Hybrid aircraft

The idea of the system comprises of a turbo generator per engine and an additional electric motor behind the tail.

Configuration:
Two gasoline engines, one per wing.
One Brushless DC electric motor, behind the tail, engine size around 15 kW. Does not require any drive shaft because the motor itself is so small and lightweight, that it can be attacted directly to the tail.
Battery that can deliver full power to the electric motor for 3 minutes.
Motor controller for each electric motor.

Possible additions:
Two wing tip turbines, one per each wing tip. Electric motor size ~5 kW.
These can produce power on cruise for the middle pusher motor.

The center pusher motor could drive a unducted fan which would have diameter around 1/3 of the diameter of the fuselage body. See NASA tech paper wake propeller, why. The fan would require adjustable pitch for each blade, so it could be changed from climb condition to cruise condition for the cruise phase (otherwise it would cause drag penalty).

Additional idea:
- the wing tip turbines could be used in case of engine failure for thrust vectoring - one small wing tip engine producing thrust could make the asymmetric thrust condition symmetric without causing drag penalty with deflected rudder.

Wing tip turbine

http://www.freepatentsonline.com/4917332.html
http://www.aiaa.org/content.cfm?pageid=406&gTable=mtgpaper&gID=80452

X-plane as educational program

It seems that X-plane educates aerodynamics, what to expect and think about different things. I was originally saying that I am not so interested in transonic region but rather interested in high altitude. I have been reading about these, but some little things like tinkering with X-plane can cause heureka moments.

And here is what happened:
I have a model of my twin concept in X-plane simulator (obviously, why wouldn't I). So I set in the latest incarnation the engine critical altitude to 50000 ft (which is feasible with two turbos in cascade plus the mentioned electric turbo compounding). I used 110 hp per side (equivalent of Rotax 912ULS equipped with two turbos doing turbo normalization plus intercooler and after cooler).

I was reading Roskam couple of days ago and noticed that the transonic drag is not a problem if the speed is mach 0.2 or below or not that much above that, e.g. 0.3-0.4 is still quite fine. So I was thinking that maybe it doesn't get that high that it would become a consideration.

So so obviously, I put the plane model to climb to 55000 ft with autopilot. I had previously added the mach meter to the hud. I came back checking how it flies after couple of tens of minutes. And oops: mach 0.56 when level at 55000 ft. The IAS was barely 100 kts. TAS was a quite a bit higher.

Then, I was thinking what happens to the Reynolds number. Indeed it gets smaller with altitude increasing. But interesting thing is what really happens, to which number it gets. I verified with atmosphere calculator, that indeed, the interesting Re range for this kind of concept with the AR=14 wing, it becomes 600000 - 1600000. That is _very_ low for an aircraft, which is full size and not a RC-model. So the low Re becomes after all a major consideration.

How a plane with AR=14 flies at 55000 ft? It requires _full_ trim aft (meaning nose high) to get the plane keep level - in this model. It became quite apparent that indeed, the tail volume coefficient is a more major concern at high altitude than at low altitude. And the control authority that felt fine at low altitude was not so fine at high altitude.

So this is what we have:
- High performance low Re airfoil is very necessary
- Cd at high lift coefficient is an important design point, the airfoil needs to be designed so that it gives high L/D at high lift coefficient rather than at low lift coefficient like for example NLF414F is targeting.
- A big tail with long enough moment arm
- Propeller with large diameter and possibly more blades than usual, e.g. 5 blades
- And of course, two turbos, intercooler, aftercooler, generator, battery, electric motor and a shaft between the prop and the engine.

Btw, my model is not yet available for download because it is not perfect, and it has couple of problems. It is very hard to get the splines right with straight sections edited by hand, and e.g. engine nacelles look really terrible at the moment. Anyway, it is a fun way for trying out things in practice.

Hybrid turbo compounding

Kate invented one day that why the turbo compounding could not be implemented with electric motors, because that way the usually unfeasible gearbox from normal turbo compounding becomes unnecessary and the gearing is instead implemented with the electric motor and the generator where the generator rotates at higher revolutions than the motor that is used to decrease the load the combustion engine sees.

We were in assumption that this was a new invention, but it seems that it has been used in heavy machinery already, e.g. by Catepillar. This in turn also means that it is feasible.

The challenge would be how to place the generator to the shaft of the turbo. Usually turbos do not have a place where to fit the generator but they are closed packages which are not easily modifiable.

The idea would be to increase fuel efficiency with the compounding and increase the shaft horse power without loading the combustion engine anything more. The electric motor could have an additional lithium polymer batter pack which could increase the power even more on takeoff, so the plane would have on critical take off situation somewhat more power than the combustion engine can output, so in other words, for example getting 80 hp out of a 60 hp HKS700E.

This would result that using impossibly small engine power would become a possibility in a wider variety of airframes. On a twin 2 x 80 hp is a lot more than 2 x 60 hp, single engine performance on 60 hp is very poor in any case without any tricks done to increase the power temporarily.

A quite small lithium polymer battery pack would be enough since assuming 300 fpm climb rate on a single engine, this results 3 minutes to 1000 feet AGL where it should be safe to turn back to the runway and perform landing even with a very low power output of a single engine. So it would be well enough for the extra power from the battery pack last only for 3 minutes. This kind of battery pack would not be that heavy, and the brushless DC electric motor is also pretty lightweight.

Any comments on this?

MGS L285 mix ratio

The ratios for L285, H285, H287 are:

100:50 by volume
100:40 by weight

I am doing some little layup today, so I decided to write the mixing ratio to my blog. I always tend to forget it and have to search from the MGS documentation. Now it is here in my blog. This same ratio applies to my MGS L285/H287.

Pushing the limits

Why am I interested in pushing the limits instead of doing a known safe solution (meaning thinking more than usual aerodynamics, low weight, low power, low production cost, manufacturing technology, integrated advanced avionics)?

This could be asked with a counter question: why not? What would be the motivation of replicating some existing aircraft with existing technology with learning nothing new?

Someone might say that because of business, to sell these things. But to be sincere, I believe that this business scene is already congested, and it there is anecdotal evidence that the profitability of this business is questionable at times. And there are already number of manufacturers which are doing this and their planes are just fine. If one wants an average plane which is not supposed to push any limits, there is plenty of selection with all kinds of colors, at least one can choose if the plane is white or white or maybe metallic gray, and gluing red or blue tape stripes is optional.

If we look back 50 years and think what has been happening in general aviation. We can conclude that we are entering yet another year with virtually a very little or no progress. Engine technology is still the same, aerodynamics have only moderately improved (laminar flow wings are nowadays somewhat utilized (Cirrus and Cessna 400), but not laminar flow fuselages) but not much, and avionics are still the same (but only gradually improved) rather than inventing a question for the answer that this is always been done that way and answering to it before it is asked.

One could ask, what is there to improve? Why to change anything. If one has nothing better to offer, then the whole venture is worthless. The easiest way to copy and not learn anything new is not to design own aircraft, but buy an existing one. If time spent for engineering is counted with any kind of monetary value, then purchasing an existing aircraft is potentially also the cheapest way to get flying. It is also the safest way, and the chances of big disappointment is small. You get what you pay and it is potentially a good compromise and biggest bugs are already fixed.

But one thing is that for some, pushing the limits is the meaning of life.

I see that the future of general aviation is not very bright without radical new designs which are better performing than the current aircraft, a lot more economical than current models, and essentially less expensive to buy and operate than the current aircraft.

This requires couple of breakthroughs to happen. I am not interested in solving all of them, I do not have unlimited time and can not have solution for everything.

I rather prefer to think everything through a filter which is the compromise I have found best suitable during the couple of years I have been thinking what do I want and now I am finding reasons why I want it. As the answer to the question can not be seen as a singularity, to make the others with a different experience base and collection of individual goals to understand the questions and the answers in this specific case in a give time fragment, can be seen as a challenging endeavor which leads to a setup where all the sides of the multidimensional coin are not fully seen.

When I know what do I want, I can more easily tackle down, what is the minimum resemblance to what I want which still is an acceptable compromise but is feasible technologically and economically and evidently this is a moving target which evolves in the flow of time.

The limits being pushed also evolve but the end result is a snapshot of the broken limits of that time and the frozen design parameters of the evolving concept and today's limits no longer exist as limits but are by then generally accepted known solutions. If this did not take place, the eventual outdatedness of the design would outweight the thought benefits of the whole reason of doing it in the first place and the understanding of the question of why to do it would be remaining essentially unanswered in a light of the evolving circumstances where the answers are integrating variables rather than constants.

The only sensible way to follow the flow of things is to ensure being a step further than the current state of the art, otherwise the train left the station before it was built and one arrived there to travel to a place which no longer exists. This way the coin has a chance to drop the right side up in a place where the relative up is defined by the eye of the beholder, which by definition, defines the answer to the questions who I am and what do I want.

Minimal twin

In the mean time, on the back of my head, I have also been thinking the twin concept. What is the minimum power feasible for the twin for being safe in single engine situation, and what can be the maximum weight and maximum wing loading of a plane which is equipped with two HKS700E engines (only 60 hp each).

Known thing is that Diamond DA42 climbs still at 22 lbs/sqft wing loading and 24 lbs/hp power loading on single engine. However, there is quite a bit more excess power on 135 hp Thielert than on a 60 hp engine. I am feeling that I am getting too optimistic results from the sizing equations with either Raymer or Anderson method.

I have estimated that the plane should not weight more than 700 kg (according to the equations) to still be able to take off and climb with single engine. This may be too optimistic figure, I have been thinking that the limit might be rather near 650 kg or maybe even a bit less.

Thinking pessimistic: the plane can have positive climb rate with 60 hp single engine mode if the gross weight is 600 kg. That gives:

600 - 55 kg - 10 kg - 55 kg - 10 kg = 470 kg for the airframe + useful load excluding engines.

For useful load, minimally needed is:
- Two big adults, 95 kg including heavy clothes per each
- 5 kg baggage per each
- 120 liters of gasoline = 85 kg

This becomes:
95 kg * 2 + 10 kg + 85 kg = 285 kg.

For the plane to be minimally useful, it must be able to carry 285 kg in addition to its own weight. There are two engines and to have useful endurance the amount of fuel has to be double the size of a single engine plane.

The airframe + systems maximum weight excluding engines then becomes:

470 kg -285 kg = 185 kg

This means that the airframe + systems excluding engine can only weight 185 kg. This is a very hard goal to achieve.

The aircraft empty weight then becomes:

185 kg + 65 kg + 65 kg = 315 kg

The empty weight to gross weight ratio becomes:
315 kg / 600 kg = 0.52

This ratio is very challenging to achieve for a twin where the airframe must be carrying in addition to the occupants instead of one engine, two engines, and their fuel.

If we could still take off at 650 kg, then this becomes:
Airframe weight can be increased with 50 kg: 185 kg + 50 kg = 235 kg

235 kg + 65 kg + 65 kg = 365 kg

Looks like now we are talkin. This looks like a figure which might be theoretically possible, even though this is still very hard goal. As seen on ultralight planes, achieving empty weight under 300 kg is very hard. Adding extra engine on top that requires aircraft that is as lightweight than best ultralights equipped, plus can still take the additional engine.

But this is just theoretical thinking and whether or not it may be feasible, the discussion can continue:

The empty weight to gross weight ratio then becomes:

365 kg / 650 kg = 0.56

Historical data shows that at least on a bit larger aircraft, the 0.56 value is pretty well achievable.

Lets consider now the performance for the 650 kg case:

Single engine produces only 60 hp power. Only the excess power can be used for climb. This means that in a side slip of asymmetric thrust and climb angle of attack, the total drag (drag due to lift + fuselage drag) must be less than the thrust of 60 hp at best climb speed with a propeller that has efficiency of 0.7 (for pessimistic evaluation, I prefer to not use 0.85) by a large margin, and then the climb rate pretty much becomes from the weight to be lifted and how much excess power is still left.

The power loading for single case would be: 23.8 lbs/hp. This would be about the same as Diamond DA42. The drag must be low in order to ensure that the power needed for level flight is small, and there is excess power for climb, even with very low power.
Then comes the disaster of increasing wing area, this increases drag, but on the other hand, increases also lift. However, to get good cruise performance on the low power, wing size should be as small as possible. So some compromise is needed here. Increase in wing loading has to be accounted with increase in aspect ratio to keep the induced drag the same. Increase of aspect ratio may increase weight, but does not necessarily always do so. For example the earlier mentioned LH10 has very light wings, despite of aspect ratio of 14. So it worths researching on this area. A good design is a synergetic design which combines couple of good things into one good compromise.

I maybe need to redo the calculation yet another time again.

Why I am thinking this?
- For a plane that I would design for myself, I could choose Rotax 912ULS, and get two used engines with about half the price of a new Rotax 912ULS. This would be roughly the cost of a pair of new HKS700E.
- However, if we think a kit-builder who wants to have a twin with shoestring budget. Many aviators are limited with budget (aviators are always rich simply does not seem to be true, and if they originally were, they no longer are after starting spending to flying). So we have been thinking of a concept of a light plane with two engines with good performance. Any twin out there, even used ones, cost many many times more than it would cost to build a plastic one with two little HKS700E engines.
- I think that twin engine aircraft are not so popular, not because they require the additional license, but because people do not opt for the additional license, because the cost of the twin is prohibitive. There is absolutely no twin out there where one could log twin engine time and which would not cost a fortune of a millionaire to own or cost a fortune of of a normal people to maintain and operate.
- It is often explained that twins are more dangerous than singles. However, the context seems to be forgotten. Single engine limits the use of the plane and with two engines, people may often go to more dangerous situations.
- And it is not only a bad thing, consider this: You live in Finland and want to visit for example Greenland. What do you do if you want to fly there by yourself and not to sit as a passenger on an Airbus? You go and start your C172 and head towards Greenland. If the one old-fashioned engine that is almost approaching car engines in reliability, that is there, quits, then you are in biiig trouble. Wouldn't it be great if there was a second engine and you could still fly even if the one failed. Even if the climb rate with single engine is poor, you could still maybe get out of there alive. Your speed would get slow, but also your fuel consumption becomes half because only one engine is drinking the fuel. You actually might make it and your relatives don't need to arrange funerals.

Any comments on this?

Boom tail microlight/LSA

I was thinking which could be a suitable configuration if target would be to the microlight category (Finnish ultralight will be aligned with European microlight), or to EASA LSA category. The US-LSA category is stupid since it has some severe limitations which removes the reason to try to optimize anything - the speed limitation and also the stall speed limitation as clean -> with these limitations, it does not worth optimizing the aerodynamic performance or flap configuration, for US-LSA, the best solution probably would to design a plane, which is as lightweight as possible and which would not have any kind of flaps and which would achieve the stall speed only with the wing area (because that is what the limitation implies anyhow), so the utilized Clmax becomes close to 1.0, which is poor.

The main criteria in this more sane European category is the weight and the second main criteria is the stall speed. These are the most important features, other features are secondary. The performance can not be optimal, but it can be optimized to the constraints given by the weight and stall speed limitations. It would be also necessary to cut the part count to minimum, an inexpensive plane will be for sure more popular than the more expensive one, in the category where the buyers are not the richest people out there (who would anyhow order a Cirrus-Jet), but normal hobbyists who are not swimming in money.

So consider this:
- Plane structure would be based on carbon fiber rods (pultrusion rods).
- The fuselage would not be a structural member of the plane, but rather a baggage pod located below the spars. The twin booms would be a pultrusion rod each. The engine would be mounted to the wing spar rather than to the fuselage. These rods could have aerodynamic fairings on top of them (which also allow space for control cables etc.).
- high aspect ratio wing, which enables good climb rate with low power
- HKS700E engine in pusher configuration
- Fixed pitch pusher propeller behind the pod (but thrust line to the wing spar).
- inverted V-tail in the ends of the two booms, and the tail would connect the
two booms with the help of a pultrusion rod which functions as spar.
- Main landing gear connected to wing spar
- Nose gear located under the pod.
- Wing structure would be solid blue styrofoam, and in wing root there would be a large fairing which contains fuel (on both sides). The wing skin could be either carbon fiber or fiberglass (fiberglass to reduce cost obviously)

Think how many parts this requires compared to how many parts and layup schedules is usually needed. The pod type cockpit could be almost a complete monococue. There would be need for only instrument panel and some structure where one can assemble the pedals. Also the instrument panel, as we know it, does not need to be like it is, a panel. There are other ways arranging instruments in the plane than having a straight panel where everything is put with tiny screws. None of the modern cars use that old-fashioned way anymore. With modern avionics, you don't need a big panel with lots of switches, knobs, circular gauges etc. You can have just two screens which display and control everything.

The only strength needed in the fuselage is for crashworthiness, it does not need to carry any loads, and it does not need to be shaped unoptimally to avoid flutter tendency for example, all structural members are straight lines and separate from the fuselage.

The idea comes from some NASA PAV concepts, but as modified. It also has some influences from the Sunseeker.

The concept could have idea of being as lightweight as possible (the lower power engine also supports this mission) and still being as highly performing as possible (that can be achieved with the light weight and aerodynamics, not so much trust is needed).

So the performance target setting for conceptual design could be:
- beat 100 hp Dynaero MCR-ULC in empty weight with large margin
- be on par with 100 hp Dynaero MCR-ULC in speed (with only 60% of the power available)
- and the rest comes from the category limitations
- be a lot cheaper than most other same category plane on the market
- climb rate 800 fpm (remember that because of the low climb speed, the climb gradient is high despite of the not so high number compared to high performance aircraft)

Compromise:
- beat 100 hp Dynaero MCR-ULC in empty weight
- cruise speed compromised to between 80 hp WT9 Dynamic and 80 hp MCR.
- climb rate 600 fpm

Failure:
- heavier than MCR-ULC
- slower than 100 kts in cruise
- climb rate less than 500 fpm

Anyone interested in a such thing or having ideas (for or against) for a such thing?

Here is an illustration about the idea (15 minutes of Rhino magic):

Aerodynamics is not a bolt-on feature

I became interested about aerodynamics through a experimental project I started building with Kate. It was Cozy MKIV. We have not been building that plane for quite a long time, but we got couple of parts done, for example the canard foams were cut with help from Rauno Viljanen and I managed to do quite poor quality chapter 4 bulkheads with zero understanding what I was doing structurally or otherwise.

Back then there was a concept of "speed modifications" very popular on canard forums, and I think it still is happening there, I haven't followed for a while. They are being invented most frequently by people that don't have even pilot's license yet or don't have flown any aircraft to the date, and they don't necessarily have much understanding on the aerodynamics either.

Back then I was really interested in them, and it felt like magic, you bolt on this and that improvement, and it becomes this and that much faster and more efficient. There were all kinds of concepts like cutting lower winglets, shortening wings, or even placing vortex generators to a laminar flow airfoil. I did not see back then what was wrong and why they wouldn't work as expected. Now I know. They were very entertaining reading, and actually inspired me to start thinking these things in more detail. And I am still on that road. They are not essentially bad but they may not work as the builders expect them to work because they don't understand why they are doing them, but are relying on non-scientific reasons to bolt them in.

Couple of years have passed and I have been reading about aerodynamics and trying to find out how it all works. It occurred to me at one point, that it is not a bolt-on feature you can add to existing design, but aerodynamics is all about the flow. And understanding it as a whole.

Someone might say that "by adding vortex generators, you get 5% fuel savings", that can be true only in a case where the flow otherwise preliminary separates. Good aerodynamics, is not fixing this and that with little this and that, but trying to get it all right and if still a problems persist, try to fix them then with some additional fix.

What happens if you consider adding vortex generators to Cozy MKIV front wing, in other words, the canard? The canard has Roncz RMS1145 airfoil which is about 45% laminar. Depending on where you put the vortex generators, you can vary between 0% laminar and 45% laminar. You can't get more than 45% laminar with that shape.

However, what else can happen is that, the turbulent flow attaches to a higher angle of attack on the airfoil which was designed to maintain its lift even if the laminar flow is disturbed by bugs or rain? You may get some more Cl out of the airfoil with the added vortex generators, and may be able to delay the stall angle of attack some.

But think the whole picture: the main wing-canard relationship was tuned so that the canard always stalls before the main wing. If this does not happen, the plane can enter into deep stall which is not recoverable on the particular type in question. If you delay the stall of the canard to a higher angle of attack, you are trying your luck with the main wing's Clmax and maximum angle of attack before it stalls. And it might be that you can achieve higher angle of attack with the canard than the main wing can function without stalling, and the result is pretty severe, everyone on board most likely die as a result, unless you are super-lucky like some that have survived from a deep stall crash. But wait, there was someone who also dropped from a passenger jet without parachute and survived. I would not try my luck based on the few exceptions.

Same thing what happens if you shorten a wing. Jet fighters have shorter wings and they are faasstt. Right? In case of subsonic aircraft you actually increase induced drag if you shorten the wing. You also increase wing loading, which also increases induced drag, although it reduces the wing wetted area which is desirable for lower drag. But in this case, the increase in induced drag can be such high that the plane actually becomes slower. There was one manufacturer that was doing light aircraft, and they were thinking how to convert their aircraft to LSA. The LSA version had longer wings, and instead of limiting the maximum IAS to 120 kts, the supposed to be LSA version became in fact faster than the plane with the shorter wing.

One could think also that a plane which would have smoothly rounded shape in the wing tip instead of a maybe less elegant looking cut shape would be faster. And surprise might be great when the person would notice that instead of making a faster plane, the plane actually got slower because of the modification. Here is also the thing: what you are trying to achieve - looks or relying someone's claims, or are you thinking what you are going to achieve in terms of flow and how it affects the wing tip turbulence and is what you are trying to achieve beneficial or not. The sharp cut hoerner shape in the wing tip might be there for a reason, it resists the flow from the bottom side to the upper side because of the sharp corner there. Rounding this shape makes the wing tip potentially worse. Only potentially, because you have to consider what is going on, and what you are going to achieve. There is no "yes this is right" and "no this is wrong", because everything affects to everything. But you always should know why you are going to do something. Because it is faster that way is a wrong answer. Right answer is the understanding of why. Would be better to first understand why before doing it rather than understanding why you did it and why you shouldn't have done that.

The point is, that the optimization of aerodynamics requires thinking as a whole. Improving something somewhere may not help if something else is really bad, and it can get worse by uninformed improvement somewhere. Only by knowing what you are doing, as a whole, you can do aerodynamic design which results better performance unless you are very lucky. In some cases, you might be lucky, but you could ruin your results by doing something additional uninformed where the whole picture what is going on is not taken into account.

If you want to do a optimized aerodynamic design, you have to begin with that basis, you can't bolt it on after. Cleaning up a existing aircraft is possible to some extent, but only to some extent, which is very small. An optimum design is a in balance from the aerodynamic and structural standpoint and everything is taken into account in every detail and they are understood as a whole with the whole thing. It is not a puzzle with small pieces you just put together, but a puzzle where the little pieces change every time you change something little.

If you want to clean up an existing airplane, what you need to do is that you have to understand what you are doing, in other words, what you are trying to achieve what you are changing. You have to consider all sides of the change, what it does. Things are not so simple as they might at first seem. And some things are simpler than believed. Impossible - there is no such word. You just can't bend what is possible with pure luck, it does not work in the long run. Understanding what you are trying to achieve and what are the potential consequences in good and bad for every detail helps doing less not so good decisions.

Laminar flow

Design of Fuselage Shapes for Natural Laminar Flow

ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19900016623_1990016623.pdf

Slotted flap design

Slotted flap slot design

AERADE Reports Archive, search keyword slotted

http://aerade.cranfield.ac.uk/ara/dl.php?filename=1947/naca-tn-1395.pdf

http://aerade.cranfield.ac.uk/ara/dl.php?filename=1947/naca-tn-1463.pdf

http://aerade.cranfield.ac.uk/ara/dl.php?filename=1950/naca-tn-2149.pdf

http://aerade.cranfield.ac.uk/ara/dl.php?filename=1949/naca-report-942.pdf

Tractor propeller effect on wing behind the prop

http://adsabs.harvard.edu/abs/1987PhDT........47H
ctn.cvut.cz/ap/download.php?id=178
http://yarchive.net/mil/laminar_flow.html
http://search.informit.com.au/documentSummary;dn=384807529637512;res=IELENG

Updated book collection

Fundamentals of Aerodynamics, by John Anderson Jr.
Aircraft Performance & Design, by John Anderson Jr.
Aircraft Design: Conceptual Approach, by Daniel Raymer
Jan Roskam: Aircraft Design parts 1-7
Jan Roskam: Airplane Flight Dynamics and Automated Flight Controls
Jan Roskam: Airplane Aerodynamics and Performance
Aerodynamics for Engineering Students
MODERN AIRCRAFT DESIGN, Volume 1 5th Edition, by Martin Hollmann.
MODERN AIRCRAFT DESIGN, Volume 2 4th Edition, by Martin Hollmann.
COMPOSITE AIRCRAFT DESIGN. REVISED 2003. By Dr. Hal Loken and Martin Hollmann.
MODERN AIRCRAFT DRAFTING by Eric and Martin Hollmann.
ADVANCED AIRCRAFT DESIGN by Martin Hollmann.
BRUCE CARMICHAEL'S PERSONAL AIRCRAFT DRAG REDUCTION
Theory of Flight
Aerodynamics for Engineers
Model aircraft aerodynamics
Smith: Illustrated guide to aerodynamics
Ron Wanttaja: Kit airplane construction
Bingelis: Sportplane construction techniques
Performance of Light Aircraft
Synthesis of Subsonic Aircraft Design
Theoretical Aerodynamics
Hoerner: Fluid Dynamic Drag
Flight Performance of Aircraft
Design of the Airplane
Burt Rutan: Moldless composite sandwitch aircraft consrtuction

Length diameter ratio of laminar pods of variable length and wing-body intersection optimization

The length-diameter ratio 3.33 was found ideal for laminar pods which are intended to the fuselage where the length Reynolds number tends to get high. The laminar flow can not sustained for very high length Reynolds number, therefore the need of relatively short pod when compared to a wing airfoil shape. That sounds like a rule of thumb, in other words, a generalization that applies to one example, but is not necessarily applicable to everything.

However, in case of engine pods, it would require some investigation to determine the optimum length/diameter ratio. On the wings, the length Reynolds number for a laminar engine pod would be similar than that of the wing. Logic says that if the wing can sustain 60% laminar flow with its chord length, then the pod with similar length diameter ratio should be able to do that as well.

Therefore, what is the ideal length diameter ratio for a engine pod if the engine pod comprises of NACA 66-series laminar symmetrical airfoil (which provides zero lift at zero degrees angle of attack)? Is it still 3.33 or something else?

I was yesterday evening also reading some documents I have got links from a Internet friend of mine (a aerodynamics-guru) and was comparing that to what was told in Bruce H. Carmichael's Personal Aircraft Drag Reduction Book. The fuselage-wing intersection optimization is described as a rule of thumb in the book, with the premise that the designer does not have access to CFD software, optimizing the streamlines of the fuselage to be similar than the streamlines of the wing, to avoid adverse pressure gradient.

However, today the CFD software does not need very expensive, in fact, OpenFoam is free software, and the situation might prove nowadays different than it used to be (still haven't had enough time to learn how to use the OpenFoam, but I will find out sooner or later, because I must). It would be enlightening to try out the wing-body intersection optimization. One thing I also learned is that the fairing between the wing and body has to be turbulent airfoil which has very late separation, because the flow at the wing intersection on the fuselage is turbulent anyway, the laminar flow can not be sustained that far without active boundary layer control. I am not planning active boundary layer control for step 1, to get things done.

Length diameter ratio for laminar pods

Laminar pods need to have low length/diameter ratio to get the benefits of laminar flow. Bruce Carmichael recommends length/diameter ratio of 3.33 in his book. I accidentally found also a pdf format article from web which talks about the same thing. You can get if from here:

www.aerorag.com/resource/aircraft/aerodynamics/carmichael/min_fus_drag_carmichael.pdf

Here is another document about the matter:
ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19860014381_1986014381.pdf

Grizzly

I was one day looking for information about Rutan's Grizzly, a three surface STOL bush plane which doesn't look at all like the Piper Cub. Today, I found a related patent, how Burt Rutan managed to implement fowler flaps without external supports which create drag on cruise.

You can read it here: http://www.freepatentsonline.com/4614320.html?query=PN%2F4614320+OR+4614320&stemming=on

60000 feet with Rotax 912, 80 hp

Here is an article which includes some text about Burt Rutan's Raptor UAV.
www.flightglobal.com/pdfarchive/view/1993/1993%20-%202623.html

Just accidentally when searching about Raptor UAV (this is off-topic to this posting, but anyhow contains interesting information including patent numbers), I found this: Burt Rutan's CV. Needless to say "Burt Rutan is my hero", but here is the CV of Mr. Rutan:

http://www.roycecarlton.com/speaker/Burt-Rutan-Curriculum-Vitae/

Why Cirrus is limited to 17500 feet?

I was thinking about over 25000 feet cruise altitude for non-pressurized version of my concept, but I was yesterday Googling about death zone and effects of high altitude to human physiology, and it became quite apparent that it is not healthy to fly at 25000-30000 feet, it is too high altitude for humans to bear even with supplemental oxygen. Even with pressure masks like those on fighter pilots, it might not be very comfortable and safe. It is therefore not a surprise after all, why some non-pressurized GA planes are limited to 17500 feet (like Cirrus SR20 and SR22).

So the need for pressurization comes a lot earlier than I was thinking, and apparently even cruising over 20000 feet would pretty much require it.

Some articles about supplemental oxygen use:
http://www.dr-amy.com/rich/oxygen/

The highest altitude non-pressurized aircraft have been certified usually are 25000 feet according to quick searches to Internet. Columbia 400 (Cessna 400) is non-pressurized and certified to 25000 feet. Flight at that altitude require oxygen mask and it is just above the "death zone" which was mentioned in one Mt. Everest page I was looking yesterday.

According to one UAV report I have (SR22 was compared to a UAV airframe), Cirrus SR22 technical service ceiling is at about 33000 feet. SR20 on the other hand with a lot less excess power does not most likely reach its limit altitude of 17500 feet most likely unless it is very lightly loaded. On our trip to Mojave it barely made it to 11000 feet at gross weight and non-standard atmospheric temperature conditions (it was hotter than on standard atmosphere).

What is important for getting desired performance out of an airframe

I have been looking quite a while how to get the aerodynamic design optimal and how to save there some drag, or a lot of drag, but a good design has also other parts taken into consideration. One of them which should not be underestimated is the structural and thus weight.

If we look for example EM-11 Orka, what is the problem with it when it is actually slower than aerodynamically less efficient and lower power Tecnam P2006T. It is pretty obvious what is the problem: it is not the aerodynamics of the plane (which is good) but the weight. The gross weight of Orka is very high, even higher than on DA42 that some people consider to be a lead-angel (lyijyenkeli). This has implications obviously to the empty weight too. That is very high as well. The empty weight-gross weight ratio is not actually bad in Orka, it is actually better than average. However, because of the gross weight being so high, the empty weight has to follow too. With the high weight, aerodynamic efficiency goes out of the door.

So it is very important that aircraft has minimum possible empty weight and as high as possible empty weight to gross weight ratio.

From the lighter end of the scale, Dynaero MCR01 is a good example. It is very lightweight, a lot lighter than its competitors. And it really shows positively in the performance. The wings in the ULC-model don't even incorporate a NLF-airfoil and the fuselage is all-turbulent behind the propeller. Still it is damn fast compared to all competition in its class with the same engine and propeller. The Dynaero's empty weight-cross weight ratio is not actually much better than on Orka, but because Orka is so much heavier and it is designed to carry so much more, the end result is very heavy (and it requires higher power engines than the Orka prototype originally had).

So this leads to a conclusion:
Previously mentioned gross weight of 818 kg for the twin concept is not unfounded. It represents ratio of 0.55 which is worse than on Orka or Dynaero MCR01. The goal has to be drawn somewhere. If the empty weight has to be more, e.g. 500 kg, that means 900 kg MTOW with ratio 0.55, and already a bit worse performance (speed (because the plane has to fly at higher Cl to maintain level flight on cruise and it is no good especially if the airfoil was designed to give its lowest drag at low Cl value) and climb performance).

Someone might be wondering why I don't talk about aerobatics much at all - Aerobatic planes require higher empty weight - gross weight ratios more than 0.55, and because of that I am not even thinking about a aerobatic plane which is intended for cross country flying. Efficient cross country machine has to be separate from aerobatic plane unfortunately because of restrictions what is achievable with even the best materials out there. Strength in airplane is not a place where a compromise can be made, it must be strong enough for the intended use or it is a deathtrap, and this leads to that the empty weight - gross weight ratio may not go much lower than 0.53 very easily on a small aircraft, especially without compromising something else like aerodynamics.

6 milestones plan for getting things done

I have been thinking the ways to achieve a design and implementation of a dream aircraft, and have concluded that it has to go in more than one step, so I was thinking the following milestones:

1. Unpressurized version, with a single turbo and fuel injection kit per engine. Possibly with a cabin similar to seen in Orka, avoid the manufacture of the doors. Woodcomp CS propellers. Target cruise altitude = 25000-30000 ft with supplemental oxygen. Corners cut where necessary to just get it done. No active boundary layer control, no wing tip propellers etc., rely on natural laminar flow to achieve efficiency. Unstable release of plans, calculations etc. Version A.
2. Open source plans stable release for the version A (CNC code, 3D models, 2D drawings, construction plans, layup schedules). Flight testing gives the final specifications for version B and ideas what to change to version B. Version A prototype is in use.
3. Optimized version of the above, version B. Modifications to version A prototype, version A becomes version B.
4. Stable release of version B plans (CNC code, 3D models, 2D drawings, construction plans, layup schedules). Version B might be alternative for a basis of a kit.
5. Pressurized version with doors, twin turbos per engine, intercooler and aftercooler per engine, computer controlled waste gates, and hybrid turbo compounding with two electric motors where one is functioning as generator and the the other runs the compounding. Possibly longer wings for high altitude flight. MT propeller or other higher end propellers. Possibly aerodynamic design changes, based on issues found in versions A and B and other improvements. Version C.
6. Open source plans stable release for the version C (CNC code, 3D models, 2D drawings, construction plans, layup schedules). Version C is a completely new aircraft and thus version B and version C coexists.

There are at least two milestones before 1.
-1 = concepting and collecting information, and creating needed softwares (present)
0 = initial concepting freezes, and version control repository (e.g. svn) exists for all data and there is a web page for the project.

Aeronautics Thesis Works from Finland

Couple of Finnish papers with aeronautics related topic

Fun factor for twin concept

I have been flying all kinds of planes and been kind of figuring slowly out what is the optimum for power loading. It turns out like 9 lbs/hp produces the "fun" experience. That is the "RV-grin" I would say.

So what comes together is:
- Optimum aircraft would consist of 2 x 100 hp engine
- Very low drag fuselage
- Very low drag wings
- High aspect ratio
- High wing loading, 22 lbs/sqft.
- Double slotted flaps
- Power loading 9 lbs/hp
-> mtow 1800 lbs = 818 kg
Empty weight should be under 450 kg to have enough useful load (368 kg, includes fuel).
=> wing area = 81 sqft.

For more general purpose use, it could be written:
- for high performance use, mtow limited to 818 kg.
- for long range use, mtow limited to 950 kg.

This becomes:
- the wing loading limit of 24 lbs/sqft can not be exceeded for the 950 kg because otherwise the stall speed gets too high
=> this becomes:
- 2090 lbs / 24 lbs/sqft
The wing area can be then assumed to be 87 sqft. 7 sqft more than on the case of high performance case.
- Wing loading calculation for the high performance case becomes:
87*22 = 1914 lbs MTOW.
1800/87.0 = 20.6 lbs / sqft

This would cause the airframe to gross weight ratio to be 0.47. This is very low and may not be realistic without special structure. A more realistic figure would be 0.55 ratio. This becomes: 450.0/0.55. Guess what, we get the 818 kg = 1800 lbs gross weight from that. So structurally the 450 kg empty weight and 818 kg gross weight should be feasible. Dynaero MCR-01 is 0.53; 260 kg / 490 kg = 0.53). The LH-Aviation LH10 is 260 kg/500 kg = 0.52. Both of these are carbon fiber structures. With lower cost materials, this may not be even nearly feasible.

If we take a pessimistic value for airframe to gross weight ratio - 0.6 and we have set the gross weight to 830 kg (based on optimizing the power loading), this gives 498 kg empty weight. This should be easily feasible if turbos and pressurization is not taken into account.

Just flew Dynaero MCR-01

A flying club friend (Samuli Pänttäjä) kindly offered a familiarization flight on his Dynaero MCR01. I flew with Pertti Husa (a flight instructor and friend).

The short story is that the plane is very interesting, it is very different from any other same category plane.

It is close to the maximum performance one can get out of Rotax 912 in tractor configuration without utilizing laminar flow over the fuselage (I don't mean only speed, but overall performance) - the climb rate, takeoff distance, climb speed, minimum speed, stall behavior and cruise speed at low altitude (IAS) and landing distance. This plane really rocks, it surely blows average Cessna-pilot away. Despite of the low horse power in the engine, this is maybe even more high performance aircraft than turbo Cirrus SR22 is with over 300 hp. This plane has 100 hp Rotax 912 with MT propeller hydraulic constant speed propeller. With Rotax 914 this...

The takeoff is very similar than on Cirrus SR22. Everything happens maybe even faster than with the Cirrus. The plane accelerates like a rocket, is airborne almost at the same moment, time to switch flap ups, trim the plane, reduce power and propeller speed all come very quickly.

The economy cruise speed (manifold pressure at 26, rpm at 4600) settled to about 250 km/h (135 kts IAS). We didn't try flying at altitude, I don't yet know how much TAS the plane collects at high altitude. At low altitude the cruise speed is anyhow about the same as on Cirrus SR20 leaned to best power setting. It really moves compared to Cessnas etc.

It also became apparent that the plane would cruise, with little more power, a lot faster. With a little pitch down causes the IAS to go over 300 km/h and it happens effortlessly and quickly. Watch out when pitching down or you will go over the VNe very quickly!

The plane takes of and lands to a very short distance. The approach speed is very low. The double slotted flaps are very effective and the plane can be flown insanely slowly. We did one approach at 80 km/h. On the other hand, in take off, the after the plane gets airborne and out of ground effect, the speed very quickly rises to 170 km/h (91 kts). Very comparable to Cirrus SR20. The big difference to Cirrus is that, on Dynaero, the climb angle is steep. It is going up like an elevator. Takeoff from very short runway is possible and it finely clears the obstacle with ease.

Feelings on landing pattern are quite similar than on SR22, one has to act quickly and not fall behind the aircraft. Pitch down, even on landing pattern, easily makes to plane go 300 km/h. If you are trying to be behind a Cessna that flies the pattern about 130 km/h, you are going to take over it, and very fast.

The "secret" of the plane is:
- very low empty weight
- very low cross sectional area
- small wetted area
- low cooling drag
- double slotted flaps (high Clmax)
- relatively high wing loading

Everything in the plane is made out of carbon fiber. Even rudder pedals are carbon fiber.

It is beneficial to have as low as possible empty weight, high Clmax, high wing loading and as great as possible power to weight ratio. This plane has those in better balance than other types I have flown to the date.

Some pictures:

Video of landing to EFHF at Youtube:

http://www.youtube.com/watch?v=SIojuZsGfUo