**2. Principles of flight**

An aircraft is a complex machine using the application of multidisciplinary engineering sciences. The major engineering groups can be indicated as follows:


h.Armament system (for military aircraft)

i. Air egress (ejection) system

j. Software (embedded as well as operational software)

#### **2.1 Atmosphere**

For the aerodynamic study, air is considered an ideal gas which follows gas laws. Therefore, the variation of air properties with respect to altitude is important. The International Standard Atmosphere (ISA) is used for comparison of performances of aircraft designed by different countries. The ISA is defined as:


The atmosphere is divided in two layers. The lower layer is called 'troposphere' where temperature decreases linearly with altitude (6.5°C/km altitude rise, known as lapse rate). Air pressure decreases with altitude as shown in **Figure 1**. Air density can be estimated from gas equation (ρ = P/RT, where R is universal gas constant). **Figure 1** also shows the different types of clouds in the atmosphere [1–5].

The upper layer of the atmosphere is called 'stratosphere' where T remains constant at −56°C (ISA condition). In troposphere, with increase in altitude, both pressure and temperature decrease reducing the density, thus lowering engine mass flow reducing thrust, lift and drag. These values optimize at 10–12 km which is known as cruise altitude for jet aircraft. Due to decrease in mass flow, jet engines cannot operate at very high altitude.

#### **2.2 Aerodynamics and flight mechanics**

Movement of air over an aircraft generates aerodynamic forces and moments. Due to change in the air properties, the aerodynamic forces and moments also vary with altitude. Flight dynamics looks at these aerodynamic forces and includes thrust and gravity forces to study aircraft motion. Further the aerodynamic forces generated due to deflections of control surfaces are added as applied control forces to study the dynamics of the flight path including stability and controllability of the aircraft [1, 5].

Airflow over a body obeys three basic aerodynamic equations: these are conservation of mass, conservation of momentum and conservation of energy. Solving

**3**

**Figure 2.**

**Figure 1.**

*Subsonic flow over [2].*

*Military Aviation Principles*

*DOI: http://dx.doi.org/10.5772/intechopen.87087*

which forces and moments can be estimated.

quantities and are represented by CL, CD, and CM.

*Variation of pressure of atmospheric pressure with altitude.*

these equations, we obtain velocity distribution over the aircraft surfaces from

Airflow over an airfoil (cross section of wing) is shown in **Figure 2**. Due to the airfoil camber, air particles traveling over the upper surface have to cover longer distance than the air flowing on the lower surface. In order to comply with the law of conservation of mass, air particle on the upper surface speeds up to cover longer distance (due to camber) than the airflow over the lower surface. According to the law of conservation of momentum, the increase in speed is compensated by the decrease in pressure. This creates differences in pressures with lower surface air pressure being more than that of the upper surface. This differential pressure gives rise to upward force. The vertical component of this upward force is the wing lift, and axial component is the drag. Further the resultant of the air pressure of the upper surface and lower surface does not pass through the same point, which creates a turning moment known as pitching moment. An airfoil of unit thickness will produce lift, drag and pitching moment coefficients. These are dimensionless

Aerodynamics of supersonic flow is, however, different. A supersonic flow over an airfoil at angle of attack (AOA) is shown in **Figure 3**. On the upper surface, flow

#### *Military Aviation Principles DOI: http://dx.doi.org/10.5772/intechopen.87087*

*Military Engineering*

control systems)

surization systems)

**2.1 Atmosphere**

i. Air egress (ejection) system

h.Armament system (for military aircraft)

Altitude (H): sea level (0 m) Temperature (T): 288.15°K

Density (ρ): 1.225 kg/m3

cannot operate at very high altitude.

**2.2 Aerodynamics and flight mechanics**

Pressure (P): 1.01325 MPa (14.7 psi)

d.Engine/power plant system

c.Mechanical system (hydraulics, pneumatics, landing gear, fuel and flight

e.Electrical system (power generations, distribution and emergency power)

and utility management system) and instrument system

j. Software (embedded as well as operational software)

f. Avionic (communication, navigation, weapon aiming, displays and warnings

g.Environmental systems (air-conditioning, life support system and cabin pres-

For the aerodynamic study, air is considered an ideal gas which follows gas laws. Therefore, the variation of air properties with respect to altitude is important. The International Standard Atmosphere (ISA) is used for comparison of performances of aircraft designed by different countries. The ISA is defined as:

The atmosphere is divided in two layers. The lower layer is called 'troposphere' where temperature decreases linearly with altitude (6.5°C/km altitude rise, known as lapse rate). Air pressure decreases with altitude as shown in **Figure 1**. Air density can be estimated from gas equation (ρ = P/RT, where R is universal gas constant).

**Figure 1** also shows the different types of clouds in the atmosphere [1–5].

The upper layer of the atmosphere is called 'stratosphere' where T remains constant at −56°C (ISA condition). In troposphere, with increase in altitude, both pressure and temperature decrease reducing the density, thus lowering engine mass flow reducing thrust, lift and drag. These values optimize at 10–12 km which is known as cruise altitude for jet aircraft. Due to decrease in mass flow, jet engines

Movement of air over an aircraft generates aerodynamic forces and moments. Due to change in the air properties, the aerodynamic forces and moments also vary with altitude. Flight dynamics looks at these aerodynamic forces and includes thrust and gravity forces to study aircraft motion. Further the aerodynamic forces generated due to deflections of control surfaces are added as applied control forces to study the dynamics of the flight path including stability and controllability of the

Airflow over a body obeys three basic aerodynamic equations: these are conservation of mass, conservation of momentum and conservation of energy. Solving

**2**

aircraft [1, 5].

these equations, we obtain velocity distribution over the aircraft surfaces from which forces and moments can be estimated.

Airflow over an airfoil (cross section of wing) is shown in **Figure 2**. Due to the airfoil camber, air particles traveling over the upper surface have to cover longer distance than the air flowing on the lower surface. In order to comply with the law of conservation of mass, air particle on the upper surface speeds up to cover longer distance (due to camber) than the airflow over the lower surface. According to the law of conservation of momentum, the increase in speed is compensated by the decrease in pressure. This creates differences in pressures with lower surface air pressure being more than that of the upper surface. This differential pressure gives rise to upward force. The vertical component of this upward force is the wing lift, and axial component is the drag. Further the resultant of the air pressure of the upper surface and lower surface does not pass through the same point, which creates a turning moment known as pitching moment. An airfoil of unit thickness will produce lift, drag and pitching moment coefficients. These are dimensionless quantities and are represented by CL, CD, and CM.

Aerodynamics of supersonic flow is, however, different. A supersonic flow over an airfoil at angle of attack (AOA) is shown in **Figure 3**. On the upper surface, flow

#### **Figure 1.**

*Variation of pressure of atmospheric pressure with altitude.*

**Figure 2.** *Subsonic flow over [2].*

**Figure 3.** *Supersonic flow [2].*

expands having higher flow area, and flow on the lower surface gets compressed due to lower area of flow. Expansion of flow is associated with increase in Mach number and decrease in static pressure. It may be appreciated that at zero AOA, there will be no difference in upper and lower surface flow. As the supersonic airfoil generates lifts only due to AOA, supersonic airfoils are symmetrical airfoils.

The lift produced increases with increase in AOA as shown in **Figure 4**. However, beyond certain AOA, flow separates from the airfoil, lift suddenly drops, and drag rises due to increase of wake area. This is known as 'stall'. With further increase of AOA, the wake region will increase and thus aggravate the situation. This being a flight safety hazard, airworthiness regulations require stall warning system and protection from stall recovery procedure to be incorporated. Subsonic aircraft stalls around 14–18° of AOA, while supersonic aircraft stalls around 24–28°

**5**

**Figure 6.**

**Figure 5.**

*6 DOF of air vehicle [4].*

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<sup>2</sup> <sup>ρ</sup>V2SCL;D <sup>=</sup> \_1

L = \_1

*DOI: http://dx.doi.org/10.5772/intechopen.87087*

audio warning with stick shaker is installed.

are the measure of controllability of the aircraft.

*The force and moment system working on an aircraft in flight [2].*

AOA. In supersonic flow, lift curve slope is very flat, and realization of stall is rather difficult. In view of this in supersonic aircraft, in addition to an AOA indicator,

An aircraft in flight has six degrees of freedom (DOF); these are three translations (motion in forward, lateral and vertical directions) and three angular motions, viz. roll (rotation about longitudinal axis), pitch (rotation around lateral axis) and yaw (rotation about vertical axis). An aircraft is steered to the desired direction by operating the respective control surface. This is shown in **Figure 5**. The lift, drag and pitching moment produced by the wings can be written as

wing area and C is the mean aerodynamic chord length. The force and moments acting on an aircraft in x-z plane during flight are shown in **Figure 6**. The lift, drag and pitching moments act at the aerodynamic centre of the aircraft. The forces and moments are in equilibrium. If the equilibrium is disturbed, the aircraft will move from its flight path. For example, if the lift is increased more than weight, the aircraft will float up, and if the thrust produced is more than the drag, the aircraft will accelerate. The reverses are also true. Similarly, if any control surface is deflected, ΣM about the axis will be disturbed, and the aircraft will rotate about that axis. The degrees of turn of flight path achieved per degree of deflection of control surfaces

<sup>2</sup> <sup>ρ</sup>V2SCCM, where S is the total wetted

<sup>2</sup> <sup>ρ</sup>V2SCD; and <sup>M</sup> <sup>=</sup> \_1

**Figure 4.** *Stall phenomenon [4].*

#### *Military Aviation Principles DOI: http://dx.doi.org/10.5772/intechopen.87087*

*Military Engineering*

**Figure 3.** *Supersonic flow [2].*

expands having higher flow area, and flow on the lower surface gets compressed due to lower area of flow. Expansion of flow is associated with increase in Mach number and decrease in static pressure. It may be appreciated that at zero AOA, there will be no difference in upper and lower surface flow. As the supersonic airfoil

generates lifts only due to AOA, supersonic airfoils are symmetrical airfoils. The lift produced increases with increase in AOA as shown in **Figure 4**. However, beyond certain AOA, flow separates from the airfoil, lift suddenly drops, and drag rises due to increase of wake area. This is known as 'stall'. With further increase of AOA, the wake region will increase and thus aggravate the situation. This being a flight safety hazard, airworthiness regulations require stall warning system and protection from stall recovery procedure to be incorporated. Subsonic aircraft stalls around 14–18° of AOA, while supersonic aircraft stalls around 24–28°

**4**

**Figure 4.**

*Stall phenomenon [4].*

AOA. In supersonic flow, lift curve slope is very flat, and realization of stall is rather difficult. In view of this in supersonic aircraft, in addition to an AOA indicator, audio warning with stick shaker is installed.

An aircraft in flight has six degrees of freedom (DOF); these are three translations (motion in forward, lateral and vertical directions) and three angular motions, viz. roll (rotation about longitudinal axis), pitch (rotation around lateral axis) and yaw (rotation about vertical axis). An aircraft is steered to the desired direction by operating the respective control surface. This is shown in **Figure 5**.

The lift, drag and pitching moment produced by the wings can be written as L = \_1 <sup>2</sup> <sup>ρ</sup>V2SCL;D <sup>=</sup> \_1 <sup>2</sup> <sup>ρ</sup>V2SCD; and <sup>M</sup> <sup>=</sup> \_1 <sup>2</sup> <sup>ρ</sup>V2SCCM, where S is the total wetted wing area and C is the mean aerodynamic chord length. The force and moments acting on an aircraft in x-z plane during flight are shown in **Figure 6**. The lift, drag and pitching moments act at the aerodynamic centre of the aircraft. The forces and moments are in equilibrium. If the equilibrium is disturbed, the aircraft will move from its flight path. For example, if the lift is increased more than weight, the aircraft will float up, and if the thrust produced is more than the drag, the aircraft will accelerate. The reverses are also true. Similarly, if any control surface is deflected, ΣM about the axis will be disturbed, and the aircraft will rotate about that axis. The degrees of turn of flight path achieved per degree of deflection of control surfaces are the measure of controllability of the aircraft.

**Figure 5.** *6 DOF of air vehicle [4].*

**Figure 6.** *The force and moment system working on an aircraft in flight [2].*

The ease with which an aircraft can be operated is judged by the handling qualities (HQ ) of an aircraft. The HQ is directly related to the aircraft stability and controllability. An aircraft is said to be statically stable if an aircraft while flying in a steady path, if unintentionally disturbed by some external forces like gust or any other reason, the aerodynamic forces and moments so created due to the disturbances bring the aircraft back to its original stable condition. This property of the aircraft is termed as stability. However, higher the stability, higher will be the demand for control power to steer the aircraft and lower will be the controllability of the aircraft. Thus, a compromise is made, and the stability requirements are specified in design regulations formulated by the airworthiness authorities. The advantages of lower stability have brought the concept of 'relaxed static stability' or 'statically unstable' aircraft. The high-performance military fighters like F-16, F-17 and F-18 are statically unstable in order to obtain dramatic increase in manoeuverability. The vehicle is kept under control by **'**fly-by-wire' (FBW) control system. In FBW, accelerometers and rate gyros are mounted in each axis which senses the aircraft position and attitude, and a FBW computer continuously monitors these data and commands the control actuators to move the control surfaces to keep the aircraft under control. In this approach, very high manoeuverability advantages can be realized without heavily taxing the pilot. Even in the transport aircraft, this leads to smaller tailplane resulting in lesser weight and drag. The light combat aircraft of India, Mirage-2000 and SU-30, belong to this category of unstable aircraft [1–5].
