Open access peer-reviewed chapter - ONLINE FIRST

# Flight Vehicle Performance

By Aram Baghiyan

Submitted: November 25th 2019Reviewed: March 13th 2020Published: May 4th 2020

DOI: 10.5772/intechopen.92105

## Abstract

In this chapter, the problem of flight vehicle performance is described. Performance parameters, such as lift-to-drag ratio, maximum and minimum level flight speed, speeds for the best rate of climb, steepest climb, maximum range and endurance, and most economical climb are described using graphical methods, such as drag polar, Zhukovsky curves with combination of analytical derivations. The approach of graphical description of flight vehicle performance allows to understand the physical basics of the aerodynamic properties of flight vehicles easier and to develop deeper connectivity between their interpretations. In addition, flight envelope and operational limits are discussed using both analytical and graphical methods for better understanding.

### Keywords

• aerodynamics
• flight vehicle
• drag polar
• flight performance
• Zhukovsky curves

## 1. Introduction

During the flight, aircraft is usually loaded by four forces (Figure 1)—gravity force G, lift L, drag D,and thrust T. The combination of these forces defines the behavior of flight and the performance of aircraft. There are also cases when the side forces Sact on aircraft due to sideslip (Figure 1), which are balanced by the vertical tail of aircraft, and during the analysis of the flight vehicle performance they are usually neglected.

As we can see from the Figure 1, the vector of aircraft speed vdoes not coincide with its body frame axes xb,yb,and zband has inclinations from axis xbexpressed by flow angles—angle of attack αand sideslip angle β. The point of acting gravity force is the center of gravity Ocgand usually the line of acting of thrust passes through the center of gravity to avoid generation of destabilizing torques or moments. Aerodynamic forces act at the point named center of pressure Ocp, which is not a fixed and changes its position depending on the angle of attack and the air speed. Therefore, a more stable point—aerodynamic center Oac—is introduced, where the changes of aerodynamic forces act, so the aerodynamic moments at that point do not change with the changes of angle of attack. However, aerodynamic center can have variations [1] with Mach number M=v/a,where vis the magnitude of aircraft speed, ais the local speed of sound.

In problems of flight vehicle performance analysis, usually a simpler model of the balance of forces is used: the side forces Sare neglected due to small values of sideslip angle βin steady flight regimes, the point of action of all unneglectable forces assumed the same, at the same time, taking the lines of action of thrust Tand drag Dforces coinciding (Figure 2).

Based on the model of force balance, we will study flight vehicle performance at several steady flight regimes, but before getting there let us consider the aerodynamic forces—lift and drag, and the effect of their relations on the aerodynamic quality of aircraft.

## 2. Drag polar of aircraft

Taking as a reference any paper on aerodynamics, we can find the formulas for lift Land drag Dforces [2]:

L=cLρv22Sref,D=cDρv22Sref,E1

where cLis the dimensionless lift coefficient; ρis the freestream density; vis the magnitude of freestream speed, which is taken equal to aircraft speed; Srefis the reference area, and cDis the dimensionless drag coefficient.

From the system of Eq. (1)

LD=cLcD.

The ratio k=L/Dis usually called lift-to-drag ratio and in eastern literature it is defined as aerodynamic quality of aircraft. This ratio has interesting properties and it changes with changes of angle of attack α.

Let us consider now the lift coefficient cLand its relations with angle of attack α(Figure 3), which is similar to the analogous relation for 2-D airfoils. The function cL=fLαcan be found using experiments with aircraft model in wind tunnels or using methods of computational fluid dynamics (CFD).

As we can see from the graph of function cL=fLα(Figure 3) there is an angle of attack αcrat which the lift coefficient is maximal, that is, cLmax. The flight at critical angle of attack αcrwill lead to stall, resulting aircraft crash. At angle of attack αts, the tip stall processes are started and the rate of increase of lift coefficient is decelerated, allowing it to get its maximum value cLmaxand decrease sufficiently. At the point αtsctsof starting tip stall, the effects of shaking of aircraft are started [3]. The range between angles of attack α0, which is called zero lift angle of attack, and αtsis the range of regular flights of aircraft and as it can be observed at this range the function cL=fLαis linear.

Before getting to the drag coefficient study, let us consider the existing types of the drag force. The drag force Dcan be represented as a sum of parasitic drag Dp, which consists of form drag Dfand skin friction drag Dsf[4], lift-induced Didrag, wave drag Dw, and interference drag Dif:

D=Df+Dsf+Di+Dw+Dif,

or, which is the same as

cDρv22Sref=cDp+cDi+cDw+cDifρv22Sref,

where cDp=cDf+cDsfis the parasitic drag coefficient, cDfis the form drag coefficient, cDsfis the skin friction drag coefficient, cDiis the lift-induced drag coefficient, cDwis the wave drag coefficient, and cDifis the inference drag coefficient.

Parasitic drag is the pressure difference in front of and behind the wing. The pressure difference depends on the shape of the wing airfoil, its relative thickness c¯and curvature. The larger the relative thickness of the wing airfoil, the greater the form drag (also known as pressure drag); on the other hand, the lower the relative thickness of the wing airfoil, the greater the effect of skin friction drag [5] (Figure 4).

The lift-induced drag is the result of the flow tilt (Figure 5). Due to the pressure difference above and under the wing on its tips, vortices are generated, leading to the downwash of air from upper surface with velocity u. Thus the effective flow speed veffbecomes the vector sum of the freestream air speed vand downwash speed u. The direction of the effective flow speed differs from the freestream velocity’s direction by angle δα, so the effective angle of attack αeffis defined as:

αeff=α+δα.

With an increase in the angle of attack or lift coefficient, the pressure difference under and above the wing increases quickly, and the coefficient of lift-induced drag increases according to the quadratic law [2]:

cDi=cL2πλe,E2

where λ=b2/Srefis the aspect ratio, bis the wing span, Srefis the wing reference area, and eis the span efficiency.

The wave drag Dwis a consequence of the compressibility of the air and occurs when there are shock waves near the aircraft.

The mutual influence of the parts of the aircraft is called interference. It occurs due to a change in the velocity field, as a result of which the nature of the flow around the aircraft changes leading to generation of interference drag Dif.

Based on the review of drag components, we can divide them into components related to the lift generation or lift-induced drag, and components not related to the lift generation:

cD=cD0+cDi,E3

where cD0is the component of drag coefficient not related with the lift generation and is called zero lift drag coefficient. Usually cD0is taken as constant and not related to the angle of attack, while cDiis proportional to the square of lift coefficient cL, which linearly depends on angle of attack αin range of the regular flight regimes. Thus the function cD=fDαshould have the graph of parabolic form (Figure 6).

Similar to cL=fLα, function cD=fDαcan be found using experiments with aircraft model in wind tunnels and using CFD tools.

Using estimated or experimental results for fLαand fDαat set of angles of attack, it is possible to draw drag polar of aircraft (Figure 7).

Based on the drag polar conditions for the best lift-to-drag ratio, zero lift drag and maximal lift can be found. By drawing tangent 1 to the curve of drag polar from the origin of coordinate frame cDcL, the best lift-to-drag ratio can be found, which corresponds to the angle of attack αbldr; the tangent 2 to the drag polar parallel to axis cDdefines maximal value of lift coefficient cLmaxat critical angle of attack αcr, and tangent 3 to the drag polar parallel to axis cLdefines zero lift drag coefficient cD0at zero lift angle of attack α0.

The angle of attack αbldrof best lift-to-drag ratio has sufficient role in the flight of aircraft, as the flight with this angle provides the maximum value of the mentioned ratio, also known as aerodynamic quality of aircraft:

kmax=fLαbldrfDαbldr.

Let us consider Eq. (3) taking into account Eq. (2):

cD=cD0+cL2πλe.

The above expression is also called drag polar equation, with the use of which we can represent the non-negative values of lift coefficient as

cL=πλecDcD0.

Taking cDderivative of cLwe get:

cLcD=πλecDcD0cD=πλe2πλecDcD0.

At point cDαbldr, this derivative is the same as the slope kmaxof tangent 1 from Figure 7, so we can write down the following expression:

πλe2πλecDαbldrcD0=cLαbldrcDαbldr=πλecDαbldrcD0cDαbldr,

which can be easily transformed to

cDαbldr=2cD0.E4

Based on the Eq. (4) we can find the maximal value of lift-to-drag ratio:

kmax=12πλecD0.

To examine the dependency of the lift-to-drag ratio on angles of attack, the graph of the function kα=fLα/fDαcan be plotted (Figure 8) and the range of angles near the αbldrstudied to find the effective flight regimes and patterns [3, 6].

Examined material is one of core bases of aircraft performance, and the results obtained through the above analysis are used in studies of different flight paths and patterns and will be referred in next subsection dedicated to the Zhukovsky curves.

## 3. Zhukovsky curves

Let us now consider steady horizontal flight. The scheme on Figure 2 will be transformed to the following form (Figure 9):

In steady horizontal flight, we have the following equation of the force balance:

L=G,T=D.E5

which is same as:

cLρv22Sref=G,T=cDρv22Sref=cD0+cL2πλeρv22Sref.

From the first equation of the above system, we can find that

cL=2Gρv2Sref,

and by substituting the value cLin the second equation we get:

T=cD0ρv22Sref+1πλe2G2ρv2Sref.

Based on the above result, we can state that the required thrust Trfor the steady horizontal flight should be equal to the sum of zero lift drag D0and lift-induced drag Di, which are defined as:

Tr=D0+Di,D0=cD0ρv22Sref,Di=cDiρv22Sref=1πλe2G2ρv2Sref.

As we can see, zero lift drag D0is proportional to the square of the air speed, while lift-induced drag Diis inversely proportional to the square of air speed.

Let us now define the conditions of minimal drag or, which is the same as, minimal required thrust at steady horizontal flight:

Trv=2cD0ρv2Sref21πλe2G2ρv3Sref=0.

The above expression can be rewritten as:

cD0ρv22Sref=1πλe2G2ρv2Sref.E6

It is obvious that the left-hand side of the above is zero lift drag D0and the right-hand side is lift-induced drag Di:

D0=Di,

or

cD0=cDi.

Thus, the drag coefficient is equal:

cD=cD0+cDi=2cD0.

As we remember from the drag polar, the same condition is true for maximal lift-to-drag ratio, so the conditions for minimal required thrust and maximal lift-to-drag ratio are the same. To complete the calculations of all parameters for minimal required thrust, let us derive the expressions for lift coefficient and air speed:

cDi=cD0=cL2πλe,

so,

cL=πλecD0,

and from Eq. (6)

vL/Dmin=1πλecD04G2ρ2Sref24.

We can also find the characteristics of available thrust Taprovided by manufacturers of engines or the estimations of available thrust from the sources of literature. In [7, 8], the forms of dependency of available thrust on the air speed for several types of engines are presented. Particularly, in [3] we can find available thrust versus air speed at several altitudes for the jet aircraft L-39 (Figure 10).

Based on the available information, Nikolay Zhukovsky developed a graphical method for the analysis of the range of the horizontal flight speeds at different altitudes. His method is based on the plotting curves of zero lift drag and lift-induced drag versus air speed at different altitudes, graphically calculating their sum, and plotting the dependencies of available thrust on air speed at corresponding altitudes for graphoanalytic estimation of ranges of available air speeds for horizontal flight at different altitudes. The set of curves obtained through the above-described procedure in memory of him are called Zhukovsky curves. The graph with curves of required and available thrusts or powers at certain altitude is also called performance diagram.

Let us implement the method proposed Zhukovsky for any altitude, for example, do a plot for the altitude of zero meters above sea level (Figure 11) and define the speed characteristics of the horizontal flight.

From Zhukovsky curves at sea level altitude, we can find the maximum speed vmaxcorresponding to the right point of intersection of curves for required thrust Trvand available thrust Tav. At speed vL/Dminfor which zero lift drag becomes equal to lift-induced drag, D0vL/Dmin=DivL/Dmin, we get minimal required thrust Trmin=2D0vL/Dmin=2DivL/Dmin. This speed is a very important quantity from the point of horizontal flight, as the flight at lower speeds is not stable to the accidental changes of speed and requires attention of pilots. On the other hand, the flight at higher speeds than vL/Dminis stable to accidental changes of speed, as if speed decreases drag force also decreases and as the thrust was at initial value it accelerates the aircraft till there is a balance between thrust and drag. We get a similar picture for accidental increase of speed at stable region II, for which the drag force decelerates the aircraft till drag-thrust balance. For unstable region I, the dangerous case is the accidental decrease of speed, which increases drag leading to decelerating the aircraft till stall. To fly at unstable region of speeds, the pilot needs always to work with the throttle to increase thrust when it is required. Minimum speed vminat sea level flight is not defined from the above curves and refers to the stall speed, which can be found from the condition of required lift:

cLmaxρv22Sref=G,

or

vmin=2GcLmaxρSref.

For flight at higher altitudes, we can get conditions when the required thrust at the above minimum speed is much higher than the available thrust at that altitude (Figure 12). In such cases, minimum speed defined as minimum thrust speed vTmincorresponds to the left point of intersection of curves for required thrust Trvand available thrust Tav.

The idea of plotting Zhukovsky curves at sea level flight allows us to have the same graph at any altitude Hfor the required trust TrvIASversus indicated airspeed (IAS), which is usually measured on aircraft and related to the true airspeed (TAS) vTASvia the expression:

vIAS=vTASρHρ0,

where ρ0is the air density at sea level and ρHis the air density at altitude Habove sea level (ASL).

Based on the Zhukovsky curves based on IAS, we can define the theoretical or static ceiling of horizontal flight (Figure 12), which is an altitude where the horizontal flight is possible only with IAS equal to vL/Dmin. The service ceiling has a more practical meaning, as it is the altitude where rate of climb (ROC) becomes less than 0.5 m/s [6, 7].

There are also defined such concepts as speed of maximum endurance ve, cruise speed vcr, and speed of maximum range vr. The maximum endurance speed and maximum range speed depend on fuel consumption characteristics of engine. The speed corresponding to the minimum hourly fuel consumption of engine is called speed of maximum endurance or economical speed. On the other hand, the speed corresponding to the minimum per-kilometer consumption of engine is the speed of maximum range and it is very close to the cruise speed (slightly more than cruise speed for real aircraft). The cruise speed is the speed at which the ratio of drag to speed is minimal, and can be found using Zhukovsky curves by drawing a tangent to the required thrust graph from the origin (Figure 13).

Let us now consider descending flight or glide (Figure 14a) and ascending (Figure 14b) flight of aircraft.

In descending flight, the throttle is usually set to minimum, so the thrust can be neglected and consider the gliding flight (Figure 14a), for which we can write down the following equations of force balance:

L=Gcosθ,D=Gsinθ.

where θis the flight path angle.

Based on the above we can get:

tanθ=DL=1k.

The flattest glide corresponds to the minimum magnitude of flight path angle θmin, which is case when the tanθis minimum, thus, we can write down the following:

θmin=atan1kmax.

The above means that the flattest glide, resulting the longest gliding distance, is also related to maximum lift-to-drag ratio kmax.

For the ascending flight the force balance is presented as follows:

L=Gcosθ,TD=Gsinθ.E7

As we know, the rate of climb is the projection of airspeed to the vertical plane:

ROC=vsinθ.

From second equation of Eq. (7) we can get:

sinθ=TDG,

or which is same as:

ROC=vsinθ=TvDvG=PaPrG,

where Pa=Tvis the available power, Pr=Dvis the required power.

From the above we can find the maximum of ROC, that is, ROCmaxcorresponding to maximum of excess power ΔPmax=PaPrmax(Figure 15).

The angle θmaxof steepest ascending flight can be found by dividing the ROC value by speed [9]:

θmax=asinROCvmax,

which corresponds to

θmax=asinTaTrmaxG.

In ideal jet aircraft case, the angle of steepest climb is defined at the speed of maximum lift-to-drag ratio. The most economical ascending flight is defined by the operational regime of aircraft engine at which the minimum fuel will be consumed for aircraft to climb to the required altitude.

## 4. Flight envelope and operational limits of aircraft

For any aircraft, performance and operational limits are defined. Performance limits (Figure 16) are mostly defined by the aerodynamic configuration of aircraft. On the other hand, operational limits are based on the type of aircraft, structural and engine limits, wind resistance parameters, and maximum Mach number. All these limits are presented in diagrams of altitude hversus airspeed vand the final diagram, which represents the intersection of all limits, is called flight envelope (Figure 17). The flight envelope can be represented in graphs of altitude versus true airspeed, altitude versus indicated airspeed, or altitude versus Mach number.

All the speeds presented on Figure 16 were described in the previous subsection. Operational limits related to the type of aircraft are described from the point of view of aircraft application: if an aircraft is a passenger jet, it should apply to the requirements of comfort for passengers and not exceed load factors of comfortable flight; on the other hand, if an aircraft is a jet fighter its operational limits from the point of view of load factors should be derived from compromise between the prevention of health issues of pilot that may occur and maneuverability for the combat use.

Structural limits are mostly related to aircraft strength, while the engine limits can be the result of its design and performance at higher altitudes. Wind resistance limits can be derived from the requirements of operational use or comfortability of flight for passengers.

Maximum Mach number can be defined from the conditions of aeroelastic effects and vibrations, effects of shifting aerodynamic center, causing severe pitching moments, which can lead to the crashes, or loosing effectiveness of aerodynamic surfaces. For example, for jet aircraft L-39 the critical Mach number is M=0.8and exceeding this condition leads to the shift of aerodynamic center of L-39, which generates pitching moment, causing a descending flight with acceleration [3]. To prevent such unstable flight, on L-39 air brakes are used, which automatically act at the Mach numbers M=0.78. The critical Mach number is defined at airspeed flight that leads to generation of shock waves on the wing due to the acceleration of airflow on the upper surface of wing.

Finally, an example of flight envelope for analysis is presented in Figure 17 where the intersection of all limiting conditions is described.

## 5. Conclusion

The material described in this chapter involves flight vehicle analysis using graphical and analytical tools for better understanding of the physical aspects of flight core parameters and development of strong and meaningful connections between them. The material from this chapter can be useful in the preliminary design and prototyping of flight vehicles and for finding the paths for further developments and improvements in the design.

## Acknowledgments

In this work, the name of Prof. Nikolay Zhukovsky (January 17, 1847–March 17, 1921) is mentioned many times, while his name may be unfamiliar to many readers. Nikolay Zhukovsky was one of the first scientists who established the mathematical base of aerodynamics; he was the initiator and the first head of the Central Aerohydrodynamic Institute (TsAGI), and on his initiative was created Zhukovsky Air Force Engineering Academy—alma mater of three famous Soviet aircraft designers Sergei Ilyushin, Artem Mikoyan, and Alexander Yakovlev.

Science lives in the research schools, and great scientists are those who both do great discoveries and develop the next generations of discoverers. Zhukovsky Air Force Engineering Academy is my alma mater as well, and with these few words I would like to express my appreciation to all my teachers who helped me on my way of professional and personal development.

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Aram Baghiyan (May 4th 2020). Flight Vehicle Performance [Online First], IntechOpen, DOI: 10.5772/intechopen.92105. Available from: