Flight Systems and Control: A Practical Approach

By Tian Seng Ng

This book focuses on flight vehicles and their navigational systems, discussing different forms of flight structures and their control systems, from fixed wings to rotary crafts. Software simulation enables testing of the hardware without actual implementation, and the flight simulators, mechanics, glider development and navigation systems presented here are suitable for lab-based experimentation studies. It explores laboratory testing of flight navigational sensors, such as the magnetic, acceleration and Global Positioning System (GPS) units, and illustrates the six-axis inertial measurement unit (IMU) instrumentation as well as its data acquisition methodology. The book offers an introduction to the various unmanned aerial vehicle (UAV) systems and their accessories, including the linear quadratic regulator (LQR) method for controlling the rotorcraft. It also describes a Matrix Laboratory (MATLAB) control algorithm that simulates and runs the lab-based 3 degrees of freedom (DOF) helicopter, as well as LabVIEW software used to validate controller design and data acquisition. Lastly, the book explores future developments in aviation techniques.

• Language: English
• Publisher: Springer
• Release date: Apr 13, 2018
• ISBN: 9789811087219

Book preview

© Springer Nature Singapore Pte Ltd. 2018

T. S. NgFlight Systems and ControlSpringer Aerospace Technologyhttps://doi.org/10.1007/978-981-10-8721-9_1

1. Introduction

Tian Seng Ng¹  

(1)

Nanyang Technological University School of Mechanical and Aerospace Engineering, Singapore, Singapore

Tian Seng Ng

Email: [e-mail]

1.1 Preliminary

The employment of the flight computer control is very common in airplanes as we used to travel around the world in the automated aircraft system. Advances in technological efforts have made air vehicle autonomous. Therefore, the studies and understandings of the scope are necessary for building automated aircraft systems. The more complex automatic systems need a thorough grasping of the related topics in the book. The materials presented in this book explains the concept of navigation from the acceleration, heading or direction of the air vehicle. The electronics based sensing components we built provides a good foundation for laboratory-based experimental study and practices. Besides, the book also introduced flight mechanics, simulators and flight accessories for the gliders. Excerpts of the three degrees of freedom helicopter system can open insights to aircraft control engineers. Moreover, the readers can also appreciate the knowledge and its applications to the modern control technique for the rotorcraft system. Interesting flight instrumentation and programmable control techniques can enhance the interest of the book explorers. We can discover up to date topics in the air vehicle automated system and the modern technology used in controlling the air automobile system.

1.2 Book Highlights

The book presents the studies of the flight vehicles and its navigational systems. It also analyzes the several forms of flight structures and its control systems. Software simulation enables us to test the hardware without actual implementation. Moreover, we also introduced the hardware means for air vehicle navigation. Altogether, they comprised the hardware and software necessary in a flight system.

The usefulness of the magneticfield finds its application in the headingsensor.

Modern satellite navigation method based on GPS is experimented in the laboratory.

Electroniccircuit for the acceleration sensor is constructed and tested.

The integration of the three sensing units comprises the flight instrumentation system.

Applied control theory presents to design control system for the flying crafts.

MATLAB algorithm runs to simulate the helicopter control behavior before real time implementation.

Software model of the air vehicle simulates to validate the controllers we design.

LabVIEW software interfaces to detect and read the flight navigational sensors. The parametersfeedback can then control the aircraft or flight movement and trajectory.

1.3 Chapters’ Organization

Chapter 2 familiarizes the basic mechanics of a flightsystem. Readers will explore the effects of aircraftcontrol surfaces and their related equations, besides the cockpit flight instruments.

Chapter 3highlights the low-cost electronic sensors for navigation. These basic components serve as dead reckoning sensors for an air vehicle.

Chapter 4 enhances the understanding of flight characteristics and their effects on flight performance with flight simulators.

Chapter 5 studies the tandem rotor helicopter for understanding the dynamics, control and stability analysis of the flight vehicle.

Chapter 6 introduces the basic communication and control hardware for the various kinds of rotorcrafts, micro-air vehicles (MAVs) and quadcopter systems.

Chapter 7 illustrates the programming and control of a quadrotor system. LabVIEW design of the flying craft verifies our controller design for the system.

Chapter 8 presents the integration of the flight instrumentation on the software environment for sensing the direction, acceleration, and location of the air vehicle. The developed software detects the navigation sensors.

Chapter 9 discusses the various natural resources available as renewable energy for the aerospace industries. It also reviews the developments of the new vertical takeoff and landing aircraft, etc.

© Springer Nature Singapore Pte Ltd. 2018

T. S. NgFlight Systems and ControlSpringer Aerospace Technologyhttps://doi.org/10.1007/978-981-10-8721-9_2

2. Flight Mechanics

Tian Seng Ng¹  

(1)

Nanyang Technological University School of Mechanical and Aerospace Engineering, Singapore, Singapore

Tian Seng Ng

Email: [e-mail]

2.1 Basic Flight System

The study of the flight mechanics provide the fundamentals of aerospace engineering relating to any conventional flight vehicles. To fly and control the aircraft, we need to know the flight mechanics of the system. In an air vehicle, we have to control the ailerons (roll), the elevator (pitch) and the rudder (yaw). Aerodynamic property such as the lift coefficient is also important to the flying vehicle. For example, how does the flap affects the take-off properties of the airplane? Furthermore, the power setting relates to the flying characteristics of the aircraft as well. The flight spoilers are controllable by the pilot. We used them together with the brake, and thruster to reduce the speed of the air vehicle, at the point of landing on the ground. The instrument panel in the cockpit of an aircraft gives the readout of any metering of the system. It is these feedback indicators in the panel that the pilot can control the airplane to perform a smooth ride. Two types of aircraft are of different characteristics ideal for studies and comparisons. We have the single engine propelled Skyhawk or Cessna C172SP airplane, and the two-turbofan engine, fixed wing Bombardier Learjet 45 airplane. The maximum fuel payload for the Cessna airplane is about 154.22 kg, which is equivalent to 340 lb. The landing gear or landing pair of wheel is retractable only in the Learjet airplane system (Figs. 2.1, 2.2, 2.3 and 2.4).

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Fig. 2.1

Cessna instrument panel

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Fig. 2.2

Learjet instrument panel

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Fig. 2.3

Cessna airplane

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Fig. 2.4

Learjet airplane

2.2 Steady Straight Level Flight

The understanding of the fundamental forces in flight is critical as it affects the flight performances of the flying plane. For aircraft flying in the atmosphere, there are mainly four forces of flight: lift (L), drag (D), thrust (T) and weight (W), as illustrated in Fig. 2.5. Lift and drag are an aerodynamics type of forces that are present due to the relative motion of the aircraft with respect to the air. The thrust produces by the propulsion system pushes the airplane forward and overcomes the drag. Weight is present due to gravity, which is a natural force pulling the aircraft downward. Flight characteristic of an aircraft is dependent on the interaction of these four forces. The standard type of flight is the so-called steady straight level flight, where the aircraft flies in a straight-line trajectory with constant airspeed and altitude with its wings in level position. In this case (see Fig. 2.6), where the engine thrust angle αT

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Fig. 2.5

Four forces of flight

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Fig. 2.6

Force equilibrium in steady straight level flight

$$ T\cos \alpha_{T} = D $$

(2.1)

$$ L + T\sin \alpha_{T} = W $$

(2.2)

indicates the angle the thrust makes with respect to the flight direction. For conventional aircraft performing conventional flight, αT is often small, and thus the following approximate relations applies (see Fig. 2.7)

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Fig. 2.7

Assumed force equilibrium in steady straight level flight

$$ T = D $$

(2.3)

$$ L = W $$

(2.4)

These approximations have been used widely for determining an aircraft’s flying characteristics. Equation (2.3) suggests that the thrust required maintaining the steady straight level flight is equivalent to the drag experienced by the aircraft. Similarly, in steady straight level flight, the lift needed is the same as the weight of the aircraft at the particular instant.

Power Required:

Power (P) is defined as the time rate of change of the work done and is equivalent to force (F) in the direction of motion times the speed (V).

$$ P = FV $$

(2.5)

The power required (PR) to maintain steady straight level flight is defined as the thrust required times the steady-state airspeed (V∞) , and by Eq. (2.3), this can be expressed as

$$ P_{R} = D\,V_{\infty } $$

(2.6)

In general, the variation of the power required with airspeed and altitude is not linear because D is both a function of air density (p∞), which varies with altitude and airspeed (V∞) is as follow,

$$ D = \frac{1}{2}p_{\infty }\,V_{\infty }{^{2}} SC_{D} $$

(2.7)

where S is the aircraft’s wing platform area, and CD is the aircraft’s drag coefficient. The plot of PR versus V∞ at a given altitude is usually called the power-required curve. For the aircraft using a propeller engine, we usually cannot directly obtain the power required from the reading of the instrument panel. The power produced by the propeller usually has a complex relationship with the engine RPM (revolution per minute), the propeller diameter, and the air density. For the case, a simplified assumption used is that the engine RPM is proportional to the power produced by the propeller for the flight condition used.

Thrust Required:

The thrust required (Tg) to maintain steady straight level flight is defined as the thrust required times the steady-state airspeed (V∞) , and from Eq. (2.3), it can be expressed as:

$$ T_{g} = D $$

(2.8)

In general, the variation of the thrust required with airspeed and altitude is not linear, because D is both a function of air density (p∞), which varies with altitude and airspeed. Thus, we can define

$$ T_{g} = \frac{1}{2}p_{\infty } V_{\infty }{^{2}} S\,C_{D} $$

(2.9)

with reference to Eq. (2.7). Where S is the aircraft’s wing platform area, and CD is the aircraft’s drag coefficient. The relationship between Tg and V∞ at a given altitude is usually called the thrust required curve.

Maximum Airspeed:

Generally, for the steady straight level flight at a specific altitude in the relatively high range of the aircraft, the drag increases as the airspeed increases. From Eq. (2.3), this implies that the thrust or power required in maintaining the steady straight level flight also increases. If the airspeed is further increased, it will reach a point where the power required is the maximum the propulsion of the aircraft can generate. In the situation, the airspeed for a steady straight level flight cannot increase further. When the aircraft is in steady straight level flight with maximum propulsion power at a specific altitude, then the steady airspeed at which the aircraft flies is the maximum airspeed for that altitude.

Coefficient of Lift:

Like drag, lift (L) is also a function of air density (p∞), which varies with altitude and airspeed (V∞), as follows:

$$ L = \frac{1}{2}p_{\infty } \,V_{\infty }{^{2}} S\,C_{L} $$

(2.10)

We calculated the coefficient of lift (CL) during the steady straight level flight from Eq. (2.4), which yields:

$$ C_{L} = \frac{2w}{{P_{\infty} V_{\infty}{^{2}} S}} $$

(2.11)

It is clear from Eq. (2.11), that the coefficient of lift can be determined if other parameters such as the instantaneous aircraft weight, wing area of the aircraft and air density at the flight altitude are known. We can assume the air density at ground (sea) level is 1.2 kg/m³, where the air density p at 3000 feet high (above sea level) is 1.1 kg/m³.

2.3 Takeoff Maneuver

During the investigation of the dynamic performance of an aircraft at takeoff, all the basic forces acting on the aircraft must be established. Figure 2.8 illustrates the force diagram for an aircraft during its ground roll.

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Fig. 2.8

Forces acting on the aircraft during takeoff

In addition to the forces such as lift (L), drag (D), thrust (T) and weight (W), the aircraft experiences a resistance force (R) due to rolling friction between the tires and the ground. For takeoff, over most of the ground roll, W is reasonably constant. However, L will vary with the velocity according to Eq. (2.10). At the onset of lifting off the ground, L = W. Note that this condition is the same as Eq. (2.4). This simple condition enables us to estimate the lift coefficient CL of a particular aircraft at takeoff. Therefore, we can estimate the takeoff lift coefficient using Eq. (2.11). For both the aircraft systems, we can calculate the lift coefficients. We achieved this by recording the take-off speed of both planes and use Eq. (2.11) to determine their lift coefficients. We then compare both their readings to determine which airplane can take-off faster. Secondly, we can find the power to overcome the drag to achieve a steady straight flight at a certain altitude. The characteristic between the engine rpm and the steady airspeed is thus plotted. Thirdly, we determined the maximum airspeed of the aircraft at specific altitude. We recorded the airspeed achievable for flying steady straight level flight at the altitude with maximum power setting.

2.4 Glider Design

We can design and build a remotely controlled glider. A set of components that might be easily available can be useful for the glider design. We need the following materials.

(1)

A Styrofoam glider

(2)

A radio transmitter

(3)

A radio receiver

(4)

A pair of servo motors

(5)

One four-cell battery

(6)

A set of push rods and servo horns

(7)

One bottle of cyano glue

(8)

One sheet of balsa wood

The glider is controllable in the pitch and yaw, or in pitch and roll modes. Note that a standard airplane is controllable in all three axes, but if the glider is stable and has sufficient dihedral, two-axis control is sufficient. We can control the pitch with the left stick on the radio controller so that pulling the stick pitches the airplane upwards. The yaw/roll is controllable with the right stick.

During your test flights, the airplane will crash frequently. Make it sturdy. It is advisable to have the wings detach easily. It will reduce the risk of breakage during a crash. We can connect the battery to any of the remaining channels on the receiver. Make sure that the polarity is correct at all times. Reverse polarity may destroy the receiver or the servos. The balance of the airplane is very important. Small shifts in weight can have a great impact on the flight performance. You may want to use blue tack to trim the airplane. You may also consider the different design configurations, such as bi-planes, canards, flying wings, twin boom, etc.