Chapter 6
Download
1 / 93

Chapter 6 - PowerPoint PPT Presentation


  • 154 Views
  • Uploaded on

Chapter 6. Elements of Airplane Performance. Un-accelerated level flight. Simple Mission Profile for an Airplane 1 Switch on + Worming + Taxi. (Cruising flight). 4. 3. Descent. Altitude. Climb. Landing. Takeoff. 5. 6. 1. 2. Simple mission profile. Airplane Performance.

loader
I am the owner, or an agent authorized to act on behalf of the owner, of the copyrighted work described.
capcha
Download Presentation

PowerPoint Slideshow about ' Chapter 6' - logan-hansen


An Image/Link below is provided (as is) to download presentation

Download Policy: Content on the Website is provided to you AS IS for your information and personal use and may not be sold / licensed / shared on other websites without getting consent from its author.While downloading, if for some reason you are not able to download a presentation, the publisher may have deleted the file from their server.


- - - - - - - - - - - - - - - - - - - - - - - - - - E N D - - - - - - - - - - - - - - - - - - - - - - - - - -
Presentation Transcript
Chapter 6

Chapter 6

Elements of Airplane Performance

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Un-accelerated level flight

Simple Mission Profile for an Airplane

1 Switch on + Worming + Taxi

(Cruising flight)

4

3

Descent

Altitude

Climb

Landing

Takeoff

5

6

1

2

Simple mission profile

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Airplane Performance

Equations of Motions

Static Performance

(Zero acceleration

Dynamic Performance

(Finite acceleration)

Thrust required

Thrust available

Maximum velocity

Takeoff

Power required

Power available

Landing

Maximumvelocity

Rate of climb

Gliding flight

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Time to climb

Maximum altitude

Service ceiling

Absolute ceiling

Range and endurance

Road map for Chapter 6

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • Study the airplane performance requires the derivation of the airplane equations of motion

  • As we know the airplane is a rigid body has six degrees of freedom

  • But in case of airplane performance we are deal with the calculation of velocities ( e.g.Vmax,Vmin..etc),distances (e.g. range, takeoff distance, landing distance, ceilings …etc), times (e.g. endurance, time to climb,…etc), angles (e.g.climb angle…etc)

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • So, the rotation of the airplane about its axes during flight in case of performance study is not necessary.

  • Therefore, we can assume that the airplane is a point mass concentrated at its c.g.

  • Also, the derivation of the airplane’s equations of motion requires the knowledge of the forces acting on the airplane

  • The forces acting on an airplane are:

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Components of the resultant flight in case of performance study is not necessary.

aerodynamic force R

  • 1- Lift force L

  • 2- Drag force D

  • 3- Thrust force T Propulsive force

  • 4- Weight W Gravity force

  • Thrust T and weight W will be given

  • But what about L and D?

  • We are in the position that we can’t calculate L and D with our limited knowledge of the airplane aerodynamics

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Total Drag of the so called

■ Drag Types [ Kinds of Drag ]

Skin Friction Drag

Pressure Drag

Form Drag (Drag Due to Flow separation)

Induced Drag

Wave Drag

Note : Profile Drag = Skin Friction Drag + Form Drag

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


►Skin friction drag of the so called

This is the drag due to shear stress at the surface.

►Pressure drag

This is the drag that is generated by the resolved components of the forces due to pressure acting normal to the surface at all points and consists of [ form drag + induced drag + wave drag ].

►Form drag

This can be defined as the difference between profile drag and the skin-friction drag or the drag due to flow separation.

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


►Profile Drag of the so called

● Profile drag is the sum of skin-friction and form drags.

● It is called profile drag because both skin-friction and

form drag [ or drag due to flow separation ] are

ramifications of the shape and size of the body, the

“profile” of the body.

● It is the total drag on an aerodynamic shape due to

viscous effects

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Skin-friction of the so called

Form drag

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


►Induced drag ( or vortex drag ) of the so called

This is the drag generated due to the wing tip vortices , depends on lift, does not depend on viscous effects , and can be estimated by assuming inviscid flow.

Finite wing flow tendencies

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Formation of wing tip vortices of the so called

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Complete wing-vortex system of the so called

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University



The origin of downwash of the so called

The origin of induced drag

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


►Wave Drag of the so called

This is the drag associated with the formation of shock waves in high-speed flight .

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


■ Total Drag of the so called ofAirplane

● An airplane is composed of many components and each will contribute to the total drag of its own.

● Possible airplane components drag include :

1. Drag of wing, wing flaps = Dw

2. Drag of fuselage = Df

3. Drag of tail surfaces = Dt

4. Drag of nacelles = Dn

5. Drag of engines = De

6. Drag of landing gear = Dlg

7. Drag of wing fuel tanks and external stores = Dwt

8. Drag of miscellaneous parts = Dms

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


● Total drag of the so called of an airplane is not simply the sum of the drag of the components.

● This is because when the components are combined into a complete airplane, one component can affect the flow field, and hence, the drag of another.

● these effects are called interference effects, and the change in the sum of the component drags is called interference drag.

● Thus,

(Drag)1+2 = (Drag)1 + (Drag)2 + (Drag)interference

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


■ Buid-up Technique of Airplae Drag D of the so called

● Using the build-up technique, the airplane total drag D is expressed as:

D = Dw + Df + Dt + Dn +De + Dlg + Dwt + Dms + Dinterference

► Interference Drag

● An additional pressure drag caused by the mutual interaction of the flow fields around each component of the airplane.

● Interference drag can be minimized by proper fairing and filleting which induces smooth mixing of air past the components.

● The Figure shows an airplane with large degree of wing filleting.

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Wing fillets of the so called

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


● No of the so called theoretical method can predict interference drag, thus, it is obtained from wind-tunnel or flight-test measurements.

● For rough drag calculations a figure of 5% to 10% can be attributed to interference drag on a total drag, i.e,

Dinterference ≈ [ 5% – 10% ] of components total drag

■ The Airplane Drag Polar

● For every airplane, there is a relation between CD and CL that can be expressed as an equation or plotted on a graph.

● The equation and the graph are called the drag polar.

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


For the complete airplane, the drag coefficient is written as

CD = CDo + K CL2

This equation is the drag polar for an airplane.

Where: CDo drag coefficient at zero lift ( or

parasite drag coefficient )

K CL2 = drag coefficient due to lift ( or

induced drag coefficient CDi )

K = 1/π e AR

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


e Oswald efficiency factor = 0.75 – 0.9 as

(sometimes known as the airplane efficiency factor)

AR wing aspect ratio = b2/S ,

b wing span and S wing planform area

Schematic of the drag polar

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Airplane Equations of Motion as

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University




  • Apply Newton’s 2 as nd low of motion:

    In the direction of the flight path

    Perpendicular to the flight path

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


I steady level flight performance

I-Steady as Level Flight Performance

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Un-accelerated (steady) Level Flight Performance (Cruising Flight)

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


● T Flight)R as a function of V∞ can be obtained by tow methods

or approaches graphical/analytical

■Graphical Approach

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


1- Choose a value of V Flight)∞

2 - For the chosen V∞ calculate CL

L = W = ½ρ∞V2∞S CL

CL = 2W/ ρ∞V2∞S

3- Calculate CD from the drag polar

CD = CDo + K CL2

4- Calculate drag, hence TR, from

TR = D = ½ρ∞V2∞S CD

5- Repeat for different values of V∞

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


6- Tabulate the results Flight)

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


(T Flight)R)min occurs at (CL/CD)max

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • Flight)Analytical Approach

  • It is required to obtain an equation for TR as a function of V∞

  • TR = D

Required equation

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


TR/D

CDo=CDi

V∞

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • Note that T Flight)R is minimum at the point of intersection of the parasite drag Do and induced drag Di

  • Thus Do = Di at [TR]min

  • or CDo = CDi

  • = KCL2

  • Then [CL](TR)min = √CDo/K

  • And [CDo](TR)min = 2CDo

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • Finally, (L/D) Flight)max = (CL/CD)max

  • = √CDo/K /2CDo

  • (CL/CD)max = 1/√4KCDo

  • Also,[V∞](TR)min =[V∞](CL/CD)maxisobtained from: W = L

  • = ½ρ∞[V]2(TR)minS [CL](TR)min

  • Thus:

  • [V](TR)min= {2(W/S)(√K/CDo)/ρ∞}½

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


L/D as function of angle of attack Flight)α

L/D as function of velocity V∞

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • Flight Velocity for a Given T Flight)R

  • TR = D

  • In terms of q∞ = ½ρ∞V2∞we obtain

  • Multiplying by q∞and rearranging, we have

  • This is quadratic equation in q∞

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • Let Flight)

  • Where (TR/W) is the thrust-to-weight-ratio

  • (W/S) is the wing loading

  • The final expression for velocity is

  • This equation has two roots as shown in figure corresponding to point 1 an 2

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Flight)When the discriminant equals zero ,then only

one solution for V∞ is obtained

●This corresponds to point 3 in the figure,

namely at (TR)min

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • Or, (T Flight)R/W)min = √4CDoK

  • Then the velocity V3 =V(TR)min is

  • Substituting for (TR/W)min = √4CDoK we have

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • Effect of Altitude on (T Flight)R)min

  • We know that

  • (TR/W)min = √4CDoK

  • This means that (TR)min is independent of altitude as show in Figure

  • (TR)min occurs at higher V∞

V∞1

V∞2

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Thrust Available T Flight)A

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Sonic speed Flight)

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Thrust Available T Flight)A and Maximum Velocity Vmax

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Prof. Galal Bahgat Salem Aerospace Dept. Cairo University



C Flight)D= 4CDo

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Prof. Galal Bahgat Salem Aerospace Dept. Cairo University



  • Power Available P Flight)A and Maximum Velocity Vmax

  • The high speed

    intersection

    between PR and

    (PA)max gives

    Vmax

  • Vmax decreases

    with altitude

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University



  • Minimum Velocity: Stall Velocity Flight)

  • Airplane minimum velocity Vmin is usually dictated by its stall velocity

  • Stall velocity Vstall is the velocity corresponds to the maximum lift coefficient (CL)maxof the airplane

  • Hence, Vmin = Vstall

  • But, L = W = ½ρ∞V2∞S CL

    V∞ = (2W/ ρ∞ S CL )½

  • Substitute for CL (CL)max

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • Finally, Flight)

    Vmin= Vstall = [2W/ ρ∞ S (CL)max ]½

CL –α curve for an airplane

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Ii steady climb performance

II-Steady Climb Performance Flight)

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • Steady Climb Flight)

  • Assumptions:

    1- dV∞/dt = 0

    2- Climb along straight line, V2∞/ r = 0

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Prof. Galal Bahgat Salem Aerospace Dept. Cairo University




[Turbojet] Flight)

,for T = constant

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


sin Flight)

Turbojet aircraft

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University



Turbojet aircraft Flight)

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • Analytical Solution for (R/C) Flight)max

  • R/C = V∞ sin ϴ

  • = (2W/ ρ∞ S CL )½ [ T/W- D/L]

  • = (2W/ ρ∞ S CL )½ [T/W-CD/CL]

  • = (2W/ ρ∞ S CL )½ [T/W-CDo +KCL2/CL]

  • =(2W/ ρ∞ S )½ [CL-½(T/W)-(CDo+KC2L)/CL3/2]

  • For turbojet T = const

  • For (R/C)max d(R/C)/dCL =0

  • So, we get

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


C Flight)L(R/C)max = [ -(T/W) + √ (T/W)2 + 12 K CD0 ] / 2K

  • So, we get:

  • And, V(R/C)max=[2W/ ρ∞ S CL(R/C)max ]½

  • (CD) (R/C)max=CDo +K C2L(R/C)max

  • (Sin ϴ) (R/C)max = T/W- (CD/CL) (R/C)max

  • (R/C)max = V(R/C)max (sin ϴ) (R/C)max

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Propeller aircraft

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • Analytical Solution for (R/C) Flight)max

  • V(R/C)max= V(CL3/2/CD)max

  • (CD) (R/C)max=CDo +K C2L(R/C)max

  • = CDo +K [√3CDo/K ]2 = 4CDo

  • (Sin ϴ) (R/C)max = T/W- (CD/CL) (R/C)max

  • (R/C)max = V(R/C)max (sin ϴ) (R/C)max

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University




Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


Flight)If PR˃ PA the airplane will descend

 In the ultimate situation when T = 0, the

airplane will be in gliding

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • Maximum Range Flight)

  • For an airplane at a given altitude h, the max. horizontal distance covered over the ground is denoted max. range R

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • For R Flight)maxϴmin

  • Where:

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


CEILINGS Flight)

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


max Flight)

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


(R/C) Flight)-1

h

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


  • Minimum Time to Climb Flight)

  • tmin =

  • Assuming linear variation of (R/C)maxwith altitude h, then

  • (R/C)max = a + b h

  • a = (R/C)max at h = 0

  • =1/b[ln(a+bh2)-lna]

max

h

b =slope

0

(R/C)max

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University


III-Range and Endurance Flight)

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University



W=Instantaneous weight Flight)

Prof. Galal Bahgat Salem Aerospace Dept. Cairo University







ad