Flight Dyanamic

8/12/2019 Flight Dyanamic

1/40

CHAPTER 1

INTRODUCTION


2/40

1.1 General

Interest in unmanned-aerial-vehicles (UAVs) and micro-aerial-vehicles (MAVs) in

recent years has increased significantly. These aircraft are useful for applications ranging

from military to scientific research because of their ability to perform dangerous missions

without risking human life. Also because their payload can be much smaller than a pilot,

there are less limitations to how small they can become. At Brigham Young University (BYU)

in particular, small UAV research has exploded during the last few years. Faculty and

students work together on many small UAV projects. Research activities include developing

new airplanes for commercial use, participating in the annual Micro-Aerial-Vehicle

competition, and developing autonomous flight vehicle systems. It seems that interest in

small UAVs will continue to grow around the world as new applications will demand new

UAV solutions and designs. Airplanes come in all shapes and sizes. Before the emergence of

UAVs, airplane designers were constrained in how small they could go because of the

necessity to carry a human pilot onboard. By removing the pilot, and due to increasing UAV

component technology, UAV designs have decreased in size significantly. As time goes on,

smaller and smaller UAV solutions will become available. Indeed, the term small UAV

undoubtedly had a much different meaning just a few years ago than it does now.

Small, remotely operated aircraft present unique challenges and advantages to both

designer and pilot. Because of a drastically higher crash frequency, it seems that small UAVs

are more susceptible to dangerous and sometimes fatal instabilities than large airplanes.

This may be due in part to quicker design cycles and the lower stakes of small UAV crashes

relative to the high stakes of a large airplane crash. Perhaps designers are sometimes

more eager to get out and see if this thing flies, than they are to do the rigorous design

work necessary to ensure a stable and successful first flight. Often, experienced airplane

designers and pilots seem to develop an uncanny intuition for diagnosing and solving

problems with all kinds of airplane instabilities. The goal of this thesis research is to capture

the intuition and knowledge of such capable engineers and pilots, evaluate it quantitatively

and provide a simple tool to new, less-experienced designers. This tool will allow them to


3/40

improve their small UAV design methods to include considerations of both static and

dynamic stability.

Airplanes, including small UAVs, are a classic engineering example of design

tradeoffs. It often seems impossible to improve one aspect of performance without

degrading another. Accordingly, stability is one aspect of airplane performance that must be

balanced with all the others. There are two types of stability which are:

Static Stability

Static stability is an essential part of the basic airplane design process already included indesign methodologies.

Dynamic Stability

Dynamic stability will effect on performance and handling qualities is generally poorly

understood by new designers. This lack of understanding makes it difficult to include

dynamic stability into a typical design process.

An airplane that has positive dynamic stability does not automatically have

positive static stability. The designers may have elected to build in, for example, negative

static stability and positive dynamic stability in order to achieve their objective in

maneuverability. In other words, negative and positive dynamic and static stability may be

incorporated in any combination in any particular design of airplane.

An airplane may be inherently stable, that is, stable due to features incorporated in

the design, but may become unstable due to changes in the position of the center of gravity

(caused by consumption of fuel, improper disposition of the disposable load, etc.). Stability

may be (a) longitudinal, (b) lateral, or (c) directional, depending on whether the disturbance

has affected the airframe in the (a) pitching, (b) rolling, or (c) yawing plane.


4/40

Figure 1.2: Stability diagram

This research will provide relatively simple methods to approximate the static

behaviour and handling qualities of small UAVs while still in the design stage, similar to

analysis used in a conventional large aircraft design process. In the future, this will hopefully

become a powerful tool in the hands of small UAV designers at BYU and elsewhere. Perhaps

in the future it will be possible to avoid fatal crashes like that shown in Figure 1.1

Figure 1.1 Damage from small UAV crashes like this one often result from a lack of

designing for stability.


5/40

1.2 The Objective of This Study

The objective of this thesis research is to better understand aircraft static stability as it

applies to small UAVs and to develop a method for including static stability analysis into the

design process. This objective can be broken down into four sub-objectives which will be the

topics covered in this thesis:

Develop a mathematical model to predict the static stability of small UAVs based on

knowledge of the geometry and inertias of the airframe.

Verify the accuracy of the model using known airplane data.

Provide analysis of the driving design parameters and guidelines for small UAV

static stability.

1.3 Stability: A Requirement for All Airplanes

Among the significant but often-overlooked obstacles to powered flight overcome by

the Wright brothers was the question of how to build an airplane that was stable enough to

be controlled and manoeuvred by a pilot. It has been shown that the Wrightsfirst powered

airplane in 1903 was so unstable that only the Wrights themselves could fly it, due to

extensive self-training on their previous glider versions in 1902. (Abzug, 3) As they and other

aviation pioneers took steps to solve the stability and controls problem, the capabilities and

performance of airplanes increased significantly. In the early days of flight, it was observed

that certain designs of airplanes were more stable and controllable than others, but it was

not until the 1930s that much theory existed to explain why. Much of the modern stability

and control theory and specifications were not developed until the 1960s or later. (Abzug,

33) Airplanes of all sizes must be capable of stable, trimmed flight in order to be controllable

by a human pilot and useful for various applications. Stable flight by a human pilot is

possible only if the airplane possesses static stability, a characteristic that requires

aerodynamic forces on the airplane to act in a direction that restores the plane to 12 a

trimmed condition after a disturbance. Dynamic stability requires that any oscillations in

aircraft motion that result from disturbances away from equilibrium flight conditions must


6/40

eventually dampen out and return to an equilibrium or trimmed condition. Certain

dynamic instabilities can be tolerated by a human pilot, depending mostly upon pilot skill

and experience. If computer-augmented feedback control is used even statically unstable

aircraft can be flown successfully. (Abzug, 312) Both static and dynamic stability

characteristics can be predicted while an airplane is still in the design stage of development.

Many companies such as DAR Corporation, whose theory will be used extensively in this

chapter, have developed software to do just that. (Roskam, I, 461) To do so, it is necessary

to have a precise knowledge of the geometric and inertial properties of the airframe. Static

stability is predicted using information about the airplane aerodynamic center and the

center of gravity as well as other geometric parameters. Dynamic stability is predicted using

the airframe geometric and inertial properties to calculate the natural frequencies, damping

ratios and time constants of the characteristic dynamic modes of the six degree-offreedom

aircraft model. The handling qualities of an airplane are said to be a measure of how well an

airplane is able to perform its designated mission. They are usually evaluated using flight

test data and pilot feedback on performance. Current military specifications relate levels of

acceptable aircraft handling qualities to the frequencies, damping ratios and time constants

of the dynamic modes.


7/40

1.4 Static Stability

An airplane possesses static stability if the aerodynamic forces and moments

introduced on the airframe as a result of it being disturbed from equilibrium tend to act in a

direction that will return the airplane to an equilibrium condition. Static stability is

analogous to a marble in a bowl. If the marble is disturbed from an equilibrium position at

the bottom of the bowl, gravitational forces at all other positions will tend to pull it back

towards the bottom. Aerodynamic forces and moments on a statically unstable aircraft will

tend to move it away from a trimmed flight condition when it is perturbed from equilibrium.

This condition is analogous to a marble on the top of a smooth hill or balancing a pendulum

upside down. This condition would be nearly impossible for a human pilot to control, but

could be possible if some form of feedback control is used. For a more complete overview of

static stability, readers should consult Anderson, Chapter 7, in the References section. Static

stability can be considered as a special case (steady-state) of the aircraft dynamics. It is

exhibited in both the decoupled longitudinal and lateral-directional axes. It will also become

clear that both longitudinal and lateral-directional static stability are a prerequisite for

dynamic stability.

1.4.1 Longitudinal Static Stability

Longitudinal static stability is essential to ensure that a human pilot can successfully fly an

airplane without stability augmentation. It depends mostly upon a parameter known as the

static margin, defined as the distance between the aircraft center of gravity and the neutral

point of the aircraft, normalized by the mean geometric chord, c , of the wing. An airplane

with longitudinal static stability must first possess a positive (nose-up) pitching moment

from the combination of the aerodynamic forces and moments on the wing and tail. For

flying wings, an airfoil with a natural positive pitching moment must be chosen or washout

and wing sweep must be combined to give the airplane a natural positive pitching moment.

If this condition is met, a positive static margin, defined as the center of gravity in front of

the neutral point, will ensure static stability. Static stability can be simply represented by


8/40

plotting the pitching moment of the aircraft about its center of gravity versus angle-of-

attack as shown in Figure 2.1.

Figure 1.1 A moment coefficient curve for an airplane possessing longitudinal static stability.

Any disturbance away from trim will result in aerodynamic forces and moments which will

act in a direction that will tend to return the plane to equilibrium.

A small static margin (center of gravity near the neutral point) will provide marginal static

stability and will be represented by a nearly flat line on the graph in Figure 2.1. A large static

margin will provide a steep line and can make an airplane feel nose-heavy.It may cause

the plane to be less controllable because it doesnt respond to control inputs. These

constraints on the static margin are presented in Figure 2.2. It is also important that the

static margin be chosen that will allow the plane to be trimmed at a reasonable angle of

attack.


9/40

Figure 1.2 Aircraft center of gravity envelope. The c.g. must fall within these limits set by the

stability and controllability of the aircraft. (Kimberlin, 2006).

To express longitudinal static stability in mathematical terms, we must first define

the aerodynamic centre, . It is the longitudinal location along the centreline of theaircraft measured from the leading edge of the wing about which the pitching moment is

constant over a range of angles-of-attack. It is also the point at which the lift effectively acts.

1.4.2 Lateral-Directional Static Stability

Similar to the longitudinal case, an UAV possesses directional (about the z-axis or

yaw-axis) static stability if a slight increase in sideslip results in a restoring yawing moment

as well as a restoring side force. This is sometimes called weathervane stability. Lateral

(about the x-axis or roll axis) static stability is expressed in terms of dihedral effect. If these

are negative, then the airplane will possess positive lateral stability and will exhibit a

negative rolling moment (left wing down) for a positive sideslip (nose left).


10/40

1.5 Dynamic Stability

An UAV possesses dynamic stability if the amplitudes of any oscillatory motions

induced by disturbances eventually decrease to zero relative to a steady-state flight

condition. This means that if an UAV experiences a small perturbation from trimmed flight,

it will eventually return to trim on its own. This is analogous to a marble in a bowl eventually

coming to rest at the bottom of the bowl. If the amplitude of oscillatory motion instead

tends to increase with time, the airplane is said to be dynamically unstable. Dynamic

instabilities are obviously undesirable, but certain mild dynamic instabilities can be

tolerated by a human pilot. If automatic controls are used, more severe dynamic instabilities

can also be tolerated. Graphical representations of dynamic stability and instability are

shown in Figures 1.3 and 1.4. To study dynamic stability, it is necessary to analyze the well-

known differential equations of UAV motion. For small perturbations, these equations can

be decoupled into longitudinal and lateral-directional portions, with 3 degrees of freedom in

each. Small perturbation theory also allows us to approximate the actual non-linear

equations as linear differential equations with constant coefficients while ignoring any less

significant non-linear aerodynamic effects. This greatly simplifies the analysis of the dynamic

modes of aircraft motion. However, it should also be explained that we will only considerthe static stability only in this study.

Figure 1.3 A graphical example of dynamically stable UAV motion relative to a steady-state

condition.


11/40

Figure 1.4 A graphical example of dynamically unstable UAV motion.

1.6 Longitudinal Dynamic Stability

The longitudinal dimensional stability derivatives represent the partial derivatives of

linear or angular acceleration due to either the displacement, velocity or acceleration

depicted by the subscript. The capital letters X, Y and Z represent derivatives of linear

accelerations in the corresponding directions while the capital letters L, M and Nrepresent

derivatives of angular accelerations according to the conventions. The dimensional stability

derivatives are the dimensionalised form of the non-dimensional stability derivatives

depicted by the letter C with appropriate subscripts. Both can be derived experimentally or

analytically with a careful analysis of the airframe geometry and inertial properties. The

variables u, , and are the perturbed velocity, angle-of-attack and pitch attitude

respectively and represent the three longitudinal degrees of freedom. By taking the Laplace

transform of this system, the differential equations become simple polynomials in the s

variable and are transformed from the time domain into the frequency domain.


12/40

CHAPTER 2

CALCULATIONS


13/40

Table 2.1 AirTRAC UAV Parameter

NO PARAMETER VALUE

1 Distance from leading edge

to the chord of wing

2 Distance from leading edge

to the center of gravity 3 Wing root chord Croot= 0.16013m

4 Wing tip chord Ctip= 0.11952m

5 Mean chord length 6 Wing area 7 Wing aspect ratio

9.22

8 Wing lift curve slope 9 Wing taper ratio 0.7464

10 Tail area 11 Tail aspect ratio 6.70612 Tail lift curve slope 2.40713 Distance from chord of tail

to center of gravity

m14 Tail efficiency 15 Tail volume ratio 0.42716 Length of fuselage 17 Fuselage body side area 18 Distance from nose to cg 19 Maximum width fuselage

20 Vertical height of fuselage at

length of fuselage 21 Vertical height of fuselage at

length of fuselage


14/40

2.1 Pitching Moment

2.2 Longitudinal static stability

An equilibrium point can be stable, unstable or neutral stable. A stable equilibrium

point is characterized by and . The criteria for longitudinal static stability canbe pictured as:

Figure 2.1 Moment coefficient curve with negative slope


15/40

Figure 2.2 Moment coefficient curve with negative slope

If the airplane is statically longitudinal stable, it has the initial tendency to return to its

equilibrium position.

For our analysis, the graph can be shown as:


16/40

Figure 2.3 CMcgversus of the AirTRAC UAV

For the calculation of the trim angle it is given by:

Neutral Point and Static Margin Determination

To find the neutral point of the AirTRAC UAV, there is several assumption that have been

made:

i)

based on the center of gravity

based on the neutral point

y = -0.944x - 0.0147

-30

-20

-10

0

10

20

30

-20 -10 0 10 20

C

oefficientofmoment,Cm

Angle of attack,

Cm versus

TOTAL

wing

tail

Linear (TOTAL)


17/40

ii) iii)

Figure 2.4 Effect of the location of the centre of gravity relative to the neutral point on

static stability

Then, equation of the neutral point is given as below:

( )

Component Method Weight

(kg)

x(mm) Wx

(kgmm)

Staticall Unstable

Neutral Stabilit

Staticall Stable

trim


18/40

Fuselage Stinton 1.6 217.00 347.20

Wing Stinton 2.4 241.20 578.88

Payload Estimation 2.2 203.66 448.05

Tail Stinton 0.4 885.43 354.17

Fuel Estimation 3.4 275.05 935.17

Powerplant Stinton 4.0 107.64 430.56

Structure Stinton 5.4 276.89 1495.21

Equipment and

services

Stinton 0.6 208.27 124.96

Total 20.0 Total 4714.20

Figure 2.5 The final location of components of the AirTRAC UAV and the weight

Thus, based on the table above, we can calculate the location of the maximum,

minimum, forward and after centre of gravity. The location of maximum centre of gravity is

including the all components of AirTRAC UAV with the fuel in the tank

The location of minimum centre of gravity is including all components with empty fuel of

AirTRAC UAV

Next, the static margin is defined as below:


19/40

The neutral point is located at aft the centre of gravity,. In other word, theallowable margin in shifting the neutral point is between 0.236 m to 0.436 m measured

from the leading edge of the AirTRAC UAV.

The pitching moment equation M, will determine the longitudinal static stability of

an aircraft. The changes of pitching moment with angle of attack will determine the

characteristics of the aircraft longitudinal stability. A longitudinal statically stable at An aircraft has a good controllability if it can be trimmed at the positive angle of

attack, . For our aircraft by referring the equation of total coefficient of moment at center

gravity, the value of which is less than zero. Thus it can be stated that our

aircraft is statically stable. Besides the value of Cmo is 0.0147 which is more than zero. It

made the aircraft can be trimmed at positive .


20/40

The location of centre of gravity is an important characteristic in order to determine

the aircraft stability while in preliminary design. In order to achieve a good longitudinal

stability, Centre of Gravity (CG) should be ahead of Neutral Point (NP), which is

Aerodynamic Centre (AC) of the whole aircraft. NP is the longitudinal location of CG when

the aircraft is neutrally stable. When CG is ahead of NP, the weight tends to correct the

upset thus lead to stable aircraft. While when the CG is behind NP, the weight worsens the

upset thus make the aircraft become unstable. NP is the most aft CG location.

During operation the CG location is not fixed. So that it is important to know the limit

and range of CG location. The determination of CG when is important to estimatethe margin of aircraft longitudinal static stability.

is called stick fixed neutral point. For

our UAV, the value of XNPis 0.272. When XCGis at XNP, aircraft is neutral stable.

The location of centre of gravity closer to tail which means the aircraft has shorter

tail moment arm, causes the aircraft to destabilize. The longer the tail moment arm, thehigher the pitching moment due to pitch rate, q and change in angle of attack, . Theaircraft pitching moment due to angle of attack, M changes with the location of CG and

aircraft aerodynamic centre. The closer is the CG to the AC, the aircraft will pitch at higher

angle of attack and it resulted in higher natural frequency. Not only it is significant in

longitudinal motion of the aircraft, but the tail moment arm also affects the yawing and

rolling moment of the aircraft due to yawing rate in lateral mode.

2.3 Lateral static stability

Since the UAV is used primarily for surveillance purposes, a higher degree of rolling

stability is required in order for the UAV to provide a clearer picture of the surroundings.

The higher the slope of the rolling moment curve, the higher degree of rolling stability the

UAV will possess.


21/40

The main contribution to an aircraft's lateral stability is the dihedral of the wings that will

create a restoring moment when the aircraft is disturbed about its lateral axis A dihedral of

6 degrees is introduced to enhance the rolling stability of the aircraft. The effect of dihedral

on can be calculated as follows:

Since plain rectangular wings are used, the integration can be simplified to be:

The aircraft's lift curve slope can be calculated using the following equation:

( ) ( )

( )

In this calculation, fuselage contribution is neglected due to the fact that there is a

relatively large gap between the fuselage and the wings which will reduce the effect of

change in flow induced by the fuselage.


22/40

Figure 2.6 Lateral Stability

As shown in Figure 2.6, the negative value of tells us that the aircraft is laterally stable.

2.4 Directional static stability

Contribution of wing and fuselage


23/40

There are two main contributions to the aircraft's directional stability, namely the fuselage

and the vertical tail. The fuselage's contribution to the yawing moment slope can be

determined as follows:

Whereis the empirical wing-body interference factor determined by the geometry of thefuselage. The value of can be determined from Figure 2.7 using the required geometricparameters. The required parameters for determination of are summarized in the tablebelow:

Table 2.2 Fuselage geometric parameter for determination of 0.5345 2.5658

1.19

1


24/40

Figure 2.7 Wing body interference factor

Referring to Figure 2.7, the value of is determined to be approximately 0.005.

The value is a correction factor that depends on the fuselage's Reynolds number,


25/40


26/40

Contribution of vertical tail

The vertical tail contribution to the fuselage's directional stability can be determined as

follows:

( )

Where the term can be estimated as follows:

( ) = the distance parallel to the z axis from the wing root quarter chord to the fuselagecenterline= the max fuselage depth

( )

( )


27/40

The total aircraft's yawing moment slope is the algebraic sum of each contribution.

1.8531rad-1

For static directional stability, the slope of the yawing moment curve must always be

positive,

as shown in Figure 2.9 below.

Figure 2.9 Static directional stability

The positive value of the aircraft's yawing moment slope is 1.8531rad-1

. It indicates that the

aircraft is directionally stable.


28/40

CHAPTER 3

HANDLING QUALITIES


29/40

3.1 Flying and Handling Qualities

The handling qualities of an airplane are said to be a measure of how well anairplane is able to perform its designated mission. They are usually evaluated using flight

test data and pilot feedback on performance. Current military specifications relate levels of

acceptable aircraft handling qualities to the frequencies, damping ratios and time constants

of the dynamic modes. It is an obvious fact that an airplanes geometric and inertial

properties, among other factors, influence how well or how poorly it flies and how

effectively it is able to perform its intended mission. The handling qualities or flying

qualities of an airplane are a measure of airplane performance relative to its intended

mission and describe how well or poorly a particular airplane flies.

MIL-STD-1797A defines what is meant by the term handling qualities. Those

qualities or characteristics of the aircraft that govern the ease and precision with which a

pilot is able to perform the tasks required in support of an aircraft role. Another definition

from the Cooper-Harper Rating Scale defines handling qualities as those qualities or

characteristics of an aircraft that govern the ease and precision with which a pilot is able to

perform the tasks required in support of an aircraft role. (Hodgkinson, 7). An analysis of

handling qualities will include both qualitative and quantitative information about the pilots

ability to control the airplane. It includes analysis of how quickly an airplane responds to

various inputs as well as the control effort that must be exerted by the pilot. Handling

qualities are used to compare various aircraft designs and are based on both subjective pilot

opinion and objective flight data.


30/40

UAVs are unique in the sense that there are no set hard and fast rules for the

employment. Mission and role dictate the general design requirements and small size with

no human on-board makes the launch, recovery methods up to the imagination of the

designer. Thus it is very difficult to classify and define all the methods, however, the general

guidelines as given in are listed below. The mini UAVs / RPVs are considered to be a

different class and the UAVs are categorized as Table 1 according to the weight and

maneuverability levels. Similarly the flight phase categories and the levels of handling

qualities are also defined in Tables 2 and 3.

Table 1:Categories of UAVs

Class Category Mission examples

I Small Light Weight Mini RPVs

II Low maneuvarability Surveillance, Reconnaissance high

altitude

III Medium maneuvarability Surveillance, Reconnaissance low

altitude

IV Highly maneuverability UCAVs


31/40

Table 2:Flight phase categories

Category Definition Flight types

A

Requiring rapid maneuvering, precision

tracking, and/or precise flight path

control.

Reconnaissance, Target

Acquisition, Terrain Following,

In-flight refuelling etc

B

Requiring gradual maneuvers, without

precision tracking, although precise

flight path control may be required.

Climb, Cruise, Loiter, Decent,

Aerial Delivery, Emergency

climb and decent.

C

Launch Recovery Phase requiring rapid

maneuvering, precision tracking, or

precise flight path control.

Arresting Gear Landings, Net

Capture, Conventional take-off

and landing

D

Launch Recovery Phase requiring

gradual maneuvers, without precision

tracking, although precise flight path

control may be required.

Catapult take off, Approach,

Parachute Recovery, Approach

to recovery envelop


32/40

Table 3:Levels of flying qualities

Level Definition

I (A/M) (Normal System Operation) Flying qualities are clearly adequate to

accomplish mission flight phase

2 (A/M) (Degraded Mission) Flying qualities remain adequate to perform

mission flight phase with moderate degradation or increase in operator

work load, a degradation of mission effectiveness or both

3 (A/M) (Recoverability) Degraded qualities remains adequate to control the

vehicle with Category A phases terminated successfully and Categories

B, C, and D phases completed sufficiently to recover the vehicle

A: Automatic Control Mode

M: Manual Control Mode

Or RPV mode


33/40

3.2 Cooper Harper Rating Scale

The dynamics and characteristics of the aircraft are related to the flying and handlingqualities. The natural frequency and damping ratio value influence on how easy or difficult

to control the aircraft. Cooper Harper rating scale is being used as a reference to give rating

for the aircraft. The rating scale gives from 1 to 10 that will influence the flying and handling

qualities to aircraft.

Figure 1:A modified form of Cooper Harper scale for UAV evaluation


34/40

3.3 Handling Qualities of UAV

Most of the military specifications are derived by setting limits on the natural

frequencies, damping ratios and time constants of the various dynamic modes. For instance,

in the longitudinal axes, the natural frequency and damping ratio of the Phugoid mode

describe what aircraft motions occur when the airplane seeks a stabilized airspeed

following a disturbance. Because the natural frequency of this characteristically slow mode

of oscillation is typically on the order of 50-100 seconds for large aircraft, a pilot is easily

able to control and compensate for unwanted motion. Therefore, no limits are placed on

the natural frequency of the phugoid mode.

However, Level 1 handling qualities require that the damping ratio be positive so as

to eventually cause any oscillation to die out over time. Negative damping ratios are

tolerated only at Level 3 because they indicate a tendency to become unstable if

uncorrected. Similar requirements, putting limits on the natural frequency and damping

ratio of the Dutch-roll mode as well as the time constants of the spiral and roll modes, exist

for the case of lateral-direction handling qualities. For the spiral mode, no specific limitation

is placed on the time constant because a slightly unstable spiral mode is typically

acceptable. However, lower limits on the time it takes for the bank angle to double from an

initial disturbance of 20 degrees are specified. For the case of a stable spiral mode, the bank

angle will actually decrease after a disturbance.


35/40

Table 4:Short period damping ratio limits

Level Category A and C Flight Phases Category B Flight Phases

Minimum Maximum Minimum Maximum

1 0.35 1.30 0.30 2.00

2 0.25 2.00 0.20 2.00

3 0.15* - 0.15* -

Table 5:Phugoid damping ratio limits

Phugoid mode

Level 1 > 0.04

Level 2 > 0

Level 3 T2> 55s

Table 6:Spiral mode (minimum time to double amplitude flying qualities

Class Category Level 1 Level 2 Level 3

I and IV A 12 s 12 s 4 s

B and C 20 s 12 s 4 s

II and III All 20 s 12 s 4 s

Table 7:Roll mode (maximum roll time constant) flying qualities (in seconds)

Class Category Level 1 Level 2 Level 3

I, IVA

1.0 1.410

II, III 1.4 3.0

All B 1.4 3.0 10

I, IVC

1.0 1.410

II, III 1.4 3.0


36/40

Table 8:Dutch roll flying qualities

Level Category Class Min *

Min n*,

rad/s

Min n,

rad/s

1

A I, IV 0.19 0.35 1.0

II, III 0.19 0.35 0.4

B All 0.08 0.15 0.4

C

I, II-C 0.08 0.15 1.0

IV

II-L, II 0.08 0.15 0.4

2 All All 0.02 0.15 0.4

3 All All 0.02 - 0.4

Where C and L denote carrieror land-based aircraft

*The governing damping requirement is that yielding the larger value of


37/40

CHAPTER 4

DECLARATION OF APPROVAL


38/40

Declaration for Approval

With a detailed research had done, we had come out with a solid design for our proposed

UAV design for air traffic surveillance. The SolidWork drawing is a general idea picked from

our group members idea that had the greatest score. So we really hope with the data and

initial estimation that had been calculate will make you confident and let us proceed with this

design.

Proposed by:

____________________

(NUR FATIN MOHAMAD IDRIS)

_____________________

(MOHD ZAMRI MHD NASIR)

_____________________

(NUR HAZIQAH BAROM)

____________________

(AHMAD AMEER ABDULLAH)

____________________

(MUHAMMAD HAIKAL ABDUL JAMAL)

Approved by:

____________________

(EN MD NIZAM B DAHALAN)


39/40

CHAPTER 5

REFERENCES


40/40

1. Anderson, John D. Introduction to Flight Fourth Edition. McGraw-Hill International

Editions, 1999.

2. Nelson, Robert C. Flight Stability and Automatic Control - 2nd Edition . McGraw-Hill

International Edition, 1998

3. Anderson, John D.Aircraft Performance and Design. McGraw-Hill International Editions,

1999.

4. Raymer, Daniel P. Aircraft Design: A Conceptual Approach 2nd Edition. American

Institute of Aeronautics and Astronautics,

5. Tang Shiao Loong, Infantry Section Unmanned Aerial Vehicle (UAV) Flight Controls and

TestingNational University of Singapore 2012

Documents

Flight Dyanamic