offset 2.60 Very High Offset 3.40
Since phase lag between cyclic input and flapping response is expected to decrease for configurations of higher hinge offset, it is expected that the results from the Bandwidth/Phase Delay test will show lower phase delay for these configurations. The determining factor for this test will be the impact of hinge offset on bandwidth frequency. Results for configurations of various
80 flapping hinge offsets are shown in Figure 5-1 for three flaperon settings (0, 0.375, and 0.75) and three stabilator settings (0, 0.375, and 0.75).
Figure 5-1. Bandwidth/Phase Delay analysis for configurations of various flapping hinge offsets using three flaperon allocations 휔푑푓푙푎푝 (top) and three stabilator allocations 휔푑푠푡푎푏 (bottom).
In general, the results are consistent with the results from Chapter 4 where variation in flaperon gearing primarily impacts the Roll Bandwidth Analysis and variation in stabilator gearing primarily impacts the Pitch Bandwidth Analysis. All three configurations are within the boundaries of desired
Level 1 handling qualities for all flaperon and stabilator settings. One interesting observation is that
81 increasing hinge offset also increases flaperon effectiveness, shown in the top left plot in Figure 5-
1. This is probably related to the trend mentioned earlier where higher hinge offset increases the moment generated by the main rotor for a given amount of cyclic pitch. Increasing hinge offset also increases the spread in bandwidth, and the configuration with highest hinge offset has and flaperon/stabilator allocation has the highest roll/pitch bandwidth. Another interesting observation is that increasing hinge offset decreases the roll phase delay, but increases the pitch phase delay.
The latter effect was unexpected and at this point the reasons are not clear. Finally, results from the off-axis (top right and bottom left) show that increasing hinge offset actually degrades the bandwidth slightly when no flaperon/stabilator allocation is used.
Next, small amplitude/moderate to high frequency response characteristics are further explored with the Cross-Coupling Analysis. As mentioned earlier, increasing hinge offset is expected to increase cross-coupling. Results for configurations of various flapping hinge offsets are shown in Figure 5-2 for three flaperon and three stabilator settings.
Figure 5-2. Cross-Coupling Analysis for configurations of various flapping hinge offsets using three flaperon allocations 휔푑푓푙푎푝 (left) and three stabilator allocations 휔푑푠푡푎푏 (right).
82 Results for no flaperon/stabilator allocation show that increasing hinge offset decreases roll-due- to-pitch (푝/푞) cross-coupling and slightly increases pitch-due-to-roll (푞/푝) cross-coupling with a net improvement in overall handling qualities. This contradicts the earlier prediction that increasing hinge offset also increases cross-coupling characteristics. Although the swashplate phasing was held constant (thus degrading handling qualities for cross-coupling with increasing hinge offset), the NLDI is designed to account for this, so it is difficult to decipher the meaning of the trends.
Finally, flaperon allocation again has a larger impact on cross-coupling behavior for higher hinge offset configurations; this time it actually degrades handling qualities by increasing roll-due-to- pitch towards the Level 1/Level 2 boundary, which is consistent with the results from Chapter 4.
Generally, it appears that stabilator allocation has a similar magnitude impact on each configuration, where increasing allocation leads to improved handling qualities. However, increasing stabilator allocation decreases roll-due-to-pitch (푝/푞) cross-coupling for the low hinge offset configuration and increases roll-due-to-pitch for the highest hinge offset configuration; the middle configuration did not show a distinct trend, although it appears that generally increasing allocation decreases roll-due-to-pitch. Pitch-due-to-roll decreases for all three configurations as stabilator allocation increases. While all three of these configurations are within the boundaries of desired Level 1 handling qualities, it appears that there is a trade-off between hinge offset and stabilator allocation, where the configuration with the best handling qualities meets in the middle.
Finally, attitude quickness is evaluated for all configurations using RAQ for flaperon allocation and PAQ for stabilator allocation. Again, the PAQ results showed very little variation with hinge offset and/or stabilator allocation, and therefore only one plot is included that features all variations. Results are shown in Figure 5-3.
83
Figure 5-3. Roll Attitude Quickness analysis for configurations of various hinge offset and flaperon allocation 휔푑푓푙푎푝 (top row and bottom left) and Pitch Attitude Dropback analysis for configurations of various hinge offset and stabilator allocation 휔푑푠푡푎푏 (bottom right).
Results show that all configurations and all flaperon/stabilator allocations result in desired Level 1 handling qualities. Without redundant control allocation (top left), it appears that increasing hinge offset improves handling qualities for large attitude changes, while there is little to no effect for small attitude changes (0.4 sec triangular pulse input). This is expected since increasing hinge offset also increases the moment generated by the main rotor for a given amount of cyclic pitch. All trend lines have a negative slope, which agrees with the earlier conclusion that nonlinearities have a larger negative impact at larger amplitudes. Generally, increasing flaperon allocation improves
84 handling qualities, shown by the upward shift of data points between plots. Results from the medium magnitude flaperon allocation (0.375) show a distinct trend with small spread between configurations of varying hinge offset, while results from the large magnitude flaperon allocation
(0.75) are not as clear. In general, it appears that increasing hinge offset for large flaperon allocations actually degrades handling qualities, especially for higher attitude changes. This may be due to the stiffer rotor system providing a larger amount of damping when most of the control is allocation to the flaperons. Thus flaperons tend to approach their saturation limit in the larger amplitude maneuvers.
Overall, the predictive analysis of flapping hinge offset with redundant control allocation revealed that increasing hinge offset is generally beneficial for small amplitude/moderate to high frequency applications. In these cases, it appears that increasing hinge offset also increases flaperon effectiveness. Additionally, no concrete trends were noted from the cross-coupling analysis. As for the large amplitude/low to moderate frequency response characteristics, increases hinge offset generally improves handling qualities when no flaperon/stabilator allocation is used. While increasing flaperon allocation to 0.375 improves handling qualities for all configurations, the results from the highest flaperon allocation (0.75) show that increasing hinge offset actually degrades handling qualities. These results should be explored in future research projects.
Wing Span and Tail Area
The purpose of incorporating airplane features, such as a wing and horizontal tail, on a rotorcraft is to create a compound aircraft that is capable of performing like a helicopter and an airplane. These added features are beneficial at high speeds, but can negatively impact the main rotor performance in hover by creating download. Nevertheless, the wing and horizontal tail strongly impact the overall stability and control or the aircraft. This section examines this impact
85 through predictive analysis of the flight handling qualities of a compound rotorcraft. The
configurations involved in this study use the baseline command filters (휔푛훷 = 3.75 푟푎푑/푠푒푐 and
휔푛훳 = 2.50 푟푎푑/푠푒푐) and flapping hinge offset (푒 = 1.80) but vary the wing span and horizontal tail area separately, resulting in 5 configurations, shown in Table 5-3, where the baseline configuration uses a wing span of 45 ft and a horizontal tail area of 45 ft2.
Table 5-2. Configurations involved in the study of wing/tail size. Configuration Type Wing Span (ft) Tail Area (ft2) Baseline 45 45 Low Wing Span 35 45 High Wing Span 55 45 Low Tail Area 45 35 High Tail Area 45 55
Aircraft dynamics relate the roll damping coefficient (퐿푝) to the wing span using Equation
5-2, where the corresponding stability and control derivative (퐶푙푝) is given in Equation 5-3 (all equations from References [1, 23].
1 (5-2) 퐿 = 휌푢 푏2푆퐶 푝 4 0 푙푝
푎푤(1 + 3휆) (5-3) 퐶푙 = − + 퐶푙 푝 12(1 + 휆) 푝퐹
In this study, all of the variables in Equation 5-1 and Equation 5-2 (density 휌, velocity 푢0, wing area 푆, wing lift slope 푎푤, taper ratio 휆, and vertical tail contribution 퐶푙 ) are held constant except 푝퐹 for wing span (푏). Therefore, it is easy to see that increasing wing span also increases the magnitude of the roll damping coefficient (퐿푝). Similarly, the pitch damping coefficient (푀푞) is defined in
Equation 5-4, where 퐶푚푞 is given in Equation 5-5.
1 (5-4) 푀 = 휌푢 푐̅2푆퐶 푞 4 0 푚푞
86
푙2 (5-5) 퐶 = (−2푎 푡 ) 푆 푚푞 푡 푆푐̅2 푡
In Equation 5-3 and Equation 5-4, all values are held constant (density 휌, velocity 푢0, mean aerodynamic chord 푐̅, wing area 푆, horizontal tail lift slope 푎푡, and longitudinal distance from center of gravity to horizontal tail aerodynamic center 푙푡) except for horizontal tail area, 푆푡. Therefore, it
can easily be seen that increasing that horizontal tail area increases the magnitude of 퐶푚푞 which in turn increases the magnitude of 푀푞. Increasing roll and pitch damping tends to increase the bandwidth of the configuration, but decrease control power. It is expected that these configurations will require more control action to achieve higher angular rates, leading to worse performance in attitude quickness.
Additionally, flaperons are sized with respect to the wing. Therefore, increasing the wing span is expected to also increase the flaperon span. The contribution of flaperons (indicated by subscript 푓) to the rolling moment of the aircraft is shown in Equation 5-6, and the corresponding stability and control derivative is given in Equation 5-7, where 푦̅ is the spanwise distance from the center of gravity to the aerodynamic center of the wing and 푐푓 is the chord.
1 2 (5-6) 퐿훿 = 휌푢0푏푆퐶푙 푓 2 훿푓
2푎푓푦̅푓푏푓푐푓 (5-7) 퐶푙 = − 훿푓 푆푏
Using the definition of the derivative from Equation 5-6 and factoring out the constants, Equation
5-5 becomes Equation 5-8.
1 2푎푓푦̅푓푏푓푐푓 (5-8) 퐿 = 휌푢2푏푆 (− ) = (−휌푢2푎 푐 )푦̅ 푏 훿푓 2 0 푆푏 0 푓 푓 푓 푓
The only terms in Equation 5-7 that change are 푦̅푓 and 푏푓. As wing span increases, the span-wise distance from the center of gravity to the aerodynamic center of the flaperon (푦̅푓) and the span of the flaperon (푏푓) also increase. Therefore, increasing wing span also increases the magnitude of
87
퐿훿푓, meaning that the flaperon effectiveness increases. As a result, flaperon gearing is expected to recover the degradation from increased damping. In all configurations, the ratio of the flaperon aerodynamic center to wing span is held constant (푦̅푓/푏 = 0.256), so the aerodynamic center increases proportional to the span.
Similarly, the stabilator is sized with respect to the horizontal tail. Therefore, increasing the tail area is expected to also increase stabilator area. The contribution of the stabilator to the pitching moment of the aircraft is shown in Equation 5-9, and the corresponding stability and control derivative is given in Equation 5-10, where 푉퐻 is the tail volume ratio.
1 2 (5-9) 푀훿 = 휌푢0푐̅푆퐶푚 푒 2 훿푠
푎푠푆푡푙푡 (5-10) 퐶푚 = −푎푠푉퐻 = − 훿푠 푆푐̅
Using the definition of the derivative from Equation 5-9 and factoring out the constants, Equation
5-9 becomes Equation 5-11.
1 푎 푆 푙 1 (5-11) 푀 = 휌푢2푐̅푆 (− 푠 푡 푡) = (− 휌푢2푎 푙 ) 푆 훿푠 2 0 푆푐̅ 2 0 푠 푡 푡
The only term in Equation 5-10 that changes is the tail area (푆푡). Therefore, increasing the tail area
also increases the magnitude of 푀훿푒, meaning that the stabilator effectiveness increases.
In summary, it is expected that increasing wing span will increase roll damping and the effect of flaperons, while increasing tail area will increase pitch damping and the effect of the stabilator. Results of the Bandwidth/Phase Delay test for configurations of various wing span and flaperon settings (0, 0.375, and 0.75) as well for configurations of various tail area and stabilator settings (0, 0.375, and 0.75) are shown in Figure 5-4.
88
Figure 5-4. Bandwidth/Phase Delay analysis for configurations of various wing sizes using three flaperon allocations 휔푑푓푙푎푝 (top) and configurations of various tail sizes using three stabilator allocations 휔푑푠푡푎푏 (bottom).
In general, results show that variation of wing span and flaperon gearing primarily affects the Roll
Bandwidth Analysis, while variation of tail area and stabilator gearing primarily affects the Pitch
Bandwidth Analysis. All configurations are within desired Level 1 handling qualities except for the configurations with increased wing/tail size and no redundant control allocation. It appears that increasing wing span decreases the roll bandwidth and slightly decreases the roll phase delay, shown in the top left plot of Figure 5-4, which results in degraded handling qualities. Additionally,
89 the impact of flaperon gearing (spread between data points of the same color) slightly increases.
This is consistent with the earlier expectation from Equation 5-7. Higher flaperon effectiveness corresponds to a larger increase in roll bandwidth for the same change in gearing (for example, from 휔푑푓푙푎푝 = 0.375 to 휔푑푓푙푎푝 = 0.75) at the cost of a slight increase in roll phase delay.
Therefore, increasing wing span generally degrades handling qualities for small amplitude/moderate to high frequency responses unless it is coupled with an increase in flaperon allocation.
In the pitch axis, it appears that increasing tail area decreases the pitch bandwidth and slightly decreases the pitch phase delay, leading to degraded handling qualities (bottom right plot of Figure 5-4). While stabilator allocation improves handling qualities, it appears that the stabilator effectiveness does not improve from the baseline configuration (45 ft2) to the high tail area configuration (55 ft2). However, decreasing tail area (35 ft2) degrades the effectiveness of the stabilator (shown by a decrease in spread of data points of the same color). This suggests that the relationship in Equation 5-7 might not impact that Pitch Bandwidth Analysis to the extent that was expected. Therefore, increasing tail area generally degrades handling qualities for small amplitude/moderate to high frequency responses.
The small amplitude/moderate to high frequency responses characteristics are further explored with the Cross-Coupling Analysis. Results for configurations of various wing span and flaperon settings (0, 0.375, and 0.75) as well for configurations of various tail area and stabilator settings (0, 0.375, and 0.75) are shown in Figure 5-5.
90
Figure 5-5. Cross-Coupling Analysis for configurations of various wing sizes using three flaperon allocations 휔푑푓푙푎푝 (left) and configurations of various tail sizes using three stabilator allocations 휔푑푠푡푎푏 (right).
Results for configurations without flaperon allocation show that increasing wing span increases the magnitude of roll-due-to-pitch (푝/푞) cross-coupling, improving handling qualities, and slightly decreases the magnitude of pitch-due-to-roll (푞/푝) cross-coupling, slightly degrading handling qualities. If the Level 1/Level 2 boundary were expended for higher magnitudes of 푞/푝, then the net effect of increasing wing area is that it improves the overall cross-coupling handling qualities.
Additionally, increasing wing span also increases the flaperon effectiveness, but this time negatively by drastically decreasing the magnitude of 푝/푞, thus degrading handling qualities. The configuration with the best cross-coupling handling qualities has a large wing span (55 ft2), but little to no flaperon allocation. This confirms the earlier prediction that increasing wing span also increases roll damping, therefore making the aircraft more resistant to roll-due-to-pitch cross- coupling. As for tail size, it appears that increasing tail area with no stabilator allocation slightly decreases the magnitude of pitch-due-to-roll (푝/푞) cross-coupling, thus slightly degrading handling qualities, and has little to no impact on roll-due-to-pitch (푞/푝) cross-coupling. Additionally, increasing tail area increases the stabilator effectiveness by increasing the magnitude of 푝/푞 for a
91 particular gearing, thus improving cross-coupling handling qualities. It appears that increasing the tail area (with stabilator allocation) might actually increase roll damping while having little to no effect on pitch damping.
Finally, the moderate to large amplitude/low to moderate frequency response characteristics are evaluated using Roll Attitude Quickness and Pitch Attitude Dropback analyses.
It is expected that the extra damping due to larger wing and tail size will decrease control power, thus requiring more control action to achieve higher angular rates. Results for configurations of various wing span and flaperon settings (0, 0.375, and 0.75) as well for configurations of various tail area and stabilator settings (0, 0.375, and 0.75) are shown in Figure 5-6.
92
Figure 5-6. Roll Attitude Quickness analysis for configurations of various wing sizes and flaperon allocation 휔푑푓푙푎푝 (top row and bottom left) and Pitch Dropback analysis for configurations of tail sizes and stabilator allocation 휔푑푠푡푎푏 (bottom right).
As expected, results from the Pitch Attitude Dropback analysis show very little impact of tail size and stabilator gearing on handling qualities. Further research should be conducted to verify these results using other analysis methods. As for the roll axis, increasing wing span degrades handling qualities by shifting the curve towards the Level 1/Level 2 boundary. In fact, the configuration with no flaperon allocation and largest wing span actually falls in the Level 2 category. Increasing wing span also increases roll damping, leading to less control power. As a result, cyclic control saturation
93 becomes more severe, and the roll quickness degrades. However, the addition of flaperon gearing relieves control saturation, thus improving handling qualities. Increasing wing size also increases the effectiveness of the flaperons, which is seen through the decrease in vertical spread for configurations with high flaperon allocation.
In summary, results supported the earlier prediction that increasing wing/tail size increases roll/pitch damping, leading to degraded handling qualities due to control saturation. However, the addition of redundant control gearing mitigates the effect of control saturation, resulting in very similar handling qualities between configurations.
94
Chapter 6
Concluding Remarks
Next-generation rotorcraft are likely to feature fly-by-wire technology and various forms of compounding to meet demanding requirements. As such, it is essential that designers include handling qualities assessment from the beginning phase of design so that the vehicle performs as expected. This has the potential to benefit the rotorcraft community by reducing the time and money spent fixing handling qualities problems during flight testing and ultimately improving future rotorcraft by incorporating a broader perspective. As a result, various rotorcraft experts from industry and academia are studying the relationship between frequency response to control deflection (quantitative/predictive metrics) and actual performance in Mission Task Elements
(MTEs, piloted simulation).
To support this endeavor, the Pennsylvania State University (PSU) studied the impact of controller and design parameter modifications on handling qualities of a generic compound rotorcraft using the industry standard ADS-33E-PRF. Results showed that increasing the command filter parameter generally improved handling qualities for all amplitudes and frequencies of cyclic pitch deflections. However, increasing the parameter past the ideal frequency might result in problems with cross-coupling. These results were verified by piloted simulation at the PSU
VLRCOE flight simulator facility in 2017, where results trended towards improved handling qualities for higher command filter parameters.
Additionally, predictive metrics revealed that the addition of redundant controls improved handling qualities by relieving control saturation, but might lead to problems with phase delay for high bandwidth configurations in the roll axis. Trends from variation of flapping hinge offset were
95 generally unclear, but there seemed to be a tradeoff between hinge offset and redundant control allocation where the configuration with desired handling qualities had a medium-sized hinge offset and some redundant control allocation. Lastly, increasing wing/tail size generally degraded handling qualities by decreasing bandwidth for low amplitude deflections and damping the response to high amplitude deflections (leading to problems with control saturation). However, the addition of redundant control allocation (flaperons) recovered the performance and resulted in desired handling qualities for all configurations.
This study is important because it identified basic trends between various design parameters (including controller parameters) and handling qualities of a next-generation compound rotorcraft. It is recommended that future research refine the metrics for handling qualities requirements in ADS-33E-PRF by incorporating more flight test and piloted simulation results.
Additionally, time should be spent developing new mission task elements and a moderate to large amplitude/low to moderate frequency predictive metric for the pitch axis. A list of conclusions and recommendations for future work are included below.
Conclusions
Command Filter Predictive Analysis
1. The Bandwidth/Phase Delay predictive analysis for variation in command filter parameter
showed that increasing the parameter generally improved handling qualities by increasing
bandwidth in the respective axis (roll command filter impacted roll axis and pitch command
filter impacted pitch axis). However, increasing or decreasing the command filter
parameter past the ideal values increased the roll-axis phase delay and decreased the pitch-
96 axis phase delay. The ideal command filter parameters were appropriate since the baseline
configuration was within Level 1.
2. The Cross-Coupling predictive analysis for variation in command filter parameter showed
that increasing roll parameter increased roll-due-to-pitch (degraded handling qualities) and
decreased the spacing between configurations, while increasing or decreasing the pitch
parameter past ideal increased pitch-due-to-roll (degraded handling qualities).
3. The Roll Attitude Quickness (RAQ) predictive analysis for variation in roll command filter
parameter showed that increasing the input length (change in attitude) led to nonlinear
effects that created a negative slope. Although increasing the roll parameter generally
improved handling qualities, larger input lengths led to control saturation that degraded
handling qualities for higher parameter configurations. Additionally, handling qualities
tended to degrade for all configurations as airspeed increased.
4. The Pitch Attitude Dropback predictive analysis for variation in pitch command filter
parameter showed that increasing the parameter did not have a significant impact on the
handling qualities. All configurations were within Level 1, which is suspicious due to the
high profile drag of the compound aircraft at high airspeeds.
Command Filter Piloted Simulation Analysis
1. Results from the Combat Break Turn MTE were consistent with the RAQ predictive
analysis, where increasing roll command filter parameter generally improved handling
qualities with some deficiencies for the configuration with the largest parameter. Low
bandwidth configurations showed problems with heading overshoot, while high bandwidth
configurations had problems with airspeed maintenance and initial attitude capture.
97 2. Results from the Roll Sum of Sines MTE showed similar trends to the Bandwidth/Phase
Delay and Cross-Coupling predictive assessments where increasing roll command filter
parameter improved handling qualities. The pilots experienced distracting pitch-due-to-roll
cross-coupling with configurations of low bandwidth, which led to large tracking errors
and moderate pilot compensation. The high bandwidth configuration performed the best
and allowed for a more aggressive strategy with reduced pilot workload. However,
increasing the roll parameter only increased roll-due-to-pitch cross-coupling, which was
the opposite of the results from the Cross-Coupling predictive analysis.
3. Results from the Pitch Sum of Sines MTE showed similar trends to the Bandwidth/Phase
Delay and Cross-Coupling predictive assessments, where increasing pitch command filter
parameter improved handling qualities. However, it was difficult to draw conclusions
because there was a large disparity between pilot scores. It was also surprising that the
pilots did not notice a difference in cross-coupling between configurations.
4. Results from the Roll Attitude Capture and Hold (BACH) MTE showed similar trends to
the Bandwidth/Phase Delay and Cross-Coupling predictive analyses where increasing the
roll command filter parameter improved handling qualities. Cross-coupling created a high
workload for all configurations, although increasing the parameter resulted in crisper
response. Overall it seemed that the predictive analyses did not provide an accurate
representation of the cross-coupling characteristics of the aircraft. This might be a result of
poorly defined boundaries of the predicted analyses.
5. Results of the Pitch Attitude Capture and Hold (PACH) MTE showed similar trends to the
Bandwidth/Phase Delay and Cross-Coupling predictive analyses where increasing the
pitch command filter parameter to the ideal value improved handling qualities and
increasing the parameter past the ideal value degraded handling qualities. Results were
98 fairly consistent between pilots. Oscillatory behavior seemed to drive the HQR scores, and
cross-coupling was not a problem.
6. Results of the Acceleration/Deceleration (Accel/Decel) MTE showed similar trends to the
Bandwidth/Phase Delay and Cross-coupling predictive analyses where increasing the pitch
command filter parameter improved handling qualities. Maintenance of speed capture and
altitude led to increased workload that limited HQR for low bandwidth configurations.
Flaperon/Stabilator Gearing Predictive Analysis
1. The Bandwidth/Phase Delay predictive analysis for variation in redundant control gearing
showed that variation of flaperon gearing only impacted the roll-axis and variation of
stabilator gearing only impacted the pitch-axis. The baseline configuration (no redundant
control allocation, ideal command filter parameters) had Level 1 handling qualities, which
showed that redundant controls were not necessary for small-amplitude/high frequency
attitude changes. However, some flaperon authority improved handling qualities, while a
large amount might actually degrade handling qualities. Additionally, increasing stabilator
authority tended to improve handling qualities for high bandwidth configurations.
2. The Cross-Coupling predictive analysis for variation in redundant control gearing showed
that increasing flaperon allocation increased roll-due-to-pitch (푝/푞), which might actually
lead to Level 2 handling qualities for a high bandwidth configuration with high flaperon
authority. However, increasing stabilator authority improved handling qualities.
3. The Roll Attitude Quickness predictive analysis for variation in flaperon allocation showed
that control saturation was recovered by allocating part of the control to the flaperons and
led to improved handling qualities. Additionally, flaperon allocation might actually delay
the negative impact of nonlinearities in the system by providing more control authority.
99 The Pitch Attitude Dropback predictive analysis showed that variation in stabilator
allocation had very little impact on handling qualities.
Flapping Hinge Offset Predictive Analysis
1. Results from the Bandwidth/Phase Delay predictive analysis for variation in flapping hinge
offset and redundant control gearing were generally consistent with the results from
Conclusion 11, where flaperon gearing primarily impacted the roll-axis and stabilator
gearing primarily impacted the pitch-axis. Although all configurations had Level 1
handling qualities, increasing hinge offset increased the spread due to flaperons which is
probably a result of the hinge offset creating a larger moment for a given amount of cyclic
pitch. Additionally, increasing hinge offset decreased roll-axis phase delay and increased
pitch-axis phase delay (unexpected).
2. The Cross-Coupling predictive analysis for variation in flapping hinge offset showed no
concrete trends. Increasing hinge offset decreased roll-due-to-pitch (푝/푞) and slightly
increased pitch-due-to-roll (푞/푝) with a small net improvement in overall handling
qualities, which was unexpected. Flaperon allocation degraded handling qualities for
higher hinge offset configurations, while stabilator allocation improved handling qualities
for all configurations. There appeared to be a tradeoff in terms of handling qualities
between hinge offset and stabilator allocation.
3. The Roll Attitude Quickness predictive analysis for variation in flapping hinge offset
showed that increasing hinge offset improved handling qualities for large attitude changes,
while there was little to no effect for small attitude changes, which was expected since
increasing hinge offset also increases the moment generated by the main rotor for a given
amount of cyclic pitch. Although all configurations were within Level 1 handling qualities,
100 trend lines had negative slopes due to the impact of nonlinearities at larger amplitudes. The
addition of some flaperon allocation improved handling qualities, while the results from
full flaperon allocation were not as clear. However, increasing hinge offset for large
flaperon allocation degraded handling qualities, which was probably due to increased
damping.
Wing Span and Tail Area Predictive Analysis
1. Results from the Bandwidth/Phase Delay predictive analysis showed that variation in wing
span primarily affected the roll-axis, while variation in tail area primarily affected the
pitch-axis. All configurations were within Level 1 handling qualities except for the
configurations with increased wing/tail size and no redundant control allocation. Increasing
wing span increased the impact of flaperons, decreased roll bandwidth, and slightly
decreased roll phase delay, which resulted in degraded handling qualities. Increasing tail
area decreased pitch bandwidth and slightly decreased pitch phase delay, which resulted in
degraded handling qualities. Stabilator allocation generally improved handling qualities,
but increasing tail area did not seem to impact its effectiveness (although decreasing tail
area degraded effectiveness).
2. Results from the Cross-Coupling predictive analysis showed that increasing wing span
increased the magnitude of roll-due-to-pitch (improved handling qualities) and slightly
decreased the magnitude of pitch-due-to-roll (slightly degraded handling qualities), with a
net improvement in overall handling qualities. Increasing wing span also increased
flaperon effectiveness, shown by drastic degradation in handling qualities. Additionally,
increasing tail area decreased the magnitude of pitch-due-to-roll (slightly degraded
handling qualities) and had little to no impact on roll-due-to-pitch. Increasing tail area also
101 increased stabilator effectiveness, thus improving handling qualities. It appeared that
increasing tail area with stabilator allocation might actually increase roll damping while
having little to no effect on pitch damping.
3. Results from the Roll Attitude Quickness predictive analysis showed that increasing wing
span degraded handling qualities by making cyclic control saturation more severe, but the
addition of flaperons relieved control saturation with improved effectiveness for larger
wing spans. Results from the Pitch Attitude Dropback predictive analysis showed very
little impact of tail area on handling qualities.
Future Work
1. It is expected that next-generation rotorcraft configurations will feature various degrees of
compounding and automatic flight control to meet the demanding requirements of future
civilian and military missions. This study analyzed a generic rotorcraft compounded with
a wing and auxiliary propulsion system to enable higher forward flight speeds. Future work
should investigate alternative configurations and their impact on flight handling qualities.
2. ADS-33-PRF predictive and piloted simulation methods should be evaluated for each new
configuration change to determine the effectiveness of these standards for next-generation
rotorcraft. Additional predictive metrics and MTEs should be developed to assist in the
assessment of flight handling qualities and to help define a stronger relationship between
predictive metrics and actual performance.
3. Placement of control effectors in the PSU Vertical Lift Research Center of Excellence
(VLRCOE) flight simulator facility should be investigated to better represent the cockpit
of next-generation rotorcraft. Additionally, the Heads-Up-Display (HUD) should be
modified to provide appropriate visual cueing for all Mission Task Elements (MTEs).
102 4. The Cross-Coupling analysis predicted that increasing the roll command filter parameter
would increase the roll-due-to-pitch, but the results from the RSOS MTE showed the
opposite (increasing the parameter decreased pitch-due-to-roll and did not impact roll-due-
to-pitch). This disparity should be further analyzed to conclude if the ADS-33E-PRF Cross-
Coupling predictive analysis truly represents the handling qualities of compounded
configurations in high-speed flight.
5. A predictive metric should be developed to evaluate moderate to large amplitude/low to
moderate frequency response characteristics in the pitch axis. Every configuration tested
in this study had Level 1 handling qualities, which was unexpected. Therefore, additional
metrics should be used to verify this result and/or identify trends. A piloted simulation
MTE should also be developed to assist in handling qualities assessment and verification
of predictive metric results.
6. The PSOS MTE should be refined since there was a large disparity between pilot scores.
Future research should incorporate more pilot evaluations to verify that trends actually
exist.
7. The relationship between cross-coupling and handling qualities should be investigated for
the Accel/Decel MTE. If a strong relationship is found, then perhaps that can be
incorporated into future MTEs to better assess pitch-axis precision tracking response
characteristics. Otherwise, new MTEs should be developed that better test for cross-
coupling issues since pilots did not notice these characteristics in the PSOS and PACH
MTEs.
8. Results from the predictive analyses of redundant control gearing should be compared to
result from piloted simulation using the Break Turn MTE to verify that trends actually
exist.
103 9. Results showed that some flaperon allocation tended to improve handling qualities, but full
flaperon allocation generally degraded handling qualities. The impact of flaperon
allocation should be further explored to ultimately define the allocation that maximizes
handling qualities.
10. Since results from flapping hinge offset and wing/tail size were often unclear, these
parameters should be investigated using alternative predictive metrics and piloted
simulation MTEs.
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