Airspeed Calculators for Pilots

Indicated airspeed (IAS) is the reading shown on the airspeed indicator. It is derived from the dynamic pressure measured by the pitot tube — the difference between total pressure (from the pitot head) and static pressure (from the static port). IAS is a direct measure of dynamic pressure, which governs aerodynamic forces on the aircraft. This means that the aircraft stalls, rotates, and flies at the same IAS regardless of altitude, making IAS the correct reference for all structural speed limits and flight envelope boundaries. True airspeed (TAS) is the actual speed of the aircraft relative to the surrounding air mass. As altitude increases and air density decreases, a given dynamic pressure corresponds to a progressively higher actual speed — so TAS increases above IAS with altitude. At sea level on a standard day, IAS and TAS are approximately equal. At 10,000 feet, TAS is roughly 17% higher than IAS. At 35,000 feet, TAS may be 50–60% higher than IAS.

The primary method for calculating TAS from IAS uses pressure altitude and outside air temperature (OAT). The formula is: TAS = IAS × v(?0 ÷ ?), where ? is the actual air density and ?0 is the sea-level ISA density. Since air density depends on both pressure (altitude) and temperature, the calculation requires both values. A practical approximation is: TAS ˜ IAS × (1 + 0.02 × pressure altitude in thousands of feet), which gives roughly 2% increase per 1,000 feet. This approximation is accurate to within 1–2% up to approximately 15,000 feet on near-standard days. For more precise calculation, particularly at higher altitudes or large ISA deviations, the full density ratio formula or the CAS-to-TAS correction that accounts for compressibility effects should be used.

Calibrated airspeed (CAS) is indicated airspeed corrected for instrument error and position error (also called pressure error). Instrument error arises from mechanical imperfections in the airspeed indicator itself. Position error arises from the location of the pitot tube and static port on the airframe — the airflow at these locations is not perfectly undisturbed and creates a small pressure measurement error that varies with angle of attack and therefore with airspeed. The correction from IAS to CAS is published in the POH as a position error correction chart. At cruise speeds the correction is typically small (1–3 knots), but at low speeds near stall it can be significant (5–10 knots). For performance calculations requiring high accuracy, CAS should be used as the starting point rather than IAS.

Equivalent airspeed (EAS) is calibrated airspeed corrected for the compressibility of air. At low airspeeds and low altitudes, air behaves as an incompressible fluid for practical purposes and CAS equals EAS. Above approximately 200 knots CAS or at altitudes above 10,000 feet, air compressibility becomes significant and must be corrected for. EAS is used in structural engineering calculations because aerodynamic forces scale directly with EAS squared at any altitude. Aircraft structural speed limits (VNE, VA, VNO) are defined in terms of EAS (or effectively in IAS for the design envelope). For flight planning and navigation purposes, most general aviation pilots work with IAS and TAS directly, as CAS-to-EAS corrections are typically small below FL250.

Mach number is the ratio of the aircraft's true airspeed to the local speed of sound: M = TAS ÷ a, where a is the local speed of sound. The speed of sound in air depends on temperature: a = 38.95 × v(T) knots, where T is the absolute temperature in Kelvin. In the ISA at sea level (288.15K), the speed of sound is approximately 661 knots TAS. At FL350 (ISA temperature -56.5°C = 216.65K), it falls to approximately 573 knots TAS. Mach number matters because compressibility effects on aerodynamic forces become significant above approximately Mach 0.3, and critical aerodynamic phenomena — shock wave formation, wave drag, Mach tuck — occur at specific Mach numbers rather than at specific indicated airspeeds. Above FL280 in most transport category aircraft, Mach number replaces IAS as the primary speed reference because it correctly captures the compressibility-related performance limits.

VMO (Maximum Operating Speed) is the maximum speed in knots IAS at which the aircraft may be operated in smooth air. It is defined in IAS terms because aerodynamic forces — which determine structural loading — scale with dynamic pressure, and dynamic pressure is directly proportional to IAS squared. MMO (Maximum Operating Mach number) is the maximum Mach number at which the aircraft may be operated. It is defined separately from VMO because at high altitudes the compressibility effects that limit the aircraft occur at a specific Mach number regardless of IAS. The operating speed limit is always the more restrictive of VMO and MMO: at low altitudes VMO governs, at high altitudes MMO governs. The crossover altitude — where VMO and MMO intersect — is typically around FL280 to FL310 depending on aircraft type.

The speed of sound in a gas depends on the square root of the absolute temperature: a = v(?RT), where ? is the ratio of specific heats for air (approximately 1.4), R is the specific gas constant for air (287.05 J/kg·K), and T is the absolute temperature in Kelvin. As altitude increases in the troposphere, temperature decreases at the ISA lapse rate of approximately 1.98°C per 1,000 feet (6.5°C per kilometre). Lower temperature means lower molecular kinetic energy and slower propagation of pressure disturbances — so the speed of sound decreases. From the ISA sea-level value of approximately 661 knots, it falls to approximately 573 knots at the tropopause (approximately 36,090 feet in the ISA). Above the tropopause, temperature stabilises at -56.5°C in the ISA and the speed of sound remains approximately constant at 573 knots throughout the lower stratosphere.

Mach critical (Mcrit) is the Mach number at which the airflow over some part of the aircraft — typically the upper surface of the wing near its thickest point — first reaches the local speed of sound, even though the aircraft itself is flying below Mach 1. This occurs because the airflow accelerates as it curves around the wing. Above Mach critical, a shock wave begins to form on the wing surface, creating a sudden pressure discontinuity. This shock wave causes boundary layer separation, dramatically increases drag (wave drag), changes the pressure distribution on the wing, and can produce Mach tuck — an uncontrollable pitch-down tendency as the centre of pressure shifts aft. Swept-wing aircraft have higher Mach critical numbers than straight-wing aircraft because the sweep reduces the effective chord-wise component of airspeed.

Fuel burn for a jet engine is primarily a function of thrust, which is related to mass airflow through the engine. Because air density decreases with altitude, a jet engine at a given throttle setting moves less mass of air per second but at higher velocity — the TAS increases while IAS may remain constant. Specific fuel consumption (SFC, fuel burned per unit of thrust) for jet engines actually improves at altitude because the lower ambient temperatures improve thermodynamic efficiency. The net result is that jet aircraft are more fuel-efficient at high altitude despite higher TAS, because the drag in terms of IAS (and therefore dynamic pressure) can be set optimally. Piston aircraft are different: their fuel burn decreases with altitude primarily because the mixture can be leaned, but power output is limited by air density and TAS advantages are more modest.

Fuel burn rate (volume or mass per hour) should be taken from the POH power setting table at the planned power setting, altitude, and temperature — it is already expressed per unit of flight time, so the airspeed type does not matter for determining burn rate. However, to determine fuel burn for a specific leg of given distance, the correct speed to use is ground speed (GS), not TAS or IAS. Ground speed determines how quickly the aircraft covers the ground distance and therefore how much time (and fuel) a specific leg requires. Using TAS instead of GS ignores the headwind or tailwind component and will over or underestimate leg fuel if there is a significant wind component along track. Fuel burn per leg = leg distance ÷ GS × burn rate per hour.