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The CORRECT answer is (a) True

For explanation I would say: Altimeter indicates the height as a FUNCTION of barometric static PRESSURE in the atmosphere. Altimeter reading is the reading of INDIVIDUAL mechanical instrument which calculates the pressure. During these calculations some corrections are made to avoid errors that are caused due to mechanical tolerance.

The correct answer is (a) It is the indicated altitude MEASURED with respect to appropriate datum pressure SETTING

To elaborate: Pressure altitude it is the indicated altitude measured with respect to appropriate datum pressure setting. It is CORRECTED for static pressure error. This error is caused due to the static source being LOCATED within the disturbed pressure FIELD caused due to presence of aircraft.

The CORRECT CHOICE is (b) It is the pressure height interval measured by the ALTIMETER

Easiest EXPLANATION: Geopotential height interval it is the pressure height interval measured by the altimeter. Geopotential height interval is corrected for temperature difference from ISA model ATMOSPHERE.

The correct answer is (a) H=11000-\(\FRAC{RT_{11}}{g_0}\)ln(\(\frac{p}{p_{11}}\))

To ELABORATE: The height of the isothermal LOWER STRATOSPHERE is given by H=11000-\(\frac{RT_{11}}{g_0}\)ln(\(\frac{p}{p_{11}}\)) where,

H=geopotential height

R=characteristic gas constant

T11=temperature at 11km

P11=pressure at 11km

g0=acceleration due to gravity

p=pressure.

Correct choice is (B) 9144 km

For explanation I would SAY: The answer is 9144 km. Given L0=-0.0065 k/m, P=30070.36 Pa.

We know T0=288.15K, p0=101325Pa, R=287 J/kg-K, g0=9.81m/s^2.

From H=\(\frac{T_0}{l_0}\)[\((\frac{p}{p_0})^\frac{-L_0R}{g_0}\)-1]

H=\(\frac{288.15}{-0.0065}\)[\((\frac{30070.36}{101325})^\frac{-(-0.0065)*287}{9.81}\)-1]

H=9144 km.

The correct OPTION is (c) H=\(\frac{T_0}{l_0}\)[(\(\frac{P}{p_0}\))^\(\frac{-L_0R}{g_0}\)-1]

For EXPLANATION I would say: The geopotential HEIGHT in TROPOSPHERE is given by H=\(\frac{T_0}{l_0}\)[(\(\frac{p}{p_0}\))^\(\frac{-L_0R}{g_0}\)-1] where

H=geopotential height

R=characteristic gas constant

T0=temperature at 11km

P0=pressure at 11km

g0=acceleration due to gravity

p=pressure

L0=lapse rate.

Correct option is (b) \(\frac{p_1}{p_2}\)=(\(\frac{\rho_1}{\rho_2}\))^γ

Easiest explanation: The relation between pressure and DENSITY in an adiabatic air flow is GIVEN by \(\frac{p_1}{p_2}\)=(\(\frac{\rho_1}{\rho_2}\))^γ where p1, P2 are pressure values and ρ1, ρ2 are density values and γ is the ratio of specific heat at constant pressure to that of specific heat at constant VOLUME.

The correct choice is (a) The RELATIVE velocity between the aircraft and the air mass in which the aircraft is FLYING

The best I can EXPLAIN: Airspeed is the relative velocity between the aircraft and the air mass in which the aircraft is flying. The unit of airspeed is knot. There are different types of AIRSPEEDS. They are ground SPEED, true airspeed, calibrated airspeed and indicated airspeed.

The correct CHOICE is (b) False

Easy EXPLANATION: Airspeeds are of four types. They are ground speed, TRUE airspeed, CALIBRATED airspeed and indicated airspeed. Airspeed is the relative velocity between the aircraft and the air mass in which the aircraft is flying. The unit of airspeed is KNOT.

Correct choice is (a) CPT+\(\frac{V^2}{2}\)=constant

The explanation: CpT+\(\frac{V^2}{2}\)=constant is the general equation that RELATES flow TEMPERATURE to true airspeed, where Cp=specific HEAT at constant PRESSURE, T=temperature, V=true airspeed.

The CORRECT option is (B) indicated air speed

Easy explanation: IAS stands for indicated air speed. Indicated air speed is the speed indicated on the airspeed INDICATOR in the cockpit. Airspeeds are of four types. They are ground speed, TRUE airspeed, calibrated airspeed and indicated airspeed.

Correct option is (a) true AIR speed

Easy explanation: TAS stands for true air speed. True AIRSPEED is the relative speed of aircraft with RESPECT to the SURROUNDING air flow. As the altitude increases the pressure decreases and the true airspeed is greater than indicated airspeed.

The CORRECT option is (a) relative speed of aircraft with respect to the surrounding air flow

For explanation: TRUE airspeed is the relative speed of aircraft with respect to the surrounding air flow. As the altitude increases the pressure decreases and the true airspeed is GREATER than INDICATED airspeed. It is also REPRESENTED as TAS.

Right OPTION is (c) ground speed

To explain I would say: GS stands for ground speed. Ground speed is the relative movement between the AIRCRAFT and the ground. It is CORRECTED for true air speed. For example, 98 knots true air speed + 10 knots tailwind=108 knots ground speed.

The correct CHOICE is (d) CALIBRATED AIR speed

Explanation: CAS stands for calibrated air speed. Calibrated air speed is the indicated air speed corrected for the instrument and positional errors. Airspeeds are of four types. They are GROUND speed, TRUE airspeed, calibrated airspeed and indicated airspeed.

Correct OPTION is (a) R=CP\(\frac{\gamma-1}{\gamma}\)

For explanation I would say: R=Cp\(\frac{\gamma-1}{\gamma}\) is the correct relationship between R and Cp where R= characteristic GAS constant, γ is the ratio of specific heat at constant pressure to that of specific heat at constant volume and Cp is the specific heat at constant pressure.

Correct option is (B) CV=\(\frac{R}{\gamma-1}\)

The explanation: Cv=\(\frac{R}{\gamma-1}\) is the correct relationship between γ and Cv where R= characteristic gas constant, γ is the ratio of SPECIFIC HEAT at constant pressure to that of specific heat at constant VOLUME and Cv is the specific heat at constant volume.

The CORRECT choice is (a) 346.61K

Explanation: The answer is 346.61K. Given T=299K. We know that R=287J/kg-K, γ for air is 1.4.

From a=\(\SQRT{\gamma RT}\)

a=\(\sqrt{1.4*287*299}\)

a=346.61K.

Correct option is (a) \(\frac{p_1}{p_2}\)=\(\Big\{1+\frac{\gamma-1}{2}(\frac{V1}{a1})^2\Big\}^\frac{\gamma}{\gamma-1}\)

The explanation: The relation between pressure and AIR SPEED in ISENTROPIC relations is

\(\frac{p_1}{p_2}\)=\(\Big\{1+\frac{\gamma-1}{2}(\frac{V1}{a1})^2\Big\}^\frac{\gamma}{\gamma-1}\) where p1, p2 are PRESSURES at two points, V1=velocity at one point, a1=speed of sound at point one and γ is the ratio of specific heat at constant pressure to that of specific heat at constant volume.

Right option is (a) \(\frac{p_1}{p_2}=(\frac{T_2}{T_1})^\frac{\GAMMA}{1-\gamma}\)

Explanation: \(\frac{p_1}{p_2}=(\frac{T_2}{T_1})^\frac{\gamma}{1-\gamma}\)is the correct isentropic relation between pressure and TEMPERATURE where P1, p2 are PRESSURES, T1, T2 are temperatures and γ is the ratio of specific heat at constant pressure to that of specific heat at constant volume.

The CORRECT OPTION is (a) 340.26m/s

To explain: The answer is 340.26m/s. Given P=101306N/m^2, ρ=1.225kg/m^3 and we know that γ for air is 1.4. From the formula, a=\(\SQRT{\FRAC{\GAMMA P}{\rho}}\)

a=\(\sqrt{\frac{\gamma*101306}{1.225}}\)

a=340.26m/s.

Right ANSWER is (a) a=\(\SQRT{\FRAC{\gamma P}{\rho}}\)

To elaborate: The formula for speed of sound in terms of pressure and density is given by a=\(\sqrt{\frac{\gamma P}{\rho}}\) where a=speed of sound, P=pressure, ρ=density and γ is the ratio of specific HEAT at constant pressure to that of specific heat at constant VOLUME.

Correct CHOICE is (a) 959.1m/s

For explanation: The ANSWER is 959.1m/s. Given V=330m/s, σ=8.447. From the EQUATION Ve=V\(\SQRT{\SIGMA}\)

Ve=330\(\sqrt{8.447}\)

Ve=959.1m/s.

Right choice is (a) The calibrated air speed corrected for scale-altitude ERROR

The explanation is: Equivalent air speed is the calibrated air speed corrected for scale-altitude error. The correction is DONE in calibrated EQUATION of the airspeed indicator which is a FUNCTION of calibrated air speed and height.

The CORRECT choice is (a) M=\(\frac{V}{a}\)

The best EXPLANATION: The formula for MACH NUMBER is M=\(\frac{V}{a}\) where M=mach number, V=velocity, a=speed of sound. There is no unit for mach number as it is a ratio of speed of the object to the speed of sound in the surrounding air.

Correct option is (a) 1.318

Easy EXPLANATION: The answer is 1.318. Given T=290K , V= 450m/s. We know γ of air=1.4 and R=287 J/kg-K.

From the formula-M=\(\frac{V}{a}\) where a=\(\sqrt{\GAMMA RT}\)

On SUBSTITUTING the VALUES to find “a”,

we have a=\(\sqrt{1.2*287*290}\)

a=341.35m/s.

Now substituting the value of “a” in the formula=\(\frac{V}{a}\),

we get M=\(\frac{450}{341.35}\)

M=1.318.

Correct answer is (a) M=V\(\SQRT{\frac{\rho}{\GAMMA P}}\)

The EXPLANATION is: M=V\(\sqrt{\frac{\rho}{\gamma P}}\) is the correct relation for mach number in terms of pressure and density. In M=V\(\sqrt{\frac{\rho}{\gamma P}}\), M=mach number, V=velocity , ρ=density, P= pressure and γ is γ is the RATIO of specific heat at constant pressure to that of specific heat at constant VOLUME.

The correct choice is (a) 1.634

To ELABORATE: The answer is 1.634. Given P=101306N/m^2, V=556m/s. We KNOW that ρ and γ of air are ρ=1.225kg/m^3 and 1.4. From the equation M=V\(\SQRT{\frac{\RHO}{\gamma P}}\)

On substituting and solving,

M=556\(\sqrt{\frac{1.225}{1.4*101306}}\)

M=1.634.

Right choice is (a) \(\frac{p1}{p2}=\Big\{1+\frac{\gamma-1}{2}(M)^2\Big\}^\frac{\gamma}{\gamma-1}\)

Easy explanation: The relation between mach number and pressure is \(\frac{p1}{p2}=\Big\{1+\frac{\gamma-1}{2}(M)^2\Big\}^\frac{\gamma}{\gamma-1}\) where p1, p2 are pressures at two points, M=mach number and γ is the RATIO of specific HEAT at constant pressure to that of specific heat at constant volume.

Correct option is (B) 1

To explain: The answer is 1. Given σ=1.893 and we know that γ for air is 1.4.

On SUBSTITUTING the values in the EQUATION \(\frac{p_1}{p_2}=\BIG\{1+\frac{\gamma-1}{2}(M)^2\Big\}^\frac{\gamma}{\gamma-1}\)

We get 1.893=\(\Big\{1+\frac{1.4-1}{2}(M)^2\Big\}^\frac{1.4}{1.4-1}\)

On SOLVING we get M=1.

Correct answer is (a) True

The EXPLANATION is: Mach number of an aircraft is affected by the shock waves CREATED at the vortex of the aircraft. There are THREE types of shock waves they are NORMAL shock waves, oblique shock waves and expanded waves.

Right choice is (a) True

Explanation: Scale height is the description of how the ALTITUDE changes in the ATMOSPHERE. It is the vertical distance measurement where the DENSITY and pressure decrease by the factor of \(\frac{1}{E}\).

Correct option is (a) T0=T\(\Big[1+\frac{\gamma-1}{2}M^2\Big]\)

The best explanation: The relationship between temperature and MACH number is T0=T\(\Big[1+\frac{\gamma-1}{2}M^2\Big]\) where T is the temperature at that ALTITUDE, T_0 is the STAGNATION temperature, M is mach number and γ is the ratio of specific heat at CONSTANT pressure to that of specific heat at constant VOLUME.

The CORRECT choice is (b) 240.125K

To explain: The ANSWER is 240.125K. Given T0=288.15K and M=1. By substituting the values in the FORMULA T0=T\(\Big[1+\FRAC{\gamma-1}{2}M^2\Big]\)

On substituting 288.15=T[1+\(\frac{1.4-1}{2}1^2]\)

T=\(\frac{288.15}{1.2}\)

T=240.125K.

Right answer is (a) True

For explanation: The ratio of indicated temperature rise to IDEAL PRESSURE rise is known as recovery FACTOR. It is given by the expression r=\(\frac{T_i-T}{T_t-T}\) where r is known as recovery factor, TI is known as indicated total temperature, Tt is known as total temperature and T is known as static temperature.

The correct choice is (b) R=\(\frac{T_i-T}{T_t-T}\)

To explain: The expression for RECOVERY factor is given by r=\(\frac{T_i-T}{T_t-T}\) where r is KNOWN as recovery factor, TI is known as indicated total temperature, Tt is known as total temperature and T is known as static temperature. The ratio of indicated temperature rise to ideal pressure rise is known as recovery factor.

Right CHOICE is (a) 1

To explain I would say: The answer is 1. The EXPRESSION for recovery factor is given by r=\(\frac{T_i-T}{T_t-T}\)where r is known as recovery factor, TI is known as INDICATED total temperature, Tt is known as total temperature and T is known as static temperature. In the formula if Tt=Ti then

r=\(\frac{T_i-T}{T_i-T}\)

r=1.

Correct choice is (a) T0=T\(\BIG[1+\frac{\gamma-1}{2}(\frac{V}{a})^2\Big]\)

To ELABORATE: The RELATION between speed of object and temperature is T0=T\(\Big[1+\frac{\gamma-1}{2}(\frac{V}{a})^2\Big]\)where T is the temperature at that altitude, T0 is the stagnation temperature, γ is the ratio of specific HEAT at constant pressure to that of specific heat at constant volume, V is speed of object and ‘a’ is speed of sound.

The correct choice is (a) True

Easy EXPLANATION: The difference in indicated values and LOCAL values of altitude, airspeed and MACH number is known as system pressure error. The indicated values of the altitude, airspeed and mach number resulting from the measured values of the local or system, pressures will differ from the values that WOULD occur when USING the undisturbed freest stream pressures. This error is known as system pressure error.

The correct choice is (a) 780K

The best EXPLANATION: The answer is 780K. Given r=2, T=300K, M=2 and we know that γ for air is 1.4. From the formula Ti=T[1+r\(\FRAC{\gamma-1}{2}\)M^2].

On SUBSTITUTING the values we get Ti=300[1+2\(\frac{1.4-1}{2}\)2^2]

On solving above equation we get Ti=780K.

Correct choice is (a) Ti=T[1+r\(\frac{\gamma-1}{2}\)M^2]

Explanation: The relation between RECOVERY FACTOR and TEMPERATURE is Ti=T[1+r\(\frac{\gamma-1}{2}\)M^2] where T is the temperature at that altitude, Ti is indicated TOTAL temperature, r is recovery factor, M is Mach number and γ is the ratio of specific heat at CONSTANT pressure to that of specific heat at constant volume.

Right ANSWER is (b) 2

For EXPLANATION I would SAY: The answer is 2. Given r=2, T=300K, Ti=780K and we know γ of air is 1.4. From the formula Ti=T[1+r\(\FRAC{\gamma-1}{2}\)M^2].

On substituting the values in the formula, we get 780=300[1+2\(\frac{1.4-1}{2}\)M^2].

On solving we get M=2.