11.1. General Airspeed. ELEC4504 Avionics Systems 139

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1 ELEC4504 Avionics Systems 139 CHAPTER 11. Air Data Computer (ADC) In this section we shall be considering all of the functions of the air data computer separately in order to study their history, purposes and principles of operation. The main measurements to be considered are: a) (indicated, true and Mach number) b) c) Temperature d) Angle of Attack Note: although angle of attack is not strictly a function of the ADC, some sensors use airstream measurements to measure it General is probably the most important single piece of information the pilot needs. Virtually every phase of flight is conducted at a prescribed airspeed or range of airspeeds. e.g. V s is the stall speed, which is probably the most important airspeed since below V s the aircraft will not fly V 1 is the takeoff reject speed. If an emergency occurs below this speed, the takeoff is rejected and the aircraft is stopped on the runway. If the emergency occurs above this speed, the takeoff is continued. This speed is calculated for every takeoff and depends on such factors as the length of the runway, the temperature, the weight of he aircraft and the altitude of the airport. V R is the rotate speed. The speed at which the aircraft is expected to fly if the control yoke is pulled back. Climb and cruise speeds are defined for maximum efficiency. The approach speed is usually about 1.3 times the stall speed. This is a compromise between having a safe margin above the stall and having a minimum speed at threshold crossing to minimize the landing roll (and the amount of braking required) V NE is the never exceed speed which is determined by the aircraft structure. i.e. above this speed, parts start to fall off V MO is the maximum operating speed An addition to these, there are specified maximum airspeeds for such things as lowering the undercarriage and extending flaps primarily for structural reasons, All of the above speeds are Indicated airspeeds, which will be defined a little later True airspeed is derived from the indicated airspeed, and, as the name suggests, is the actual speed of the aircraft relative to the air mass in which it is moving. This is used almost exclusively for navigation purposes

2 140 ELEC4504 Avionics Systems Mach number is the speed of the aircraft relative to the local speed of sound. It is used mainly for cruise The second most important piece of information is altitude. Uses of altitude information: a) - by the air traffic control system to provide vertical separation between aircraft. b) - to avoid terrain (assuming one knows where the aircraft is in relation to the terrain) c) - to convert indicated airspeed to true airspeed d) - to control the pressurization system in the aircraft (to avoid rapid changes in pressure) Temperature Temperature information is used to compute Mach number and true airspeed and to indicate when external conditions are such that icing is likely Angle of Attack The angle of attack is used primarily to drive the stall warning and stall prevention systems. Older types of aircraft have a distinctive pattern of behaviour as they approach the stall but many modern aircraft do not. Also, in T tail aircraft it may not be possible to recover from a stall. Thus artificial means of providing warning and prevention are required At a certain angle below the stall angle, a stick shaker is activated. This is similar to the natural stall warning which occurs in older aircraft (due to turbulent airflow over the horizontal stabilizer and elevators from the stalling wing). If the pilot does not take appropriate action to decrease the angle of attack and if it increases, a stick pusher forces the stick forward to prevent the stall from occurring.

3 ELEC4504 Avionics Systems Principles of Operation Indicated (IAS) is measured using a pitot tube and a static vent. The pitot tube is a tube that is open at the front end and closed at the rear and is aimed directly into the relative wind Direction of flight Static Pressure Stagnation Pressure Figure 60: Schematic Diagram of Pitot Static System It produces a pressure which depends on the speed of the aircraft and which is called the stagnation pressure, the dynamic pressure or total pressure. The pitot tube is supplied with a heater to prevent icing. The static vent consists of a hole or an array of holes in the skin of the aircraft and is intended to measure the absolute pressure of the still air surrounding the aircraft. But because the air is moving past the aircraft the pressure at various places on the aircraft s skin is slightly higher or lower than the free stream pressure by an amount called the static defect. The static defect at a particular location depends on the speed, angle of attack, yaw angle, and altitude. The optimum location (i.e. the location which gives the smallest errors) of the static vents is determined by wind tunnel tests. In compressible subsonic flow the total pressure p t is γ ( γ 1) γ 1 ρv 2 p t = p s (1) γ 2 p s where γ = specific heat ratio of air (= 1.4) V = true airspeed ρ s = free-stream air density At low speeds, this reduces to: ρv 2 p t = p s

4 142 ELEC4504 Avionics Systems Thus p t p s ρv 2 2( p = and V t p s ) 2 IAS = ρ SL Indicated airspeed is proportional to the square root of p t p s note that p t p s is proportional to ρv 2 which is a common factor in the expressions for such effects as drag and lift which is why it is so widely used. When corrected for static defect and scale factor error, this is called the calibrated airspeed. Multiplying by ρ SL ρ gives effective airspeed which, below Mach 0.8 and 30,000 Ft. is within 7% of TAS. True (TAS) could theoretically be derived from equation (1) but this requires an accurate measurement of ρ which is difficult to do. Instead TAS is usually derived from the Mach number via the equation V TAS = T m Ma k( Ma) where Ma is the Mach number and k is the temperature probe recovery factor. Mach number at subsonic speeds is derived from ratio of the stagnation pressure and static pressure as shown in the equation γ Ma 2 = ---- p t p s γ γ Temperature Temperature is measured by a thermometer element on the exterior of the aircraft. On a moving aircraft this is higher than the free stream temperature due to frictional heating and compression of the air impinging on the thermometer. It is also altered by radiation from the thermometer to the sky and airframe. A probe similar to the pitot tube is used. It points along the aircraft axis and compresses the incoming air to zero relative speed, thus causing total, or stagnation temperature to exist at the thermometer. The temperature actually measured at the thermometer is γ 1 T m = T s η( Ma) 2 (2) 2 where absolute temperatures (usually in Kelvin) are used and where Ma is the local Mach number T s can be calculated from equation 2

5 ELEC4504 Avionics Systems 143 The temperature transducer is usually a small coil of wire whose resistance is a function of temperature. A leakage hole at the rear of the probe reduces the probe s time lag when the temperature changes. It also drains water from the probe. A shield surrounds the transducer in order to reduce heat exchange due to radiation from the aircraft skin or sky Electrical heaters can deice the probes without seriously affecting the accuracy of the measurement is measured using the static pressure only since external pressure decreases with increasing altitude. The relationship between pressure and altitude is defined by the ICAO standard atmosphere. This is a fictitious atmosphere defined as follows: Temperature at sea level: 15 C Lapse rate (rate of change of temperature with altitude): 1.98 C/1000Ft. to Ft. then C Sea level pressure: in Hg, kp or 1013 mbar All altimeters are calibrated to read correctly in such an atmosphere. Of course this atmosphere does not occur in practice but the only correction that is made is for the sea level pressure. Since the local pressure varies from hour to hour as weather systems move through the area is essential to make a correction so that at least near the ground at an airport altimeters will read correctly. This is done by providing the altimeter with a bias adjustment known as the altimeter setting. The altimeter setting is computed by measuring the actual air pressure at the level of the airport and then calculating, using the ICAO standard atmosphere, what pressure at sea level would produce the measured surface pressure. This is called the altimeter setting and causes all altimeters to read correctly on the airport s surface. While it is recognized that at any appreciable altitude above the airport, the actual altitude will be incorrect this can be ignored because the higher the aircraft the less danger of running into terrain. Since all aircraft flying at low altitudes use the altimeter setting from the closest airport, their altimeters all have the same errors and thus vertical separation can be maintained. Due to the high speed of jet aircraft, it has been found that maintaining the altimeter setting of the nearest airport required frequent changes. Thus for flight above Ft. in North America (called the transition altitude), all aircraft set their altimeters to a standard pressure of in Hg, or 1013 mbar. The transition altitude is different in other countries e.g. in the UK it is 8000 Ft. In this region altitudes are referred to as flight levels and are expressed as 100s of feet. e.g Ft. is flight level 360.

6 144 ELEC4504 Avionics Systems Note that in extremely cold weather, altimeters at high altitudes (20,000 Ft.) will read several thousand feet high which can be a hazard in mountainous regions. This is due to the fact that the cold air is denser than the air in the standard atmosphere and thus the pressure changes more quickly with increasing altitude leading to an erroneously high indicated altitude. The density is inversely proportional to the Absolute (Kelvin) temperature. Thus a temperature which is 27 degrees below standard gives a 27/273 or approximately 10% higher altitude reading. This is a 2000 Ft. error at 20,000 Ft. Question: Since temperature information is available, why doesn t the ADC correct for this? Angle of Attack There are two main methods of measuring angle of attack. The first is readily understood. A hinged vane with a wedge on the trailing edge is attached to the side of the aircraft. The hinge is connected to an angle measuring device such as a synchro and the angle information is sent to the air data computer or the cockpit instrument.

7 ELEC4504 Avionics Systems 145 The other method involves a hollow probe but with static vents on the top and bottom. Each vent is connected to a different static line and the air data computer determines angle of attack from the difference in pressure from the two ports α centre line of aircraft Airflow Figure 61: Schematic Diagram of Angle of Attack Sensor Figure 62: Angle of Attack Sensor on CL601 Challenger

8 146 ELEC4504 Avionics Systems A block diagram of a typical air data computer is shown in the diagram Static Defect Correction Static Pressure Transducer Computation Rate Memory - Hold Mach number computation Mach Number Maximum Allowable Maximum Allowable Speed Pitot Pressure Transducer computation q c = p t - p s True computation True Total Temperature Probe Static-air Temperature computation Static Air Temperature