• The pitot-static system of an aircraft is a system in which total pressure created by the forward motion of the aircraft and the static pressure of the atmosphere surrounding it are sensed and measured in terms of speed, altitude and rate of change of altitude (vertical speed).

Slide 1

The Earth's Atmosphere and Characteristics

The earth's atmosphere, is the surrounding envelope of air, which is a mixture of a number of gases the chief of which are nitrogen and oxygen. This gaseous envelope is conventionally divided into several concentric layers extending from the earth's surface, each with its own distinctive features.

TroposphereThe lowest layer, the one in which we live and in which conventional types of aircraft are mostly flown. It to a height of about 28,000 ft at the equator. This height boundary is termed the tropopause.

StratosphereIt is above the tropopause and extends to the stratopause at an average height between 60 and 70 miles.

At greater heights the remaining atmosphere is divided into further layers or regions which from the stratopause upwards are termed the ozonosphere, ionosphere and exosphere.

Slide 2

Atmospheric Pressure

The atmosphere is held in contact with the earth's surface by the force of gravity, which produces a pressure within the atmosphere. Gravitational effects decrease with increasing distances from the earth's center, so that atmospheric pressure decreases steadily with altitude. The units used in expressing atmospheric pressure are: pounds per square inch, inches of mercury and millibars. The standard sea-level pressure is 14.7 Ibf/in. and is equal to 29.92 1 in Hg or 101 3.25 mbar.

Standard Pressure Lapse Rate1 inHg per 1000ft. of altitude.

1 inHg =3386 Pa=0.0339 bar=0.0334 atmosphere=25.4 mmHg=0.49 psi

Slide 3

Atmospheric Temperature The air in contact with the earth is heated by conduction and radiation, and as a result its density decreases and the air starts rising. In rising, the pressure drop allows the air to expand, and the expansion in turn causes a fall in temperature.

The assumptions are as follows: (i) the atmospheric pressure at mean sea-level is equal to 1013.25 mbar or 29.92 1 in Hg; (ii) the temperature at mean sea-level is 15°C (59° F); (iii) the air temperature decreases by 1.98°C for every 1,000 ft increase (Lapse rate) in altitude from 15°C at mean sea-level to -56.5°C (69.7° F) at 36,089 ft. Above this altitude the temperature is assumed to remain constant at -56.5°C.

Slide 4

Measurement of Atmospheric Pressure

There are two principal methods of measuring the pressure of the atmosphere.

They are:

(i) balancing of pressure against the weight of a column of liquid

(ii) magnifying the deflection of an elastic sensing element produced by the pressure acting on it.

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Slide 6

Basic pitot static system

Slide 7

Aneroid Barometer and Altimeter

The aneroid comes from the Greek aneros, 'not wet‘.

•The pressure-sensing element of the instrument is an evacuated metaI capsule.Since there is approximately zero pressure inside the capsule, and assuming the instrument to be at sea-level, approximately 14.7 lbf/in2 on the outside, the capsule will tend to collapse.

• Balancing force is given by leaf spring which always tends to open outwards.

• When atmospheric pressure decreases, the force tending to collapse the capsule is decreased but the spring tension remains the same and consequently is able to open out the capsule a Iittle further than before.

• The resulting expansion and contraction of the capsule, which is extremely small, is transformed into rotary motion of the pointer by means of a magnifying lever system and a very finely-linked chain.

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Slide 9

Altimeter Dial Presentations

Slide 10

Effect of atmospheric temperature on an altimeter

Slide 11

Relation between various altitudes

Slide 12

QFE Setting the pressure prevailing at an airfield to make the altimeter read zero on landing and take-off.

QNE Setting the standard sea-level pressure of 1,013.25 mbar (29.92 in Hg) to make the altimeter read the airfield elevation.

QNH Setting the pressure scale to make the altimeter read airfield height above sea-level on landing and take-off,

Slide 13

Pitot Pressure

This may be defined as the additional pressure produced on a surface when a flowing fluid is brought to rest, or stagnation, at the surface.

work must be done by the mass of liquid and this raises an equal volume of the fluid above the level of the fluid stream.

coming to rest at the stagnation point, kinetic energy of the fluid is converted into pressure energy.

Slide 14

•If the mass of the fluid above this level is m pounds then the work done in raising it through a height h feet is given by

Work done = m × g × h foot-poundswhere g is the acceleration due to gravity. In the British Isles at sea level, g = 32.2 ft/s2 = 9.81

m/s2

•The work done is also equal to the product of the ratio of mass to density (ρ) and pressure (p):Work done = (m/ ρ) × p

•The kinetic energy of a mass m before being brought to rest is equal to 1/2m V2 , where V is the

speed, and since this is converted into pressure energy,(m/ ρ) × p = to 1/2m V2

•Therefore p=1/2ρV2 and is additional to the static pressure in the region of fluid flow.

Slide 15

Pitot-Static Probes

Figure : Basic form of pitot-static probe. 1 heating element, 2 static slots, 3 pitot tube connection, 4 static tube connection, 5 heater element cable, 6 external drain hole, 7 pitot tube drain hole.

•The tubes are mounted concentrically, the pitot tube being inside the static tube, which also forms the casing.

•Static pressure is admitted through either slots or small holes around the casing.

•The heating element is fitted around the pitot tube, or in some designs around the inner circumference of the outer casing,

Slide 16

Heating Circuit Arrangements

The direct current for heating is controlled by a switch located on a cockpit control panel

Figure : Typical heating circuit arrangements. (a) Light and relay; (b) ammeter; (c) magnetic indicator and relay.

Slide 17

Pressure (Position) Error:

The accurate measurement of airspeed and altitude, by means of a pitot-static probe, has always presented two main difficulties

i. to design a probe which will not cause any disturbance to the airflow over it

ii. to find a suitable location on the aircraft where the probe will not be affected by air disturbances.

pressure or position error (PE) which is defined as the amount by which the local static pressure at a given point in the flow field differs from the free-stream static pressure.

Due to the aircraft itself these effects of such disturbances are greatest on the static pressure measuring section of a pitot-static system giving rise to a pressure or position error (PE)

Slide 18

As a result of PE, an altimeter and an airspeed indicator can develop positive or negative errors. The vertical speed indicator remains unaffected.

PE corrections are performed automatically and in a variety of ways.

•One method is to employ aerodynamically compensated pilot-static probes, i.e. probes which are so contoured as to create a local pressure field which is equal and opposite to that of the aircraft, so that the resultant PE is close to zero.

•Other methods more commonly adopted utilize correction devices within separate transducers.

Slide 19

Pressure Error Correction Transducers

•The transducer consists of a static pressure-sensing capsule and a dynamic pressure-sensing capsule, each capsule being connected to a d.c. potentiometer.

Slide 20

•The PE corrections are applied to the potentiometers by a network of pre-adjusted resistors connected across various tapping points on the potentiometers. Thus, their out puts are `shaped' electrically to provide the signals necessary to correct the known errors of a particular type of aircraft at various speeds and altitudes.

•The generator also amplifies the shaped signals which, after modulation, are finally supplied to the associated servo altimeter.

•The power supply circuit of the transducer is energized from a 26 V 400 Hz source, and the circuit also provides a d.c. output for operating a correction failure warning device.

Slide 21

Alternate Pressure Sources

If failure of the primary pitot-static pressure source occur, for example, complete icing up of a probe due to a failed heater circuit, then it is obvious that errors will be introduced in the indications of the instruments .

As a safeguard against failure, therefore, a standby system may be installed in aircraft employing pitot-static probes whereby static atmospheric pressure and/or pitot pressure from alternate sources can be selected and connected into the primary system.

The required pressure is selected by means of selector valves connected between the appropriate pressure sources and the flight instruments, and located in the cockpit within easy reach of the flight crew.

Slide 22

Figure: Alternate pitot pressure and static pressure system

Slide 23

Drains:

In order for a pitot-static system to operate effectively under all flight conditions, provision must also be made for the elimination of water that may enter the system as a result of condensation, rain, snow, etc., thus reducing the probability of `slugs' of water blocking the lines. Such provision takes the form of drain holes in probes, drain traps and drain valves in the system pipelines. Drain holes are drilled in probe pitot tubes and casings, and are of such a diameter that they do not introduce errors in instrument indications.

Pipelines

Pitot and static pressures are transmitted through seamless and corrosion-resistant metal (light alloy and/or tungum) pipelines and flexible pipes, the latter being used for the connection of components mounted on anti-vibration mountings.A minimum acceptable limit to the internal diameter, namely 1/4 in. A smaller internal diameter would present the hazard of a blockage due to the probability of a `slug' of water developing in such a way as to span the diameter.

Slide 24

Vertical speed indicators

It is very sensitive differential pressure gauges, designed to indicate the rate of altitude change from the change of static pressure alone.

This indicator consists basically of three main components, I.a capsule,II.an indicating element and III.a metering unit

The metering unit forms part of the static pressure connection and is connected to the interior of the capsule to prevent pressure surges reaching the capsule. The other end of the metering unit is open to the interior of the case to apply static pressure to the exterior of the capsule.

Slide 25

(a) Level flight: zero differential pressure across capsule;

(b) aircraft descending: metering unit maintains case pressure lower than capsule pressure, changing it at the same rate and thereby creating a constant differ ential pressure across the capsule;

(c) aircraft climbing: metering unit creates a constant differ ential pressure across cap sule by maintaining case pressure higher than cap sule pressure.

Slide 26

Figure: Vertical speed indicator metering units.(a) Capillary and orifice type;(b) capillary and orifice characteristics

Metering Unit

Slide 27

Airspeed indicators

Airspeed indicators are in effect very sensitive pressure gauges measuring the difference between the pitot and static pressures detected by the pressure probe, in terms of the 1/2ρV2 formula.

p = 1/2 ρV2 (1+V2 /4a02)

where p = Pressure difference (mmH20)

ρ = Density of air at sea-levelV = Speed of aircraft (m.p.h. or knots) ao = Speed of sound at sea-level (m.p.h.)

Present-day airspeed indicators are calibrated to the law

Slide 28

Square-law Compensation

Airspeed indicators measure a differential pressure which varies with the square of the airspeed. If the capsule were coupled to the pointer mechanism so that its deflections were directly magnified, the instrument scale would be like this.

The non-linearity of such a scale makes it difficult to read accurately, particularly at the low end of the speed range; further more, the scale length for a wide speed range would be too great to accommodate conveniently in the standard dial sizes.

Slide 29

The method for making the dial linear is to change the length of a lever as progressive deflections of the capsule take place, causing the mechanism, and pointer movement to be increased for small deflections and decreased for large deflections. In other words, it is a principle of variable magnification.

Figure: Methods of square law compensation. Rocking-lever/sector-arm mechanism.

Slide 30

Airspeed terminology

Slide 31

Combined Airspeed Indicator and Warning Switch

The system comes into operation when the approach speed of the aircraft is reached and the landing gear is not extended and locked in position. If this should happen, a warning flag visible through an aperture in the dial, adjacent to the approach speed graduations, commences to oscillate.

Figure: Combined air speed indicator and warning system. 1 Flag operating coil, 2 contact assembly, 3 capsule, 4 rocking shaft, 5 bimetal arm, 6 sector, 7 hand staff, 8 pointer, 9 rocking shaft flag assembly.

Slide 32

• When the landing gear is retracted, the direct-current supply is applied to the indicator, but while airspeeds are above the preset value, the capsule holds the contacts open.

• As the airspeed decreases, the capsule contracts until at the preset value the contacts close and

complete two parallel circuits.1. One circuit energizes the relay. Energizing of the relay causes its contacts to

change over, thus interrupting the supply to the relay coil and also connecting a supply to the flag operating coil, causing the flag to move into the dial aperture.

2. At the same time the capacitor is supplied with direct current by other circuit and it starts charging. When it discharges it does so through the relay coil and holds the contacts in position until the discharge voltage reaches a point at which it is sufficient to hold the relay energized. The contacts change over once again and de-energize the flag actuator coil, causing the flag to move away from the dial aperture.

The cycle is then repeated and at such a frequency that the flag appears in the aperture at approxi mately half-second intervals.

Slide 33

Machmeters

With the advent of the gas turbine, the propulsive power available made it possible for greater flying speeds to be attained, but at the same time certain limiting factors related to the strength of an air frame structure, and the forces acting on it, were soon apparent.

Mach number (M) = V/aWhere,V= aircraft's speed, a= the speed of sound,

The passage of an aircraft through the air sets up vibratory disturbances of the air which radiate from the aircraft in the form of pressure or sound waves. At speeds below that of sound, termed subsonic speeds, these waves radiate away from the aircraft in much the same way as ripples move outward from a point at which a stone is thrown into water. When speeds approach that of sound, however, there is a drastic change in the sound-wave radiation pattern. The air craft is now travelling almost as fast as its own sound waves and they begin to pile up on one another ahead of the aircraft, thus inceasing the air resistance and setting up vibrations of the air causing turbulence and buffeting of the aircraft, thus imposing severe stresses.

Slide 34

critical Mach number (Mcrit) is the mach number when the aircraft speed is below sonic speed but for the location, aerodynamic shape and profile of parts of the structure may allow the airflow over them to reach or exceed sonic speed. The critical Mach number for a particular type of aircraft is indicated by a pre-adjusted lubber mark located over the dial of the Machmeter.

Figure: Machmeter. 1 Airspeed capsule, 2 'altitude capsule, 3 altitude rocking shaft, 4 sliding rocking shaft, 5 calibration spring (square law compensation), 6 cali bration screws (square-law compensation).

Slide 35

The first variable for mach meter is airspeed and therefore a mechanism based on the conventional airspeed indicator is adopted to measure this in terms of the pressure difference (p – s), where p is the total or pitot pressure and s the static pressure.

The second variable is altitude because the speed of sound cannot be measured by the instrument, but since it is governed by static pressure conditions, the altimeter mechanism can do the best thing.

Mach number V/a in terms of the pressure ratio = (p - s)/s

Slide 36

Mach Warning System

In many types of high-performance aircraft, a switch unit utilizing the principle of the Machmeter is provided; its purpose being to give an aural warning on the flight deck whenever the maximum operating limit speed is exceeded. The mechanical and electrical arrangement of a switch unit, based on that employed in a typical aircraft, is shown in figure.

Figure: Mach warning system.

Slide 37

The switch contacts are actuated by the airspeed and altitude capsule assemblies, and remain closed as long as the speed is below the limiting value. The 28 V d.c. supply passing through the contacts energizes the control relay which interrupts the ground connection of the aural warning device, or `clacker' as it is called from the sound which it emits. When the limiting speed is exceeded, the switch contacts open thereby de-energizing the control relay to allow d.c. to pass through its contacts to activate the `clacker' via the now completed ground connection. The sound is emitted at a frequency of 7 Hz.

Slide 38

AIR DATA COMPUTERS

The pressures on which the operation of the primary flight instruments is dependent are transmitted through a system of pipelines. It should be apparent therefore, that the length and quantity of the pipelines will vary according to the size of the aircraft, and also on the number of stations at which indications of the relevant air data are required.

In order to minimize the `plumbing arrangements, the concept of supplying the pressures to a special unit at some centralized location, and then transmitting the air data electrically to wherever required, was developed and resulted in the design of units designated as central air data computers (CADC).

CADC is a device that produces electrical signal equivalents of pitot and static pressures by the combined operation of mechanical and synchronous transmission devices. The final computed output signals are then supplied to the appropriate indicators which, unlike their conventional counterparts, contain no pressure-sensing elements.

Slide 39

Figure: Modular arrangement of a CADC system.

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Questions?

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Thank You