2 aircraft flight instruments

Aircraft Flight Instruments

SOLO HERMELIN

Updated: 04.12.12

1

Table of Content

SOLO Aircraft Avionics

2

McDonnell Douglass F-4B Phantom Instrument Panel 3

http://forum.combatace.com/index.php?automodule=downloads&showfile=3395

McDonnell Douglass F-4 Phantom Cockpit4

5

Earth Atmosphere

6

Earth Atmosphere

7

Earth Atmosphere

The basic variables representing the thermodynamics state of the gas are the Density, ρ, Temperature, T and Pressure, p.

SOLO

8

Earth Atmosphere

• The Density, ρ, is defined as the mass, m, per unit volume, v, and has units of kg/m3.

v

mv

0lim

• The Temperature, T, with units in degrees Kelvin ( 1 K). Is a measure of the average kinetic energy of gas particles.

• The Pressure, p, exerted by a gas on a solid surface is defined as the rate of change of normal momentum of the gas particles striking per unit area.

It has units of N/m2. Other pressure units are millibar (mbar), Pascal (Pa), millimeter of mercury height (mHg)

S

fp n

S

0

lim

kPamNbar 100/101 25

mmHginHgkPamkNmbar 00.7609213.29/325.10125.1013 2 The Atmospheric Pressure at Sea Level is:

Speed of Sound (a)This is the speed of sound waves propagation in ambientair. The speed of sound is given by

SOLO

9

Earth Atmosphere

Sa TRa

γ air = 1.4 Ra =287.0 J/kg--1K TS – Static Air Temperature

True Airspeed (TAS)The True Airspeed is the speed of the aircraft’s center of mass with respect to the ambient air through which is passing.

Indicated Airspeed (IAS)The Indicated Airspeed is the speed indicated by a differential-pressure airspeed indicator.

Mach Number (M)Is the ratio of the TAS to the speed of sound at the flight condition.

SOLO

10

Earth Atmosphere

aTASM /

Dynamic Pressure (q)The force per unit area required to bring an ideal(incompressible) fluid to rest: q=1/2∙ρ∙VT

2 (where VT is True Air Speed-TAS, and ρ is the density of the fluid).

Impact Pressure (QC) The force per unit area required to bring moving air to rest. It is the pressure exerted at the stagnation point on the surface of a body in motion relative to the air.

PT – Total Pressure, PS – Static Pressure

22/1 TSTC VPPQ


11

Air Data Computer

Air Data Computer uses Total and Static Pressure and Static Temperature of the external Air Flow, to compute Flight Parameters.

SP

Pitot StaticProbe

Total PressureSensor/Transducer

Static PressureSensor/Transducer

Total TemperatureProbe

TP

mT

Airspeed Indicator(ASI)

Vertical Speed Indicator

(VSI)

Altimeter

Air DataComputer

(ADC)

PressureAltitude, H

VerticalSpeed, H

CalibratedAirspeed, CV

Mach Number

True Aispeed,TV

Static AirTemperature ST

Air DensityRate S /

12

Earth Atmosphere

Atmospheric Constants

DefinitionSymbolValueUnitsSea-level pressureP01.013250 x 105N/m2

Sea-level temperatureT0288.15ͦ1 K

Sea-level densityρ01.225kg/m3

Avogadro’s NumberNa6.0220978 x 1023/kg-moleUniversal Gas ConstantR*8.31432 x 103J/kg-mole -1 KGas constant (air)Ra=R*/M0287.0J/kg--1K

Adiabatic polytropic constantγ1.405Sea-level molecular weightM028.96643

Sea-level gravity accelerationg09.80665m/s2

Radius of Earth (Equator)Re6.3781 x 106m

Thermal Constantβ1.458 x 10-6Kg/(m-s-1 K1/2)

Sutherland’s ConstantS110.4ͦ1 KCollision diameterσ3.65 x 10-10m


13

Flight InstrumentsAir Data Calculation (Collison)

Geopotential Pressure Altitude

• Low Altitude (Troposphere) : H< 11000 m (36.089 ft ),

kPaHPS255879.551025577.21325.101

• Medium Altitude: 11000 m ≤ H ≤ 20000m (36.089 ft - 65.617 ft )

kPaeP HS

000,1110576885.1 4

6325.22

Air Density Ratio ρ/ρ0

S

S

T

P

35164.00


14


Mach Number

• Subsonic Speeds (M ≤ 1),

2/722.01 MP

P

S

T

• Supersonic Speeds (M ≥ 1),

Static Air Temperature TS 1 K

102.01 2

rMr

TT m

S

2/52

7

17

9.166

M

M

P

P

S

T

True Airspeed (TAS) VT m/s

smTMV ST /0468.20


15


Speed of Sound a m/s

• Subsonic Speeds (VC ≤ a),

• Supersonic Speeds (VC ≥ a),

Sa TRa γ air = 1.4, Ra =287.0 J/kg--1K

Calibrated Airspeed (CAS) VC m/s

kPaV

Q CC

1

294.3402.01325.101

2/72

kPaV

V

Q

C

C

C

1

1294.340

7

294.34092.166

325.101

2/7

2/52

2

16Central Air Data Computer

Earth Atmosphere

ComputePressure Error

Correction

TotalPressure

PS Ind

Correct PT

Correct PS

StaticPressure

PT Ind

Pressure Altitude HP

Calibrated Airspeed VC

MachNumber M

True Airspeed VT

Static Air Temperature

T

ComputePT/PS

ComputeQC=PT-PS

Total/Indicated AirTemperature

Ti Ind

RecoveryRatio r

QC

PT/PS

PS

H

M ComputePressure Altitude

ComputeCalibrated Airspeed

ComputeMach Number

ComputeTrue Airspeed

ComputeStatic

TemperatureTi

M

r

M

T

∆PT

∆PS

17Central Air Data Computer

Earth Atmosphere

Flight Instruments


18The Flight Panel - Understand Your Aircraft, Youtube

Flight Instruments


19

Flight InstrumentsSOLO Aircraft Avionics

20


21

zdgpd

TRp KsmR 22 /287

zdaTd

aR

g

T

za

p

p

00

1


22

Altimeter


23

Flight InstrumentsAirspeed Indicators

SOLO Aircraft AvionicsFlight InstrumentsAirspeed Indicators

Centurion T210 AirSpeed Gauge


25


2

2

1vpp StatTotal

The airspeed directly given by the differential pressure is called Indicated Airspeed (IAS). This indication is subject to positioning errors of the pitot and static probes, airplane altitude and instrument systematic defects. The airspeed corrected for those errors is called Callibrated Airspeed (CAS).Depending on altitude, the critic airspeeds for maneuvre, flap operation etc change because the aerodynamic forces are function of air density. An equivalent airspeed VE (EAS) is defined as follows:

0

VVE

V – True Airspeedρ – Air Densityρ0 – Air Density at Sea Level


26


2

2

1VPQPP StatCStatTotal

V – True Airspeedρ – Air Densityρ0 – Air Density at Sea Level

Air Density changes with altitude. Assuming an Adiabatic Flow, the relation between Pressure and Density is given by

constCP

γ = Cp/CV= 1.4 for air

Momentum differential equation for the Air Flow is

VdVC

PPdVdVPd

C

P

/1

/1

0

Subsonic Speeds

StaticOrifice

Pitot TubeT

S

VV

PP

T

S

VV

PP

0

V

PP T

SP TP

PressureTransducers

IASAirspeed Indicator

SP TP

A B

SoundofSpeedP

a S


27


In the free stream P = PS and V = VT,At the Probe face P = PT and V=0

01 0

/1/1

T

T

S V

P

PVdV

CPdP

Subsonic Speeds (continue)

2

1

1

2

/1

11T

ST

V

CPP

/1/1

1

SPC

12

12

2

11

2

11

2

a

VV

PP

P T

Pa

TSS

T

S

1

2

111 1

2

a

VP

P

PPPPQ T

SS

TSSTC

StaticOrifice

Pitot TubeT

S

VV

PP

T

S

VV

PP

0

V

PP T

SP TP

PressureTransducers


SP TP

A B


28


StaticOrifice

Pitot TubeT

S

VV

PP

T

S

VV

PP

0

V

PP T

SP TP

PressureTransducers


SP TP

A B

In the free stream P = PS and V = VT,At the Probe face P = PT and V=0

Supersonic Speeds

1

12

12

11

12

21

aV

aV

P

P

T

T

S

T

1

11

12

21

11

12

12

aV

aV

PP

PPPPQ

T

T

SS

TSSTC

Assume Supersonic Adiabatic Air Flow we obtain


29


StaticOrifice

Pitot TubeT

S

VV

PP

T

S

VV

PP

0

V

PP T

SP TP

PressureTransducers


SP TP

A BMach Number

1

1

2

12

11

12

21

M

M

P

P

S

T

Subsonic Speeds (M ≤ 1)

1

1

21

S

TT

P

P

a

VM

12

2

11

2

M

P

PSP

a

S

T

From

Supersonic Speeds (M ≥ 1)


30

Flight InstrumentsAirspeed Indicators (Calibrated Airspeed)

Calibrated Airspeed is obtained by substituting the Sea Level conditions, that is PS = PS0 , VT = VC , a0 = 340.294 m/s.

Subsonic Speeds (VC < a0=340.294 m/s)

1

2

11 1

2

00

a

VPQ C

SC

StaticOrifice

Pitot TubeT

S

VV

PP

T

S

VV

PP

0

V

PP T

SP TP

PressureTransducers


SP TP

A B

20

00

2

0

2

0 2

1

/21

21 C

S

CS

CS

aV

C VP

VP

a

VPQ

C

Supersonic Speeds (VC > a0=340.294 m/s)

1

11

12

21

1

12

0

12

0

0

a

V

a

V

PQ

C

C

SC

mmHginHgkPamkNmbarPS 00.7609213.29/325.10125.1013 20

γ air = 1.4


31


By measuring (TT) the Temperature of Free Airstream TS, we can compute the local Speed of Sound

StaticOrifice

Pitot TubeT

S

VV

PP

T

S

VV

PP

0

V

PP T

SP TP

PressureTransducers


SP TP

A B

Sa TRa

True Airspeed (TAS)

By using the Mach Number computation we can calculate the True Airspeed (TAS)

MM

TRMTRMaV T

aSaT

2

21

1


32

Flight InstrumentsGoodrich Air Data Handbook – Basic Air Data Calculation

Altitude

• Low Altitude: h<36.089 ft = 11000 m, PS > 6.6832426 in Hg

190255.0

1190255.0

190255.0190255.0

140000131252.092126.29140000131252.0

92126.29hP

Ph S

S

• Medium Altitude: 36.089 ft = 11000 m ≤ h ≤ 65.617 ft = 20000m6.6832426 in Hg> PS > 1.6167295 in Hg

hS

S

eP

P

h

30000480635.07345726.16832426.630000480635.0

6832426.6ln7345726.1

163156.34163156.34

1

96.710793

16.6451776167295.116.645177

6167295.196.710793

hP

Ph S

S

• High Altitude: h >65.617 ft = 20000m, PS < 1.6167295 in Hg


33


Impact Pressure STC PPQ where:

QC=½∙ ρ∙V2= Impact PressurePT = Total PressurePS = True Static Pressure

Indicated Airspeed (IAS)

1192126.29

1026.14797/2

CQ

IAS

Subsonic Flight (M ≤ 1)

1

1026.1479192126.29

2/72IAS

QC

Supersonic Flight (M ≥ 1)

1

14748.661

7

6

8411.60392126.29

2/5

2

7

IAS

IASQC

where:IAS = Indicated Airspeed in knotsQC = PT – PS Impact Pressure in Hg


34


Mach Number M = TAS/a

15115

7/27/2

S

T

S

C

P

P

P

QM

Subsonic Flight (M ≤ 1)

Supersonic Flight (M ≥ 1)

117

2.72.11

2/5

2

22

M

MM

P

P

P

Q

S

T

S

C

where:TAS = True Airspeed in knotsa = Speed of Sound in knots

where:QC=½∙ ρ∙V2= Impact Pressure in HgPT = Total Pressure in HgPS = True Static Pressure in Hg


35


Mach Number M = TAS/a

Altitude(feet)

75 KIAS(Qc=0.2701

In Hg)

100 KIAS(Qc=04814

In Hg)

200 KIAS(Qc=1.9589

In Hg)

300 KIAS(Qc=4.5343

In Hg)

400 KIAS(Qc=8.3850

In Hg)

500 KIAS(Qc=13.7756

In Hg)

600 KIAS(Qc=21.0749

In Hg)

700 KIAS(Qc=30.7642

In Hg)

S.L..113.151.302.454.605.756.9071.058

10,000.137.182.363.541.716.8881.0571.230

20,000.167.222.440.651.8541.0471.2421.453

30,000.207.276.541.7911.0231.2481.4891.754

40,000.262.347.672.9651.2361.5201.8292.171

50,000.331.438.8311.1711.5091.8752.2762.717

60,000.418.5491.0141.4261.8622.3352.8523.419

70,000.524.6841.2301.7542.3182.9283.5924.318

80,000.653.8421.4972.1722.8973.6784.5265.450


36


Static Temperature

22.01 M

TT T

S

True Airspeed (TAS)

where:TS = Static Temperature 1KTT = Total Temperature 1K

22.0196695.38

M

TMMTAS T

a

where:TAS = True AirspeedM = Macha = Speed of SoundTT = Total Temperature 1K


37


Vertical Speed Indicator

SOLO

38

Aircraft AvionicsFlight Instruments

SOLO

39

Aircraft AvionicsFlight InstrumentsAirspeed Indicator (ASI)

White Arc – Flaps Operation Range VSO – Stalling Speed Flaps Down VSI - Stalling Speed Flaps Up VFE – Maximum Speed Flaps Down (Extendeed)

Green Arc – Normal Operation Range VNO – Maximum Speed Normal Operation

Yellow Arc - Caution Range VNE – Not to Exceed Speed

Private Pilot Airplane – Flight Instruments ASA, Movie

http://www.youtube.com/watch?v=vICokuUAVAM

SOLO

40


Altimeters

SOLO

41

Aircraft AvionicsFlight InstrumentsAltimeters

SOLO

42


Gyroscopic Flight Instruments

Turn Indicator

SOLO

43


Attitude Indicator

SOLO

44


Attitude Indicator

SOLO

45


Turn Coordinator

SOLO

46


Turn-and Slip Indicator

SOLO

47


Attitude Heading Reference

SOLO

48


Heading IndicatorThe Magnetic Compass is sensitive to Inertia Forces. It is a reliable Heading Instrument in the long yerm, but during maneuvers it may swing and be hardly reliable. To provide a more precise Heading Instrument a Directional Gyro is used.

SOLO

49

Aircraft Avionics

Flight Instruments

SOLO

50

Aircraft AvionicsFlight InstrumentsFlux Gate Compass System

The Gate Compass System is connected to Radio Magnetic Indicator (RMI) and to Heading Situation Indicator (HSI).

Heading Situation Indicator (HSI).Radio Magnetic Indicator (RMI)

SOLO

51

Aircraft Avionics

Flight Instruments

SOLO

52


SOLO

53


SOLO

54

Aircraft Avionics

Flight Displays

SOLO

55

Aircraft Avionics

Flight Instruments

SOLO

56

Aircraft Avionics

Flight Displays

Chelton’s Flight Logic Reconfigurable Panel Display

SOLO

57

Aircraft Avionics

Flight Displays

Avidyne’s Entegra Reconfigurable Panel Display

SOLO

58

Aircraft Avionics

Flight Cockpit

SOLO

59

Aircraft Avionics

Flight Displays

SOLO

60

Aircraft Avionics

Flight Instruments

SOLO

61

Aircraft Avionics

Flight InstrumentsAutomatic Dependent Surveillance (ADS)

SOLO

62

Aircraft Avionics

Flight Instruments

SOLO

63

Aircraft AvionicsFlight InstrumentsAlert Systems

SOLO

64


Alert Systems

SOLO

65

Aircraft AvionicsFlight InstrumentsAlert Systems

SOLO

66


SOLO

67

Aircraft AvionicsFlight InstrumentsHelmet-up-Display

SOLO

68

Aircraft Avionics

NavigationInstrument Landing System (ILS)

SOLO

69

Aircraft Avionics

Flight Instruments

SOLO

70

Aircraft Avionics

SOLO

71

Aircraft Avionics

Cockpit

SOLO

72

Aircraft Avionics

Instrument Flight

SOLO

73


SOLO

74

Aircraft Avionics

To be replaced

SOLO

75

Aircraft AvionicsAerodynamics of Flight

76

SOLO

TechnionIsraeli Institute of Technology

1964 – 1968 BSc EE1968 – 1971 MSc EE

Israeli Air Force1970 – 1974

RAFAELIsraeli Armament Development Authority

1974 – 2013

Stanford University1983 – 1986 PhD AA

http://upload.wikimedia.org/wikipedia/en/a/ae/RAFAEL_logo.png

77

SOUND WAVESSOLO

SupersonicV > a

SubsonicV < a

a t a t

V tV t

M

1sin 1

Soundwaves

Machwaves

Disturbances propagate by molecular collision, at the sped of sound a,along a spherical surface centered at the disturbances source position.

The source of disturbances moves with the velocity V.

- when the source moves at subsonic velocity V < a, it will stay inside the family of spherical sound waves.

-when the source moves at supersonic velocity V > a, it will stay outside the family of spherical sound waves. These wave fronts form a disturbance

envelope given by two lines tangent to the family of spherical sound waves. Those lines are called Mach waves, and form an angle μ with the disturbance

source velocity:a

VM

M

&

1sin 1

78

SOUND WAVESSOLO

Sound Wave Definition: p

p

p p

p1

2 1

1

1

2 1

2 1

2 1

p p p

h h h

For weak shocks

up

1

2

11

11

1

11

11

2

12

1

1uuuuuu

)C.M.(

ppuuupuupu

11

111122111

211

)C.L.M.(

21

au 1

1p

1

1T

1e

112 uuu

112 ppp

112

112 TTT

112 eee

SOUND

WAVE

Since the changes within the sound wave are small, the flow gradients are small.Therefore the dissipative effects of friction and thermal conduction are negligibleand since no heat is added the sound wave is isotropic. Since

au 1

s

pa

2valid for all gases

79

SPEED OF SOUND AND MACH NUMBERSOLO

21

au 1

1p

1

1T

1e

112 uuu

112 ppp

112

112 TTT

112 eee

SOUNDWAVE

Speed of Sound is given by

0

ds

pa

RTp

C

C

T

dT

R

C

pT

dT

R

C

d

dp

dR

T

dTCds

p

dpR

T

dTCds

v

p

v

p

dsv

p

00

0

but for an ideal, calorically perfect gas

pRTa

TChPerfectyCaloricall

RTpIdeal

p

The Mach Number is defined asRT

u

a

uM

1

2

1

1

111

a

a

T

T

p

pThe Isentropic Chain:

a

ad

T

Tdd

p

pdsd

1

2

10

80

NORMAL SHOCK WAVESSOLO

Normal Shock Wave ( Adiabatic), Perfect Gas

G Q 0 0,

Mach Number Relations (1)

122

22

1

21

22

222

21

221

22

2

222

1

1

21

1222

2

11

1

22221

211

2211

2

1

2

12

1

2

1

*12

1

2

1

12

1

14..

...

..

uuu

a

u

a

uaa

uaaau

h

au

h

aEC

uuu

p

u

p

pupuMLC

uuMCp

a

Field Equations:

1222

2

11

2

2

1

2

1

2

1

2

1uuu

u

au

u

a

u u a1 22

u

a

u

aM M1 2

1 21 1

Prandtl’s Relation

u

p

T

e

u

p

T

e

11

q

1

1

1

1

1

2

2

2

2

2

1 2

2

1

2

11

2

1

2

1

2

1

21

2

12122

21

12

uu

auuuua

uu

uu

Ludwig Prandtl(1875-1953)

81



G Q 0 0,


M

MM

M

M

M

M

MM

22

22

1

12

12

12

12

12

21

1

2

1 1

2

11

1 21

2 1 2

1 1 1 1 1

12

or

M

M

M

M

MH H

A A

2

12

12

12

121 2

1 21

1

21

2

2

1

11

2

12

11

2

1

1

2

12

1 2

12

2 12 1

2

12

1 2 1

1 2

A A u

u

u

u u

u

aM

M

M

u

p

T

e

u

p

T

e

11

q

1

1

1

1

1

2

2

2

2

2

1 2

82



G Q 0 0,


p

p

up

u

u

u

a

MM

MM

M M

M

2

1

12

1

1

2

1

12

12

1

2

12 1

2

12 1

2 12

12

12

1 1 1 1

1 11 2

11

1 1 2

1

or

(C.L.M.)

p

pM2

1121

2

11

h

h

T

T

p

pM

M

M

a

a

h C T p R Tp2

1

2

1

2

1

1

212 1

2

12

2

1

12

11

1 2

1

s s

R

T

T

p

pM

M

M2 1 2

1

12

1

1

12

1

112

12

1

12

11

1 2

1

ln ln

s s

RM M

M2 1

1 1

2 12 3

2

2 12 41

2 2

3 11

2

11

Shapiro p.125

u

p

T

e

u

p

T

e

11

q

1

1

1

1

1

2

2

2

2

2

1 2

83

STEADY QUASI ONE-DIMENSIONAL FLOWSOLO

STAGNATION CONDITIONS

)C.E.( constuhuh 222

211 2

1

2

1

The stagnation condition 0 is attained by reaching u = 0

2

/

21202020

2

11

12

12

122

12

MTR

u

Tc

u

T

T

c

uTTuhh

TRa

auM

Rc

pp

Tch pp

Using the Isentropic Chain relation, we obtain:

2

10102000

2

11 M

p

p

a

a

h

h

T

T

Steady , Adiabatic + Inviscid = Reversible, , q Q 0 0, ~ ~ 0

G 0 t

0

SOLO

84

Civilian Aircraft AvionicsFlight Cockpit

CIRRUS PERSPECTIVE

Cirrus Perspective Avionics Demo, Youtube Cirrus SR22 Tampa Landing in Heavy Rain, Movie

http://www.youtube.com/watch?v=2-DZIysbrx0

http://www.youtube.com/watch?v=tq7eS1wXTeo

SOLO

85

Flight Displays

CIRRUS PERSPECTIVE

Civilian Aircraft Avionics

SOLO

86

Flight Displays

CIRRUS PERSPECTIVE


SOLO

87

Flight Displays

CIRRUS PERSPECTIVE


SOLO

88

Flight Displays

CIRRUS PERSPECTIVE


SOLO

89

Flight Displays

CIRRUS PERSPECTIVE


SOLO

90

Flight Displays

CIRRUS PERSPECTIVE


SOLO

91

Flight Displays

CIRRUS PERSPECTIVE


SOLO

92

Flight Displays

CIRRUS PERSPECTIVE


93

New integrated flight control system, Movie

https://www.youtube.com/watch?v=D3cH5ydHMzw


http://www.youtube.com/watch?v=D3cH5ydHMzw

2 aircraft flight instruments

Science

Transcript of 2 aircraft flight instruments