Archive for the ‘ untested idea ’ Category

Idea: Series hybrid in airplane using auto engine and avoiding the pitfalls of auto conversions

I have been thinking this back and forth now quite some time. This idea is quite simple, the purpose is to fix the most critical problem with auto conversions, achieve better aerodynamics, propeller placement and mass and inertia distribution.

Auto conversions most often fail, no surprise, because of the reduction gear or belt. The core engine is not the root cause in the problems and many problems with the reduction belt or gear system can not be seen beforehand because the dynamics of the vibrations of the engine, propeller and their inertia forces affecting each other is a bit more complicated than one could think at first – it is not that simple to make these parts to last for hundreds or thousands of hours.

So we came up (with Kate, we usually talk with Kate about these things and we kind of invent these things together, I usually happen to be the one who writes them down – and it is usually so that Kate is the opponent into which I test my idea’s feasibility before I write it here) with the idea of having a auto engine, possibly a diesel engine, running at constant power, most likely exactly at the optimum point of the engine, always. Then all the power variation would come from the electric motors which would drive the propellers. The idea is that the diesel engine only runs a generator.

The downside of this idea is the additional weight from the generator, batteries, motor controllers, electric motors and the props (depending how many electric motors are used, it is also possible to use just one if that is preferred). However, there are two several things possibly good about this:

– First the diesel engine burns less fuel, resulting smaller fuel tanks.
– Secondly the gearbox system is saved. The gearbox system can be very heavy duty in a high power aircraft engine and they still have tendency to fail. Possibly something like 40-50 kg is saved straight away.
– Thirdly the aerodynamic advantage – optimal aerodynamic shape without using long extension shafts and couplings to deal with the dynamics of the rotating shaft connected to a non-optimally rotating propeller and the power pulses of the diesel engine. Now there is the chance to put the engine anywhere in the airframe where it best fits and propeller drive don’t need to be considered at all.

Then there is the redundancy thing. Brushless DC electric motors usually never fail, but the prop can still fail in bad circumstances. Therefore having two independent props for the one diesel engine could be advantageous. Same thing with the batteries – if the diesel engine fails, the batteries could be sized such that the aircraft can fly without the diesel engine for example for 30 minutes in level flight. That might be enough in most cases to get safely on the ground, except on middle of an ocean. The most likely place for the engine to fail is the takeoff. This takeoff stress would never happen with this engine configuration – the engine would be run always at optimum and safe power, never on takeoff power. The extra power for the takeoff can be easily taken from the batteries if they have proper capacity and the electric motors are powerful enough. On takeoff the batteries at full power are not discharging that quickly, because the diesel engine is recharging the batteries at the same time. The takeoff power can be rarely used for longer than 5 minutes on an aircraft equipped with Lycoming engine either, so having a limited period of time for the full power is not that big problem.

Generator and electric motor can have very high efficiency, and the gap to a efficiency of a reduction belt system is not that great. Best electric motors (though heavy ones) are around 98% efficient.

On descent the diesel engine could be shut down providing there was enough battery capacity. The motors could actually regenerate also batteries when the pilot wants to decelerate the plane.

Maintenance cost would be like a single engine aircraft, but the reliability geared towards a twin. Of course there is the one little fine print: the battery pack is expensive and it has an expiration time and date, unfortunately. But nothing is perfect and without compromises.

Any comments about this idea? This surely would not be a racer as the power to weight ratio would be rather poor, but anyhow I am thinking, providing it would be efficient enough to climb adequately, this would be a quite economical thing to fly and also easy conversion-wise, almost stock auto engine would be okay, no reduction gear and prop installation and an assembly that takes the push or pulling loads, would be needed. Also waiting on the airport would not waste any energy, since props can be completely stopped when the plane does not need to move. For example Lycoming IO-360 consumes about the same amount of gasoline per hour when waiting on IFR clearance on the ground than our Toyota Prius car on highway. Consuming zero amount of fuel when still on the ground, but still being ready, would save some liters.

And answer to the question, why diesel and not gasoline when gasoline engines can be run very lean and quite great specific fuel consumption values can be achieved in optimal conditions – it is quite simple: availability of the 100LL/Avgas seems to be becoming poor. There has been three 100LL operators in Finland, but two of them decided to discontinue this year. There is only one left. When that only one decides that it is not profitable enough, there is no 100LL available for anybody and the whole country’s fleet of Lycoming and Continental based planes are grounded. The Jet-A1 is not going anywhere, so engine that can burn the jet fuel would be a safe bet. Jet engine, turboprop, or turbofan are out of the question because those are not available in meaningful sizes and power classes – there is not a small turbofan that would have high pressure ratio and bypass ratio available, nobody manufactures such a thing. And it is unlikely anybody will in the future because this personal flying all is a very niche market unfortunately until it changes for better (if it ever does).

The implementation possibilities have challenges; namely no such electric motor available (would require custom motors possibly), etc. And the weight also causes penalty for the efficiency and speed of the plane. But the power to weight ratio will be with this arrangement a lot better than on a pure electric aircraft. And pure electric aircraft is feasible, why an electric aircraft with a generator and a fueltank added would not be.

And by the way, even if it is first of April at the time of writing this, this blog post is not an April fool.

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>Suction stabilization for low fineness ratio pusher engine pod

>I got an idea how to achieve suction to the read of the engine pod.
* The prop is located just after the laminar-turbulent transition to the pod and the remainings of the pod is a very large spinner which is open from the center.
* The air tunnel inside the pod has venturi-shape.
* There are tunnels that connect the venturi tube and the ring that is supposed to have suction.
* Airflow (which is used to engine cooling) inside the venturi (helped with the propeller part that is inside the pod) causes suction to the rear of the pod. The air exits at the end of the venturi tube, which happens to be the center of the spinner.
* The exhaust in the center makes the cut aft end of the spinner to still maintain low drag, it functions in the same way as the rear cut fuselages in jets

I have not tested this idea and don’t know it it would work, but I think it would be pretty easy to try out in the model scale, even with an electric motor. This interests me enough that I think I am going to try it out of someone doesn’t tell me (with better knowledge, as a fact that has been proven and tested) that it is not gonna work.

I hereby license this invention under the terms and conditions of GNU General Public License, version 3, or any later version. (C) 2008 Karoliina Salminen. All rights reserved. By reading this text, you aknowledge this and agree with the terms and conditions of the GPL license.

Suction stabilization for low fineness ratio pusher engine pod

I got an idea how to achieve suction to the read of the engine pod.
* The prop is located just after the laminar-turbulent transition to the pod and the remainings of the pod is a very large spinner which is open from the center.
* The air tunnel inside the pod has venturi-shape.
* There are tunnels that connect the venturi tube and the ring that is supposed to have suction.
* Airflow (which is used to engine cooling) inside the venturi (helped with the propeller part that is inside the pod) causes suction to the rear of the pod. The air exits at the end of the venturi tube, which happens to be the center of the spinner.
* The exhaust in the center makes the cut aft end of the spinner to still maintain low drag, it functions in the same way as the rear cut fuselages in jets

I have not tested this idea and don’t know it it would work, but I think it would be pretty easy to try out in the model scale, even with an electric motor. This interests me enough that I think I am going to try it out of someone doesn’t tell me (with better knowledge, as a fact that has been proven and tested) that it is not gonna work.

I hereby license this invention under the terms and conditions of GNU General Public License, version 3, or any later version. (C) 2008 Karoliina Salminen. All rights reserved. By reading this text, you aknowledge this and agree with the terms and conditions of the GPL license.

Idea: Full span flaps

Full span flaps with flapped ailerons:

In board wing has 60% span fowler flaps. Outboard wing, the remaining 40% consists plain flap type flaperons with similar mechanism than used in Mini-Sytky.
deltaClmax_fowler = 0.6 * 1.67 + 0.4 * 0.9 = 1.362
For airfoil with Clmax 1.2 the maximum Clmax on landing configuration is thus 1.32 + 1.362 = 2.68
This allows smaller wing area and higher wing loading to be used without sacrificing takeoff and landing performance too much.
Another variation with single slotted flaps:
deltaClmax_singleslotted = 0.6*1.18 + 0.4*0.9 = 1.06
+1.06 in Clmax still is a very good value and better that would be obtained with full span flaperon (+0.9). For airfoil with Clmax of 1.32 this yields Clmax of 2.37.
This idea has not been tested in practice and is not guaranteed to work.
Effects on aircraft:
Aircraft with 60% span plain flap and Wortman FX 38-153 (no full span high lift device):
Clmax = 1.3 + 0.9*0.6 
deltaClmax = 0.54
Clmax => 1.84
86 hp required for 200 kts cruise
wing loading: 92 kg / m2
wing area: 7.2 m2
stall speed: 55 kts
design cruise: 200 kts
Cdtot = 0.011 (with boundary layer suction)
Same aircraft with full span flaperon and Wortman FX 38-153:
Clmax = 1.3+0.9 =  2.20
Same aircraft parameters:
76 hp required for 200 kts cruise
wing loading: 110 kg / m2
wing area: 6 m2 
Aircraft with full span flaps with slotted inboard section:
Clmax = 1.3 + 1.06 = 2.36 
Same aircraft parameters:
74 hp required for 200 kts cruise
wing loading: 118 kg / m2
wing area: 5.6 m2
Aircraft with full span flaps with fowler inboard section:
Clmax = 1.3 + 1.362 = 2.66
70 hp required for 200 kts cruise
wing area: 5 m2
wing loading: 134 kg / m2
For the most extreme case theoretical savings over usual configuration:
Power = 86-70 = 16 hp (18%)
wing loading: 134-92 = 42 kg/m2 (31%)
wing area: 7.2 m2 – 5 m2 = 2.2 m2 (30%)

>Idea: Full span flaps

>Full span flaps with flapped ailerons:

In board wing has 60% span fowler flaps. Outboard wing, the remaining 40% consists plain flap type flaperons with similar mechanism than used in Mini-Sytky.
deltaClmax_fowler = 0.6 * 1.67 + 0.4 * 0.9 = 1.362
For airfoil with Clmax 1.2 the maximum Clmax on landing configuration is thus 1.32 + 1.362 = 2.68
This allows smaller wing area and higher wing loading to be used without sacrificing takeoff and landing performance too much.
Another variation with single slotted flaps:
deltaClmax_singleslotted = 0.6*1.18 + 0.4*0.9 = 1.06
+1.06 in Clmax still is a very good value and better that would be obtained with full span flaperon (+0.9). For airfoil with Clmax of 1.32 this yields Clmax of 2.37.
This idea has not been tested in practice and is not guaranteed to work.
Effects on aircraft:
Aircraft with 60% span plain flap and Wortman FX 38-153 (no full span high lift device):
Clmax = 1.3 + 0.9*0.6 
deltaClmax = 0.54
Clmax => 1.84
86 hp required for 200 kts cruise
wing loading: 92 kg / m2
wing area: 7.2 m2
stall speed: 55 kts
design cruise: 200 kts
Cdtot = 0.011 (with boundary layer suction)
Same aircraft with full span flaperon and Wortman FX 38-153:
Clmax = 1.3+0.9 =  2.20
Same aircraft parameters:
76 hp required for 200 kts cruise
wing loading: 110 kg / m2
wing area: 6 m2 
Aircraft with full span flaps with slotted inboard section:
Clmax = 1.3 + 1.06 = 2.36 
Same aircraft parameters:
74 hp required for 200 kts cruise
wing loading: 118 kg / m2
wing area: 5.6 m2
Aircraft with full span flaps with fowler inboard section:
Clmax = 1.3 + 1.362 = 2.66
70 hp required for 200 kts cruise
wing area: 5 m2
wing loading: 134 kg / m2
For the most extreme case theoretical savings over usual configuration:
Power = 86-70 = 16 hp (18%)
wing loading: 134-92 = 42 kg/m2 (31%)
wing area: 7.2 m2 – 5 m2 = 2.2 m2 (30%)