Archive for the ‘ thinking ’ Category

High altitude without turbo

I was thinking one day about the Bohannon B1. It is basically a modified RV (Harmon rocket something) with very high power to weight ratio and that’s it. This plane climbed to something like 14 km.

So consider this (high excess power) case hypothetically:
– Airplane with high aspect ratio (low span loading) with high power engines with high power to weight ratio. Example: Chevrolet LS9 (600 hp).
– If the plane can maintain level flight with minimal power. 35000 ft we have remaining power 0.2 * 600 = 120 hp.
– Diamond flies nicely with 120 hp, actually 90 hp is quite sufficient for it for normal cruise speed. With lower span loading much less should keep the plane level.

So now the naysay would be “nah, LS9 can not sustain 600 hp continuous without breaking”. However, 120 hp is hardly 600 hp continuous even if the engine is at full throttle and giving all it can at the altitude. It is still stressed only for the 20 percent power.

Same engine, with single stage turbocharger, it should be possible to extend this quite a bit further. With two stage turbocharger even higher altitude should be possible, 70000 ft might be feasible given that the other challenges that come with the altitude are solved somehow.

So you could have a 1200 hp airplane with 240 hp used at altitude for cruise (in case of twin). This should give a quite generous cruise speed at the altitude given that the props are big enough (disc loading low enough).

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Airplane design from structural efficiency point of view combined with aerodynamics point of view – multi-domain optimization

So far I have been looking only the aerodynamics side, but it is quite evident that compromises are needed on the aerodynamics side to achieve the best structural efficiency. I think one good example is Virgin Global Flyer (Scaled Composites model 311). I have not analysed yet the structure, but common sense says that trimaran has weight placed more evenly along the wing span avoiding a very large point load in the middle where the single fuselage would normally exist. The trimaran may have more wetted area than a single fuselage, but on the other hand, weight savings in the very high aspect ratio wing and space gains for the extra fuel are in this concept very important factors.

I find the trimaran configuration quite interesting – several different engine placement configurations for example can be used with this configuration without changing the aerodynamic shape of the concept very much. It is also interesting because it allows placement of the main gear away from the center fuselage and thus provides greater stability on the ground when the aspect ratio is high even if there is fuel placed to the wings very far away from the center of gravity. And as can be seen the same design suits several different missions: Global Flyer is very much like White Knight 2 with SpaceShipTwo under it on the center. Almost the same configuration, adapted to different kind of mission for very different kind of parameters (Global Flyer = long range cruise, White Knight 2 = optimized for climb).

Global flyer drawing Google found from some site
Wikipedia has another great photo, this is from front

The configuration is not really so new and not so unproven either, as people might expect, here is one example where a similar configuration has been used a long time ago:
Northrop Widow
The only difference here is that the Northrop Widow was optimized for different mission than either of the abovementioned and that it had piston engines in front of the outer “fuselages” which were interconnected from the tail section similarly than in Adam A500 whereas the Global Flyer and White Knight Two have two separate tails. It is quite apparent why the tails are separate in these aircraft – because the outer fuselages are placed so widely apart from each other, connecting the tails would have made the tail unnecessarily large which would have caused negative effect for the drag despite it would have had fewer intersections. On the other hand, I have been looking different HALE concepts, and it is quite apparent that the number of intersections is not the major drag source in high altitude aircraft, but the induced drag is, and to minimize induced drag, more intersections can be allowed as the penalty from them is lesser than limiting the aspect ratio would be. This is why there are even some concepts considered at the moment which have wing struts – even if everybody knows that they produce drag, in some concepts, the significance of that drag can be proportionally small whereas the increased aspect ratio has major effect on minimizing the total drag of the aircraft. HALE aircraft have to be quite different than those which are designed to cruise at low altitude, the drag percentages of each contributors are quite different and “one size does not fit all”.

It is quite interesting area to explore when the structural efficiency is added to the equation in addition to the aerodynamics and the result is a compromise on both structures and aerodynamics instead of being optimized for either aerodynamics or for structures. The mission parameters tend to heavily affect both and best suited results can be achieved by combining these two and by knowing the intended use exactly, potentially bigger gains can be realized than in a concept that is a general purpose in everything (GA = GENERAL aviation).

Why Diamond uses Wortmann FX63-137?

I have been thinking over and over again why Diamond has chosen the Wortmann high lift airfoil FX63-137 on its aircraft. However, I am suspecting what might be the reason (not confirmed though since anybody on Diamond booth e.g. in Oshkosh is usually never able to answer to my questions). Here is my theory about it:
– The FX63-137 has high L/D at fairly high alpha and thus Cl (as the airfoil is such that the Cl rises rapidly as a function of alpha). This is maybe not the best configuration for cruise where a low drag bucket at low Cl is desirable. On an airfoil which has best L/D at low Cl, the climb has more D component (because high lift devices cause drag) and while getting more L with high lift devices. It might be close to the optimal climb optimisation on the chosen aspect ratio on those planes and compromise is drawn to cruise and it is not seen as a bad thing because competition is not faster but usually slower, it does not take so much to win e.g. a C172 in efficiency and speed after all. So it might be that with a lower drag cruise airfoil e.g. DA42NG with the very heavy diesel engines might have somewhat poorer climb rate on single engine situation or it might not climb alltogether if the airfoil was not optimised to provide low drag on high Cl.
– Comparison between the DA40 and Cirrus SR20 kind of potentially shows this: the Diamond shows significantly better climb rates with a quite similar AR and quite similar wing loading (SR20 takes some toll on that, but not that much in comparison if a light loaded SR20 and heavy loaded DA40 is compared), despite of the fact that the SR20 has more sophisticated flaps and the SR20 has 20 hp more engine power available.
– This can be also evidenced on best climb rate speed: with similar wing loading, the best climb rate speed is much higher on the SR20 than it is on the DA40, which partly indicates that the sweet point in the L/D occurs at lower alpha on SR20 than on DA40. SR20 also requires quite accurate angle of attack and thus speed to climb optimally whereas the DA40 is not that critical which would also indicate that the low drag bucket of the FX63-137 is broader than on the (according to UIUC data site) Roncz airfoil on the SR20.

So this is just my home-brewn theory style thinking, is based on collected information and my experience with flying the Diamond DA40, DA42 and Cirrus SR20 and SR22. I might be wrong as always, but here is some food of thought if you have been thinking why there is this airfoil with high L/D at high Cl and the airfoil also has fairly high pitching moment which some find undesirable because of for example trim drag.

>Random thinking about the feasibility of fabrication of a partial pressure suit

>I have been thinking this topic for quite a long time and my old conclusion was that it is not feasible. However, there is very little information about pressure suits out there, but looking at the very little there is about the old Mercury suits etc., I have got indications that actually fabrication of a partial pressure suit might be feasible. Buying one might not be feasible because of lack of availability and insane pricing.

I would like to learn more about the topic, but in the Internet at least, there is very little or nothing. If someone has some insight, please leave some comments.

Random thinking about the feasibility of fabrication of a partial pressure suit

I have been thinking this topic for quite a long time and my old conclusion was that it is not feasible. However, there is very little information about pressure suits out there, but looking at the very little there is about the old Mercury suits etc., I have got indications that actually fabrication of a partial pressure suit might be feasible. Buying one might not be feasible because of lack of availability and insane pricing.

I would like to learn more about the topic, but in the Internet at least, there is very little or nothing. If someone has some insight, please leave some comments.

>Looking at historical data

>One interesting aircraft in the historical data:

Lancair Evolution

P/W Power to weight ratio 7.54 kg/kW, 12.28 lbs/hp.
W/S Wing loading 142 kg/m^2, 29.05 lbs/sqft
Stall speed 61 kts
Empty weight to gross weight ratio: 0.55
Fuel to gross weight ratio: 0.2
Aspect ratio: 10.3

Aircraft with 12.28 lbs/hp power loading and 29.05 lbs/sqft wing loading in other words can
be made to climb, and it can also meet FAR 23 in stall speed requirement (61 kts). According to an article, the LC Evolution demonstrated glide ratio of 1:22 which is amazing compared to the competition, especially achieving this with only AR=10.3.

With these parameters, a smaller Rotax powered twin aircraft would be sized as follows:
Engines: 2 x Rotax 912UL, each turbocharged at 100 hp
Gross weight: 1116 kg
Empty weight: 0.55 * 1116 kg = 613 kg
Fuel weight: 0.2 * 1116 kg = 223 kg
Fuel volume: 314 l
Wing area: 1116 kg / 142 kg/m2 = 7.75 m2
Useful load (including fuel): 503 kg
Useful load full fuel: 280 kg
Endurance: 10.4 hours

Challenges:
– achieving the stall speed of 61 kts, requires very high Clmax for the flapped airfoil
– achieving > 20 glide ratio with lower Re, requires higher AR most likely
– achieving positive climb rate with single engine
– achieve the Clmax with wings that carry two engine pods on them (blanketing potentially flap and part of the wing).
– The fuel potentially does not fit inside the wing of this low wing area.

What it shows:
– Still with this high wing loading, it would be possible to fit three adults on the plane with full fuel. The result is not at all bad compared to any production aircraft.
– Empty weight looks realistic taking in account there are two Rotax engines on the craft. It is higher than it would be if it was relatively as lightweight as a Dynaero.

Bottom line: The parameters of the Lancair Evolution are very impressive and inspiring.

Realism hits:

Reduce the wing loading to 120 kg/m2
Wing area becomes: 1116 kg / 120 kg/m2 = 9.3 m2

-> It ends up in the magic 9.3 m2 wing area I have ended up from many directions already several times before.

Looking at historical data

One interesting aircraft in the historical data:

Lancair Evolution

P/W Power to weight ratio 7.54 kg/kW, 12.28 lbs/hp.
W/S Wing loading 142 kg/m^2, 29.05 lbs/sqft
Stall speed 61 kts
Empty weight to gross weight ratio: 0.55
Fuel to gross weight ratio: 0.2
Aspect ratio: 10.3

Aircraft with 12.28 lbs/hp power loading and 29.05 lbs/sqft wing loading in other words can
be made to climb, and it can also meet FAR 23 in stall speed requirement (61 kts). According to an article, the LC Evolution demonstrated glide ratio of 1:22 which is amazing compared to the competition, especially achieving this with only AR=10.3.

With these parameters, a smaller Rotax powered twin aircraft would be sized as follows:
Engines: 2 x Rotax 912UL, each turbocharged at 100 hp
Gross weight: 1116 kg
Empty weight: 0.55 * 1116 kg = 613 kg
Fuel weight: 0.2 * 1116 kg = 223 kg
Fuel volume: 314 l
Wing area: 1116 kg / 142 kg/m2 = 7.75 m2
Useful load (including fuel): 503 kg
Useful load full fuel: 280 kg
Endurance: 10.4 hours

Challenges:
– achieving the stall speed of 61 kts, requires very high Clmax for the flapped airfoil
– achieving > 20 glide ratio with lower Re, requires higher AR most likely
– achieving positive climb rate with single engine
– achieve the Clmax with wings that carry two engine pods on them (blanketing potentially flap and part of the wing).
– The fuel potentially does not fit inside the wing of this low wing area.

What it shows:
– Still with this high wing loading, it would be possible to fit three adults on the plane with full fuel. The result is not at all bad compared to any production aircraft.
– Empty weight looks realistic taking in account there are two Rotax engines on the craft. It is higher than it would be if it was relatively as lightweight as a Dynaero.

Bottom line: The parameters of the Lancair Evolution are very impressive and inspiring.

Realism hits:

Reduce the wing loading to 120 kg/m2
Wing area becomes: 1116 kg / 120 kg/m2 = 9.3 m2

-> It ends up in the magic 9.3 m2 wing area I have ended up from many directions already several times before.