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I remember learning and memorizing that TAS increases as altitude increases, however I don't understand why. Especially since my C-152's PoH seems to indicate the opposite in the cruise performance section. Can someone please explain what I'm misunderstanding here? Image of C-152 Cruise Performance

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  • $\begingroup$ Welcome to Av.SE! $\endgroup$
    – Ralph J
    Dec 18 '20 at 21:39
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    $\begingroup$ "I remember learning and memorizing that TAS increases as altitude increases, however I don't understand why."-- do you mean when IAS is held constant, or what? When the throttle position is constant? When the horsepower output is constant? In questions like this, it is always important to say what the constraint is, i.e. what variable(s) are being held constant. $\endgroup$ Dec 19 '20 at 2:47
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    $\begingroup$ . . . and AlexK I think you raised a very good question regarding the POH values of TAS reducing. If something looks wrong, don't wish it away, try and figure it out. $\endgroup$
    – skipper44
    Dec 19 '20 at 11:41
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To the extent that 75% cruise power available is constant with increasing altitude, it's simply that TAS goes up because the constant HP has less and less resistance to work against as the air thins out, while you are able to maintain 75% cruise power by opening throttle more. On a non-turbocharged engine, that is possible up to the point at which the engine is maxed out making cruise power (75%), somewhere around 8000 feet give or take (where the available manifold pressure at wide open throttle drops to about 22").

At this point the engine is wide open throttle just to make 75% power and is maxed out there. As you go above 8000 ft with wide open throttle, by the time you get to 10000 ft it can produce only 68 thrust HP (per the chart, in standard conditions) with wide open throttle and maximum TAS drops to 103 kt.

If the engine had turbocharging, it would be able to maintain 75 HP (75% of 100) to some much higher altitude and you could climb to say, 15000 ft, still making 75 HP in the even thinner air thanks to the turbo, and your TAS would be up to something like, say, 115 kt (you would also need a constant speed prop to exploit that power, but that's another issue).

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Imagine that IAS measures the number of air particles hitting your aircraft per second. As you get higher the air is thinner — there are progressively fewer air particles around the higher you go so to get the same number of particles hitting you per second (ergo the same IAS) you need to move faster through the air - your TAS increases.

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  • $\begingroup$ You have explained the difference between IAS and TAS, but you haven’t explained why TAS is higher at high altitude. $\endgroup$ Dec 19 '20 at 0:56
  • $\begingroup$ Roger. I added that the air gets thinner as you go up. $\endgroup$
    – Arkhem
    Dec 19 '20 at 2:32
  • $\begingroup$ With that explanation speed should change with the inverse of density when in fact it changes with the square root of that inverse. I think there is still room for improvement. Hint: The Newtonian impact theory only works for very high Mach numbers. $\endgroup$ Dec 19 '20 at 6:26
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    $\begingroup$ @PeterKämpf to clarify, at no point did I address the specifics of calculating dynamic pressure with altitude. Pedagogically, I am explaining only in simple terms to try and target the estimated level of understanding required to aid the questioner’s mental model. I have supposed them to be a new or student pilot without the requisite training to jump directly to the maths, if it turns out to be incorrect I apologise for the disservice! 😊 $\endgroup$
    – Arkhem
    Dec 19 '20 at 9:29
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True air speed (TAS) does indeed increase with altitude. The reason is because with increasing altitude the air density decreases. The Indicated air speed (IAS) which is measured by the pitot is a function of the dynamic pressure (Q) acting on the aircraft. The equation for the dynamic pressure is:

Q = 1/2 * rho * V^2

where:

Q = dynamic pressure.

rho = density.

V = True air speed.

According to the equation, if we want to keep the IAS or Q constant in a climb, the TAS should increase, because there is a decrease in density (rho) with altitude. So, there is absolutely nothing wrong with what you have learnt, because it is correct in the sense TAS increases with altitude.

So, why is your POH contradictory. Let us look at it, shall we? I will use the data given at 2000 ft and 12000 ft, under standard temperature to explain what is happening.

enter image description here

If we look at the table we can see that at 2000 ft, if you cruise at 2300 RPM, the engine is able to produce 66 Brake horsepower. The result is a TAS of 96 knots. But at 12000 ft the same engine is only able to produce 54 Brake horsepower at the same 2300 RPM. The reason why the TAS dropped to 92 knots is because your engine is unable to generate enough power at higher altitudes due to the reduced air density (C-152 has a normally aspirated engine). Hence, the aircraft is unable to fly at a higher TAS. If you want to increase your TAS you could increase the RPM to 2450 (more air rammed inside the cylinders) and get a TAS of 100 knots, but that will cost you more fuel.

It is expected that you will climb at the speed for best rate of climb (Vy). Once you reach your cruise altitude, you will push the nose down to level off, pull the power back to the desired RPM and trim it. If you climbed to 12000 ft and set 2300 RPM and trim the aircraft, the IAS that is shown in your air speed indicator will result in the tabled value of 92 knots. The value of IAS is not mentioned in cruise performance is because TAS is what determines your aircraft navigation performance and fuel consumption.

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  • $\begingroup$ Flying at 12,000’ at 2450 rpm will NOT cost you more fuel because 100 KTAS @ 5 GPH is better than flying at 2000’ at 2300 rpm with only 96 KTAS @ 5.4 GPH. $\endgroup$ Dec 20 '20 at 7:25
  • $\begingroup$ I was comparing the power at 12000 ft. You burn more at 2450 than at 2300. $\endgroup$
    – Anas Maaz
    Dec 20 '20 at 9:47
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Although, it is a backwards way of thinking about it, consider that the True AirSpeed is not changing with altitude for an aircraft keeping a constant speed. In actuality, the Indicated AirSpeed is what is changing. This is due to the IAS being measured by using Ram Air Pressure inside of the pitot tube. The RAP is directly related to the density of the air. The higher the air density, the higher the RAP at the same speed.

More gas molecule mass in the air is entering the pitot at any given time at lower density altitudes (higher air density). Think of it this way. A single bb would have less impact than 100 shot gun pellets fired at the same muzzle velocity.

Your pitot static system is actually measuring the impact of the air molecules and converting that into IAS. In order to fly at a constant IAS at a higher altitude, you have to increase your TAS because the air is less dense (fewer air molecules).

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  • $\begingroup$ You have explained the difference between IAS and TAS, but you haven’t explained why TAS is higher at high altitude $\endgroup$ Dec 19 '20 at 0:52
  • $\begingroup$ @MikeSowsun - Because TAS is not higher at higher altitudes. IAS is lower at higher altitudes for the same aircraft speed. But, IAS is measurable with the ASI. Whereas TAS is calculated. Also, IAS is more useful in determining aircraft performance (V speeds, etc). Unfortunately, that is the best way I can explain it. If I can up with something better, I will revisit this. $\endgroup$
    – Dean F.
    Dec 19 '20 at 2:23
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For a given power setting, True Airspeed increases with altitude because there is less drag due to the air being less dense. Aircraft are more efficient at high altitude because of this simple fact. Providing the engine can produce enough power, any aircraft will fly faster and further, without burning more fuel, if you fly at a higher altitude.

Your C-152 chart shows this quite clearly.

2000’ 59% power = 91 KTAS @ 4.8 GPH

12,000’ 59% power = 97 KTAS @ 4.8 GPH

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Inflight, the IAS is measured via the pitot/static system and is displayed on your ASI, the primary airspeed display instrument. TAS cannot be directly measured by a probe or transducer, it has to be calculated by using the IAS as a starting reference.

IAS: The pitot pressure is static pressure + the pressure on account of air molecules impinging on the pitot probe from the direction of flight. If the air density reduces, as it does with altitude, in order to fly a given IAS, it requires the airplane to fly faster through the thinner air for the pitot to experience the same pressure and indicate the same IAS. This faster speed is your TAS. Aerodynamic forces such as lift, drag, propeller torque/thrust etc. all depend on the flow of air molecules over the wings and body of the airplane and stall speeds, maneuvering speeds and limit speeds all relate therefore to the IAS.

TAS: is a calculated speed, as there is no instrumentation to physically measure it - in it's simplest form it is arrived at by plugging the IAS into a formula and based on the International Standard Atmosphere. TAS is useful for planning and actual navigation as it is your true speed in relation to the ground in nil winds.

Atmospheric Air Density as per ISA: As per the Standard atmosphere around 6600ft the density of air is about half of what it is at Mean Sea Level, at 12000ft around a quarter, and around 18000ft it is a tenth of the sea level density and so on.

You memorized "TAS increases as altitude increases" but the statement is incomplete in that it gives no info on the reference speed you are flying on your ASI. So a more complete statement is "IAS being constant, TAS increases as altitude increases".

POH 'Contradiction': Your C-152 POH Cruise Performance table seems to indicate a reducing TAS as altitude increases. This is because the page does not indicate the IAS to be flown, instead it gives you an RPM setting for level flight and the resultant TAS (and inflight we accept the resultant IAS). Take 2100rpm, the TAS is:
86KTAS at 2000ft
85KTAS at 4000ft
84KTAS at 6000ft
83KTAS at 8000ft
82KTAS at 10000ft
81KTAS at 12000ft
Seems to reduce with altitude, but calculating the IAS equivalent to the TAS and Alt at ISA using the 'aerotoolbox' webpage gives:
86KTAS at 2000ft = 83KIAS
85KTAS at 4000ft = 80KIAS
84KTAS at 6000ft = 77KIAS
83KTAS at 8000ft = 74KIAS
82KTAS at 10000ft = 71KIAS
81KTAS at 12000ft = 67KIAS

Here are values of TAS for fixed IAS, good for any airplane, and clearly showing the TAS increasing with altitude:
77KIAS
= 79KTAS at 2000ft
= 87KTAS at 8000ft
= 93KTAS at 12000ft

In the cockpit you can either fly a speed on the ASI (adjusting the RPM to suit) or you can fly the manufacturers recommended RPM and accept the resultant speed on the ASI.

As per the POH page, you input altitude and RPM (with mixture leaned) and the table gives you the Power, TAS, and Fuel consumption. The chart gives you this data for an airplane weight of 1670lbs and various RPM settings of the engine at even thousands of feet altitude ranging from 2,000 to 12,000 ft. (And for ISA, ISA-20degC, ISA+20degC.)

An input of altitude - RPM - OAT/ISAdev gives the TAS and fuel consumption per hour. These are the parameters you need alongwith forecast winds to make out a navigation and fuel flight plan. TAS and Wind gives Groundspeed/leg times and Heading to steer, and further with Fuel cons/hr it gives you the fuel for trip etc. The fact that the table outputs no IAS is of no particular disadvantage in the cockpit.

The data could have been shown by using a fixed IAS strategy, replacing the RPM settings with different IAS values in which case you would input altitude, IAS, Temp/dev the chart would output an RPM, TAS and Fuel consumption.

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  • $\begingroup$ You have explained the difference between IAS and TAS, but you haven’t explained why TAS is higher at high altitude. $\endgroup$ Dec 19 '20 at 0:51
  • $\begingroup$ 1. That was already explained by other posts, but i should add it, you're right. 2. The attempt was to explain why his POH shows lower TAS with higher altitudes which seems to contradict the fact of TAS increasing with altitude. I shall make the answer more explicit. @AlexK may comment if he's read my answer. $\endgroup$
    – skipper44
    Dec 19 '20 at 9:04

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