Aircraft Flight Instruments
SOLO HERMELIN
Updated: 04.12.12
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Table of Content
SOLO Aircraft Avionics
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McDonnell Douglass F-4B Phantom Instrument Panel 3
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.
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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
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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.
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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
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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 /
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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
SOLO Aircraft Avionics
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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
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Flight InstrumentsAir Data Calculation (Collison)
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
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Flight InstrumentsAir Data Calculation (Collison)
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
SOLO Aircraft Avionics
18The Flight Panel - Understand Your Aircraft, Youtube
Flight Instruments
SOLO Aircraft Avionics
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Flight InstrumentsSOLO Aircraft Avionics
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Flight InstrumentsSOLO Aircraft Avionics
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zdgpd
TRp KsmR 22 /287
zdaTd
aR
g
T
za
p
p
00
1
Flight InstrumentsSOLO Aircraft Avionics
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Altimeter
SOLO Aircraft Avionics
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Flight InstrumentsAirspeed Indicators
SOLO Aircraft AvionicsFlight InstrumentsAirspeed Indicators
Centurion T210 AirSpeed Gauge
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Flight InstrumentsAirspeed Indicators
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
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Flight InstrumentsAirspeed Indicators
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
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Flight InstrumentsAirspeed Indicators
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
IASAirspeed Indicator
SP TP
A B
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Flight InstrumentsAirspeed Indicators
StaticOrifice
Pitot TubeT
S
VV
PP
T
S
VV
PP
0
V
PP T
SP TP
PressureTransducers
IASAirspeed Indicator
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
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Flight InstrumentsAirspeed Indicators
StaticOrifice
Pitot TubeT
S
VV
PP
T
S
VV
PP
0
V
PP T
SP TP
PressureTransducers
IASAirspeed Indicator
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)
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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
IASAirspeed Indicator
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
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Flight InstrumentsAirspeed Indicators
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
IASAirspeed Indicator
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
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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
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Flight InstrumentsGoodrich Air Data Handbook – Basic Air Data Calculation
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
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Flight InstrumentsGoodrich Air Data Handbook – Basic Air Data Calculation
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
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Flight InstrumentsGoodrich Air Data Handbook – Basic Air Data Calculation
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
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Flight InstrumentsGoodrich Air Data Handbook – Basic Air Data Calculation
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
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Flight InstrumentsAirspeed Indicators
Vertical Speed Indicator
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Aircraft AvionicsFlight Instruments
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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
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Aircraft AvionicsFlight Instruments
Altimeters
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Aircraft AvionicsFlight InstrumentsAltimeters
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Aircraft AvionicsFlight Instruments
Gyroscopic Flight Instruments
Turn Indicator
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Aircraft AvionicsFlight Instruments
Attitude Indicator
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Aircraft AvionicsFlight Instruments
Attitude Indicator
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Aircraft AvionicsFlight Instruments
Turn Coordinator
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Aircraft AvionicsFlight Instruments
Turn-and Slip Indicator
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Aircraft AvionicsFlight Instruments
Attitude Heading Reference
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Aircraft AvionicsFlight Instruments
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.
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Aircraft Avionics
Flight Instruments
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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)
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Aircraft Avionics
Flight Instruments
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Aircraft AvionicsFlight Instruments
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Aircraft AvionicsFlight Instruments
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Aircraft Avionics
Flight Displays
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Aircraft Avionics
Flight Instruments
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Aircraft Avionics
Flight Displays
Chelton’s Flight Logic Reconfigurable Panel Display
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Aircraft Avionics
Flight Displays
Avidyne’s Entegra Reconfigurable Panel Display
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Aircraft Avionics
Flight Cockpit
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Aircraft Avionics
Flight Displays
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Aircraft Avionics
Flight Instruments
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Aircraft Avionics
Flight InstrumentsAutomatic Dependent Surveillance (ADS)
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Aircraft Avionics
Flight Instruments
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Aircraft AvionicsFlight InstrumentsAlert Systems
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Aircraft AvionicsFlight Instruments
Alert Systems
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Aircraft AvionicsFlight InstrumentsAlert Systems
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Aircraft AvionicsFlight Instruments
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Aircraft AvionicsFlight InstrumentsHelmet-up-Display
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Aircraft Avionics
NavigationInstrument Landing System (ILS)
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Aircraft Avionics
Flight Instruments
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Aircraft Avionics
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Aircraft Avionics
Cockpit
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Aircraft Avionics
Instrument Flight
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Aircraft AvionicsFlight Instruments
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Aircraft Avionics
To be replaced
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Aircraft AvionicsAerodynamics of Flight
76
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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
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
NORMAL SHOCK WAVESSOLO
Normal Shock Wave ( Adiabatic), Perfect Gas
G Q 0 0,
Mach Number Relations (2)
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
NORMAL SHOCK WAVESSOLO
Normal Shock Wave ( Adiabatic), Perfect Gas
G Q 0 0,
Mach Number Relations (3)
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
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84
Civilian Aircraft AvionicsFlight Cockpit
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