Aircraft Minimum Fuel Requirements

Aircraft Minimum Fuel Requirements

Commercial Flights

Per ICAO Annex 6, Part I, section 4.3.6 “Fuel Requirements,” airplanes should calculate their required fuel quantity as follows (summary; see below for actual ICAO text):

  • Taxi fuel
  • Trip fuel (to reach intended destination)
  • Contingency fuel (higher of 5% of “trip fuel” or 5 minutes of holding flight)
  • Destination alternate fuel (to fly a missed and reach an alternate)
  • Final reserve fuel (45 minutes of holding flight for reciprocating engines, 30 minutes for jets)
  • Additional fuel (if needed to guarantee ability to reach an alternate with an engine failure or at lower altitude due to a pressurization loss)
  • Discretionary fuel (if the pilot in command wants it)

General Aviation

For general aviation, ICAO Annex 6 Part II, section 2.2.3.6 “Fuel and oil supply” requires:

  • For IFR, enough fuel to reach destination, then alternate (if required), plus 45 minutes
  • For day VFR, enough fuel to reach destination plus 30 minutes
  • For night VFR, enough fuel to reach destination plus 45 minutes

ICAO Annex 6 Part I

From the ninth edition:

4.3.6.3 The pre-flight calculation of usable fuel required shall include:

a) taxi fuel, which shall be the amount of fuel expected to be consumed before take-off;

b) trip fuel, which shall be the amount of fuel required to enable the aeroplane to fly from take-off, or the point of in-flight re-planning, until landing at the destination aerodrome taking into account the operating conditions of 4.3.6.2 b);

c) contingency fuel, which shall be the amount of fuel required to compensate for unforeseen factors. It shall be five per cent of the planned trip fuel or of the fuel required from the point of in-flight re-planning based on the consumption rate used to plan the trip fuel but, in any case, shall not be lower than the amount required to fly for five minutes at holding speed at 450 m (1 500 ft) above the destination aerodrome in standard conditions;

Note.— Unforeseen factors are those which could have an influence on the fuel consumption to the destination aerodrome, such as deviations of an individual aeroplane from the expected fuel consumption data, deviations from forecast meteorological conditions, extended taxi times before take-off, and deviations from planned routings and/or cruising levels.

d) destination alternate fuel, which shall be:

  1. where a destination alternate aerodrome is required, the amount of fuel required to enable the aeroplane to:
    i) perform a missed approach at the destination aerodrome;
    ii) climb to the expected cruising altitude; iii) fly the expected routing;
    iv) descend to the point where the expected approach is initiated; and
    v) conduct the approach and landing at the destination alternate aerodrome; or
  2. where two destination alternate aerodromes are required, the amount of fuel, as calculated in 4.3.6.3 d) 1), required to enable the aeroplane to proceed to the destination alternate aerodrome which requires the greater amount of alternate fuel; or
  3. where a flight is operated without a destination alternate aerodrome, the amount of fuel required to enable the aeroplane to fly for 15 minutes at holding speed at 450 m (1 500 ft) above destination aerodrome elevation in standard conditions; or
  4. where the aerodrome of intended landing is an isolated aerodrome:
    i) for a reciprocating engine aeroplane, the amount of fuel required to fly for 45 minutes plus 15 per cent of the flight time planned to be spent at cruising level, including final reserve fuel, or two hours, whichever is less; or
    ii) for a turbine-engined aeroplane, the amount of fuel required to fly for two hours at normal cruise consumption above the destination aerodrome, including final reserve fuel;

e) final reserve fuel, which shall be the amount of fuel calculated using the estimated mass on arrival at the destination alternate aerodrome, or the destination aerodrome when no destination alternate aerodrome is required:

  1. for a reciprocating engine aeroplane, the amount of fuel required to fly for 45 minutes, under speed and altitude conditions specified by the State of the Operator; or
  2. for a turbine-engined aeroplane, the amount of fuel required to fly for 30 minutes at holding speed at 450 m (1 500 ft) above aerodrome elevation in standard conditions;

f) additional fuel, which shall be the supplementary amount of fuel required if the minimum fuel calculated in accordance with 4.3.6.3 b), c), d) and e) is not sufficient to:

  1. allow the aeroplane to descend as necessary and proceed to an alternate aerodrome in the event of engine failure or loss of pressurization, whichever requires the greater amount of fuel based on the assumption that such a failure occurs at the most critical point along the route;
    i) fly for 15 minutes at holding speed at 450 m (1 500 ft) above aerodrome elevation in standard conditions; and
    ii) make an approach and landing;
  2. allow an aeroplane engaged in EDTO to comply with the EDTO critical fuel scenario as established by the State of the Operator;
  3. meet additional requirements not covered above;

Note 1.— Fuel planning for a failure that occurs at the most critical point along a route (4.3.6.3 f) 1)) may place the aeroplane in a fuel emergency situation based on 4.3.7.2.

Note 2.— Guidance on EDTO critical fuel scenarios is contained in Attachment D;

g) discretionary fuel, which shall be the extra amount of fuel to be carried at the discretion of the pilot-in-command.

ICAO Annex 6 Part II

2.2.3.6 Fuel and oil supply

A flight shall not be commenced unless, taking into account both the meteorological conditions and any delays that are expected in flight, the aeroplane carries sufficient fuel and oil to ensure that it can safely complete the flight. The amount of fuel to be carried must permit:

a) when the flight is conducted in accordance with the instrument flight rules and a destination alternate aerodrome is not required in accordance with 2.2.3.5, flight to the aerodrome of intended landing, and after that, for at least 45 minutes at normal cruising altitude; or

b) when the flight is conducted in accordance with the instrument flight rules and a destination alternate aerodrome is required, flight from the aerodrome of intended landing to an alternate aerodrome, and after that, for at least 45 minutes at normal cruising altitude; or

c) when the flight is conducted in accordance with the visual flight rules by day, flight to the aerodrome of intended landing, and after that, for at least 30 minutes at normal cruising altitude; or

d) when the flight is conducted in accordance with the visual flight rules by night, flight to the aerodrome of intended landing and thereafter for at least 45 minutes at normal cruising altitude.

Fuel & oil (propeller –driven A/c)

A) When a destination alternate is required

B) When destination aerodrome is VMC /alternate not required

C) When a destination is isolated and no suitable destination alternate is available

i. Fuel to destination ………. A (includes one approach and a missed approach)

ii. + fuel is specified diversion ……….B

iii. + 45 mts i.e A+B+45 mts

i. Fuel to destination …. A

ii. +45 mts i.e A+45 mts

i. Fuel to destination …. A

ii. + 45 mts+15 5 of cruising time or 2 hrs ….B

which is less

i.e A+B

Fuel & oil (Turbo jets)

A. When a destination alternate is required

Fuel destination including an approach and a missed approach …A

Fuel to alternate aerodrome as specified in flights plan …B

  • 30 mts holding at 1500’ at holding speed above alternate aerodrome in ISA
  • Approach
  • land

Additional contingency fuel as specified by operator …C

i.e A+B + 30 mts holding at 1500’ +approach + landing fuel +C

B. When a destination alternate is not required because destination is VMC

Fuel to destination + 30 mts holding fuel at 1500’ above the aerodrome destination at holding speed in ISA …. B

Additional contingency fuel as specified by operator …C

i.e A+B+C

C. When destination is isolated and no suitable destination alternate aerodrome is available

Fuel to destination + 2 hrs fuel at normal fuel cruise consumption …. B

i.e A+B

Hadley Cells About Atmosphere

Hadley Cells About Atmosphere

The atmosphere transports heat throughout the globe extremely well, but present-day atmospheric characteristics prevent heat from being carried directly from the equator to the poles. Currently, there are three distinct wind cells – Hadley Cells, Ferrel Cells, and Polar Cells – that divide the troposphere into regions of essentially closed wind circulations. In this arrangement, heat from the equator generally sinks around 30° latitude where the Hadley Cells end. As a result, the warmest air does not reach the poles. If atmospheric dynamics were different, however, it is plausible that one large overturning circulation per hemisphere could exist and that wind from the low-latitudes could transport heat to the high-latitudes. As an explanation for equable climates, Brian Farrell presented this idea in 1990 and advocated that during equable climates, the Hadley Cells extended from the equator to the poles (Farrell, 1990).

Atmospheric Wind Cells

Hadley cells, Ferrel (mid-latitude) cells, and Polar cells characterize current atmospheric dynamics.

Hadley Cells are the low-latitude overturning circulations that have air rising at the equator and air sinking at roughly 30° latitude. They are responsible for the trade winds in the Tropics and control low-latitude weather patterns. Held and Hou (1980) outlined the dynamics of this circulation through a simplified model of the Hadley Cell. For the model, there are three main assumptions. First, the Hadley Cell circulation is constant. Second, the air moving toward the poles in the upper atmosphere conserves its axial angular momentum, while the surface air moving equatorwards is slowed down by friction. Third, the thermal wind balance holds for the circulation (Vallis, 2006). For simplicity, the model is also symmetric around the equator. These initial assumptions make the explanation of Hadley Cell dynamics much simpler.

Angular momentum is defined as the cross product of a particle’s distance from the axis of rotation, r, and the particle’s linear momentum, p. In the Hadley Cell as an air particle moves toward the high-latitudes, it becomes closer to the Earth’s spin axis, so r becomes smaller. If angular momentum is conserved in the Hadley Cell as Held and Hou (1980) assume, p must become larger to balance the decrease in r. P equals mass times velocity. Since the mass of the air particle cannot change, the velocity of the particle must increase. In the case of the Hadley Cell, the velocity in question is the zonal (east-west) velocity, so as the particle moves poleward, the velocity must increase in the eastward direction. Eventually, the zonal velocity is so strong that the particle stops moving poleward and only travels to the east. At this latitude, air sinks, and then to close the loop, it returns to the equator along the surface. Therefore, because of the conservation of angular momentum, Hadley Cells exist only from the equator to the mid-latitudes.

Angular Momentum

To conserve angular momentum, velocity must increase as the radius decreases. (Image courtesy of Lyndon State College Atmospheric Sciences)

This scenario holds as long as the initial assumptions are valid. Brian Farrell, however, argues that the assumptions are not accurate for equable climates and that during equable climates, angular momentum is not conserved in poleward moving particles (1990). He claims that angular momentum sinks, essentially sources of friction, could have been stronger during the Eocene and the Cretaceous. Farrell estimates that the friction term in his model would increase by eightfold under equable climate conditions. This change would prevent angular momentum from being conserved. In this situation, the zonal velocity would not become strong enough to stop air from moving poleward. Instead, air from the equator would be able to travel all the way to the poles in extended Hadley Cells.

Based off of Venus’ atmosphere’s behavior, Farrell argues that another way to extend the Hadley Cells would be to increase the height of the tropopause. This change would increase the poleward moving air’s Rossby number. The Rossby number describes the importance of the Coriolis force in atmospheric dynamics. A higher Rossby number means that the Coriolis force has a smaller impact on a particle, so if the height of the tropopause increased enough, the Rossby number would become high enough to make the Coriolis force negligible. As a result, particles would not diverge from their path as they moved poleward, and the Hadley Cells would reach the poles. To explain how the tropopause height could increase, Farrell states that the height is correlated to surface temperature and that a 1°C increase in sea surface temperature causes the tropopause potential temperature to rise by roughly 7.5°C. Raising the average equatorial sea surface temperature to 32°C from its current 27°C would increase the potential temperature of the tropopause by 37°C. There temperature increases would almost double the static stability at the tropopause. For the height to increase, the stratosphere would also have to become less stable. If CO2 concentrations increased and if stratospheric ozone concentrations decreased, the stratosphere would cool substantially, and this change would destabilize the stratosphere. As a result of the alterations to tropospheric and stratospheric stability, the tropopause height would increase. Farrell estimates the height would have doubled under Cretaceous conditions, and as a result, the Rossby number would have doubled. This change would have allowed the Hadley Cells to extend to the poles and would have made equable climates more likely.

Hadley cells extended all 
	the way to the poles

Hadley cells could extend all the way to the poles. (Image courtesy of Lyndon State College Atmospheric Sciences).

While each of these alterations to the atmosphere would extend the Hadley Cells, Farrell found that a combination of the two effects was necessary to make his model’s results agree with proxy data from equable climates. He graphed the atmosphere’s potential temperature versus latitude at different tropopause height and friction values. The results reveal that as tropopause height and friction increase, the EPTD decreases. A doubling of the tropopause height combined with an eightfold increase in the friction term leads to an EPTD of 16°C. This value agrees with Cretaceous climate reconstructions. As a result, Farrell’s theory seems to be a reasonable explanation for equable climates.

Potential Temperature vs. Latitude at 
	Tropopause Height = 15 km
Potential Temperature vs. Latitude at 
	Tropopause Height = 30 km

Farrell’s results show that as the friction term (Γ) and the tropopause height increase, the EPTD decreases. (Farrell, 1990)

The main problem is that Farrell does not provide any explanation for why angular momentum sinks would have become stronger during the Cretaceous and the Eocene. He provides a few examples of potential momentum sinks: “small scale diffusion…, cumulus momentum flux…, gravity wave drag…, and the net westward force arising from potential vorticity mixing by large scale waves” (Farrell, 1990). However, he does not explain why any of these sinks would have become stronger during the Eocene and, thus, would have prevented angular momentum from being conserved. This lack of information in the argument makes the theory harder to accept, and until this portion of the argument is explored in greater depth, Farrell’s theory cannot be accepted as the correct explanation of equable climates.

Farrell’s theory about the extension of the Hadley Cells has a mathematical basis, and he argues his case through the derivation of atmospheric dynamics equations. He starts with an overview of the work done in Held and Hou (1980) and then modifies the equations to account for his ideas. In a similar manner, the following information will walk through the work done by Held and Hou and then will present the modified equations derived by Farrell. These steps follow sections 11.2 and 11.3 in Atmospheric and Oceanic Fluid Dynamics by Geoffrey K. Vallis (2006).

For the model of the Hadley circulation, the atmosphere is assumed to be a Boussinesq atmosphere. As a result, the zonally averaged momentum equation in the absence of friction is

Zonally Averaged Zonal Momentum (1)

The overbars represent zonal averages. If vertical advection is considered to be small in comparison to horizontal advection and if we ignore the eddy terms on the right-hand side of the equation, then a steady state solution is

Steady State Equation (2)

If we assume that meridional flow is not zero, then f + ζ = 0. Another way of writing this equation is

Modified Steady State Equation (3)

where the left side equals f and the right side equals ζ. We assume that the zonal flow at the equator is zero because air rises from the surface there where the flow is weak by assumption. A solution to equation 3 is then

Zonal Velocity Equation (4)

This equation provides the zonal velocity of a particle moving toward the poles in the upper part of the circulation. We can also derive this equation from the conservation of angular momentum, m, of an air parcel at certain latitude ϑ. Angular momentum can be represented as

Angular Momentum
		Equation (5)

If zonal velocity equals zero at the equator and if a polewards moving air parcel conserves momentum, equation 5 leads directly to equation 4. Also taking the derivative of equation 5 with respect to latitude reveals

Derivate of 
		the Angular Momentum Equation (6)

These steps demonstrate that if friction and eddy fluxes can be ignored, then air moving towards the pole in the Hadley Cell will conserve its angular momentum. This fact means that an air parcel moving polewards must accelerate zonally as it moves toward the high-latitudes, so the zonal velocity will increase with latitude.

Since the zonal velocity is assumed to be low near the surface and since equation 4 provides the zonal velocity in the upper branch of the Hadley Cell, an estimate for the velocity gradient over the height of the cell is known. This fact means that we can use the thermal wind equation to find the vertically averaged temperature. While the geostrophic wind relation does not hold at the equator, it is accurate until very close to the equator. Therefore, the traditional thermal wind equation works for the model. This equation is

Thermal Wind Equation (7)

“where b = g δθ/θ0 is the buoyancy and δθ is the deviation of potential temperature from a constant reference value θ0” (Vallis 460). Substituting in equation 4 for u and vertically integrating the equation from the ground to a height H provides

Integrated 
		Thermal Wind Equation (8)

where θ = H-1* ∫H0δθ dz is the vertically averaged potential temperature. Assuming that the latitudinal extent of the Hadley Cell is not too great, we can use the small angle approximation and can replace sin ϑ with ϑ and cos ϑ with 1. After then integrating equation 8, it becomes

Potential 
		Temperature Equation (9)

where y = aϑ and θ(0) is the potential temperature at the equator. The value of this term is still unknown at this point, so we must use thermodynamics to obtain it. For this model, we will assume “that the thermodynamic forcing can be represented by a Newtonian cooling to some specified radiative equilibrium temperature, θE” even though this assumption is a big simplification (Vallis 460). The thermodynamic equation we shall use is

Thermodynamic 
		Equation (10)

where τ is a relaxation time scale. If we assume that the radiative equilibrium temperature falls from the equator to the pole and that it increases with height, a simple equation for it is

Radiative Equilibrium 
		Temperature Equation (11)

“where ΔH and ΔV are non-dimensional constants that determine the fractional temperature difference between the equator and the pole and [between] the ground and the top of the fluid, respectively” (Vallis 460). P2 is the second Legendre polynomial and is often the leading term in the Taylor expansion of symmetric functions around the equator. P2(y) = (3y2 – 1)/2. With the small angle approximation and at z = H/2, equation 11 becomes

Modified Radiative Equilibrium
		Equation (12)

or

Modified Radiative Equilibrium
		Equation 2 (13)

“where θE0 is the equilibrium temperature at the equator, Δθ determines the equator-pole-radiative equilibrium temperature difference, and” (Vallis 461)

Variables (14)

If we assume that equation 9 is valid between the equator and some latitude ϑH where the meridional velocity is zero, we create a closed system between the equator and this latitude. To conserve potential temperature, the solution to equation 9 must uphold

Constraint 1 (15)

where YH = aϑH. This requirement creates a closed system in which heat flows from the equator to ϑH to balance the atmosphere’s natural heating and cooling by absorption and emission of heat. A second constraint is that the solution must be continuous at y = YH, meaning

Constraint 2 (16)

Plugging equations 9 and and 13 into equations 15 and 16 yield

Y<sub>H</sub> (17)

and

Potential Temperature at the Equator (18)

These final equations reveal that the Hadley Cell should have a finite meridional extent. Held and Hou reached this conclusion for an inviscid atmosphere (1980), and their work agrees with the dynamics seen in the real atmosphere. Farrell uses this work as a starting point for his idea, but he modifies the equations to include friction. Instead of fully walking through all of his steps, which are similar to those of Held and Hou, the final equations that Farrell reaches are

Farrell's Final Equations

In these equations, R stands for the Rossby wave number, and Γ represents “a measure of the relative dominance of the radiation and momentum time scales” and can be thought of as a friction term (Farrell, 1990). R = gHϑ/ω2 a2, and Γ = SτRhτ v. These equations have a solution with a Hadley Cell beginning at the equator where Vb(0) = 0 and ending at the poles where Vb (yH) = 0. R and Γ determine this solution and, thus, are important factors for the Hadley Cell circulation.

As mentioned before, the height of the tropopause has a significant impact on the extent of the Hadley Cells. Looking at Farrell’s equations, one can now see how this fact is true. The height, H, is influential in determining the value of the Rossby wave number. If the height doubles, the Rossby wave number will also double. Using the first of Farrell’s equations, one can see that an increase in the Rossby wave number will decrease the change in potential temperature over latitude. This change will cause the temperature gradient to decrease throughout the extent of the Hadley Cell. Since the Hadley Cells extend as the Rossby wave number increases, the Hadley Cells will extend to the poles if R increases enough, and thus, the EPTD will decrease significantly.

Similarly, Γ plays an important role in atmospheric dynamics. It is important to note that the amount by which a torque will decrease angular momentum depends on the mass flux Vb, which is determined by SτR, which is part of Γ. In an inviscid atmosphere, Γ is set to zero, but in an atmosphere with friction, it must have a non-zero value. Since Γ determines how much a torque will decrease angular momentum, the zonal velocity decreases as Γ increases. Therefore, if Γ increases substantially, a large enough torque could be generated to prevent the formation of a zonal wind strong enough to stop an air parcel from moving poleward. As a result, a large Γ value could enable the Hadley Cells to extend all the way to the poles. This change would allow warm air from the equator to reach the high-latitudes and would reduce the EPTD to levels seen during equable climates.

Therefore, one can see that R and Γ are highly influential on the extent of the Hadley Cells. If both parameters increased enough during the Cretaceous and the Eocene, the Hadley Cells could have extended all the way to the poles. Under these conditions, air from the equator would have traveled all the way to the high-latitudes and would have heated the poles enough to have caused the equable climates.

Jet Blast

Jet Blast

Jet blast !

Jet blast is the phenomenon of rapid air movement produced by the jet engines of aircraft, particularly on or before takeoff.

A large jet-engined aircraft can produce winds of up to 100 knots (190 km/h; 120 mph) as far away as 60 metres (200 ft) behind it at 40% maximum rated power. Jet blast can be a hazard to people or other unsecured objects behind the aircraft, and is capable of flattening buildings and destroying vehicles

Despite the power and potentially destructive nature of jet blast, there are relatively few jet blast incidents. Due to the invisible nature of jet blast and the aerodynamic properties of light aircraft, light aircraft moving about airports are particularly vulnerable. Pilots of light aircraft frequently stay off to the side of the runway, rather than follow in the centre, to negate the effect of the blast.

Propeller planes are also capable of generating significant rearwards winds, known as prop wash.

Some airports have installed jet blast deflectors in areas where roads or people may be in the path of the jet blast on take off.

Jet blast deflector !

A jet blast deflector (JBD) or blast fence is a safety device that redirects the high energy exhaust from a jet engine to prevent damage and injury. The structure must be strong enough to withstand heat and high speed air streams as well as dust and debris carried by the turbulent air. Without a deflector, jet blast can be dangerous to people, equipment and other aircraft.

Jet blast deflectors range in complexity from stationary concrete, metal or fiberglass fences to heavy panels that are raised and lowered by hydraulic arms and actively cooled. Blast deflectors can be used as protection from helicopter and fixed-wing aircraft propwash. At airports and jet engine service centers, jet blast deflectors can be combined with sound-deadening walls to form a ground run-up enclosure within which a jet aircraft engine can safely and more quietly be tested at full thrust.

Purpose

High energy jet engine exhaust can cause injury and damage. Jet blast has been known to uproot trees, shatter windows, overturn automobiles and trucks, flatten poorly made structures and injure people. Other aircraft in the jet blast, especially lightweight ones, have been blown around and damaged by jet exhaust. Hurricane-force air streams moving at speeds up to 100 knots (190 km/h; 120 mph) have been measured behind the largest jet-powered aircraft at distances of over 200 feet (60 m). A Boeing 777’s two General Electric GE90 engines combine to create a thrust of approximately 200,000 pounds-force (900,000 N)—this level of force is high enough to kill people. To prevent these problems, jet blast deflectors redirect the air stream in a non-dangerous direction, frequently upward.

Airports

An illustration of a Christmas tree at Glasgow Air Force Base, showing the positioning of the jet blast deflectors

Jet blast deflectors began to appear at airports in the 1950s. Airports in the 1960s used jet blast deflectors with a height of 6 to 8 feet (1.8 to 2.4 m), but airports in the 1990s needed deflectors that were twice as high, and even up to 35 feet (11 m) high for jet airliners such as the McDonnell Douglas DC-10 and MD-11, which have engines mounted in the tail above the fuselage. Airports often place their deflectors at the beginnings of runways, especially when roadways or structures are adjacent. Airports that are in dense urban areas often have deflectors between taxiways and airport borders. Jet blast deflectors usually direct exhaust gases upward. However, a low-pressure zone can form behind the blast fence, causing ambient air and debris to be drawn upward with the jet exhaust, and hot, toxic gases to circulate behind the blast fence. Jet blast deflectors have been designed to counteract this problem by using multiple panels and various angles, and by using slotted panel surfaces.

Ground run-up enclosure

After a jet engine has been overhauled or has undergone the replacement of parts, it is normal to run the engine up to full thrust to test it.Rural airports rarely provide more than a distant portion of the airfield within which to test engines at full thrust, but urban airports surrounded by residential areas often specify that engine tests be conducted within a ground run-up enclosure (“hush house”), so that the engine noise can be reduced for residents.

Aircraft carriers

In 2003 aboard the USS Abraham Lincoln (CVN-72), a jet blast deflector is raised hydraulically to protect one F/A-18 Hornet from the jet exhaust of another.

Aircraft carriers use jet blast deflectors at the rear of aircraft catapults, positioned to protect other aircraft from exhaust blast damage. Jet blast deflectors are made of heavy duty material that is raised and lowered by hydraulic cylinders or linear actuators. The jet blast deflector lies flush with and serves as a portion of the flight deck until the aircraft to be launched rolls over it on the way to the catapult. When the aircraft is clear of the deflector, the heavy panel is raised into position to redirect the hot jet blast. As soon as the deflector is raised, another aircraft can be brought into position behind it, and flight deck personnel can perform final readiness duties without the danger of hot, violent exhaust gases. Such systems were installed on aircraft carriers in the late 1940s and early 1950s, as jet-powered aircraft began to appear in navies.

Jet blast deflectors aboard aircraft carriers are placed in very close proximity to the 2,300 °F (1,300 °C) temperatures of modern jet fighter exhaust. The non-skid decking surface of the deflector suffers heat damage and requires frequent maintenance or replacement. Additionally, the hot deflector surface cannot be used as normal decking until it has cooled enough to allow aircraft tires to roll over it. To mitigate the heat problem, active cooling systems were installed in the 1970s, tapping the fire mains (fire suppression water systems) to use seawater circulating through water lines within the deflector panel. However, the water cooling system adds more complexity and failure points, and requires additional maintenance. The most recent method tried by the United States Navy for solving the heat problem was introduced in 2008 with USS George H.W. Bush which uses heavy-duty metal panels covered in heat-dissipating ceramic tiles similar to those used on the Space Shuttle. The tiled panels can be changed quickly and easily—the ship carries a large replacement supply. Without active water lines, the passively-tiled deflector is expected to require much less maintenance.

25 Insanely Dangerous Airports Around The World

25 Insanely Dangerous Airports Around The World

25 Insanely Dangerous Airports Around The World

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I love flying. Seeing the world from the clouds and traveling to foreign places is always a magical experience. But like any good meal, there’s always a bill to pay at the end. And for flying, it sometimes comes with a treacherous landing. While it’s only just a bit jarring when flying into a “normal” airport, just wait until you see these death defying approaches. You may never fly again.

1. Toncontin International Airport, Honduras

Although Boeing 757s land and take off from this airport all the time, the History Channel rated it the second-most dangerous airport in the world due to the mountainous terrain.

2. Agatti Aerodrome, Lakshadweep, India

Constantly up for a potential extension (it’s only 4,000 feet long), planes have to risk the short runway to serve the local 36 Indian tourist islands.

3. Cleveland Hopkins International

Not terrible as an actual airport, but it’s notorious for having incompetent staff. This includes sleeping through landing protocols and being too engrossed in DVDs to help park planes at the gates. Fun times.

4. Gibraltar International Airport, Gibraltar

With a major road running through the middle, traffic has to stop for every plane that lands.

5. Kansai International Airport, Japan

Large enough to be viewed from space, it faces earthquakes, cyclones, and an unstable seabed.

6. Wellington International Airport, New Zealand

A truly hazardous and short runway, it ends with a death-defying fall straight into the ocean.

7. Savannah/Hilton Head International Airport

Not dangerous as much as it is creepy, embedded in Runway 10’s tarmac is a pair of grave markers belonging to the previous owners of the land upon which the airport now sits.

8. Qamdo Bamda Airport, Tibet

At 14,000 feet, the thin oxygen often plays extreme havoc with engines, leaving pilots and passengers scared out of their minds.

9. Princess Juliana International Airport, Saint Martin

This one pretty much speaks for itself. Wow.

10. Paro Airport, Bhutan

With only eight pilots qualified to land at this Himalayan airport, it’s no wonder passengers take anti-anxiety meds before landing.

11. Old Mariscal Sucre International Airport, Ecuador

Closed last year for being too dangerous, it was beset by mountainous terrain, active volcanoes, and a valley prone to fog.

12. Narsarsuaq Airport, Greenland

One of the most difficult approaches in the world, it requires flying up a fjord with severe turbulence and wind shear even on the calmest days of the year.

13. Lukla Airport, Nepal

This runway makes for a harrowing takeoff, with the end 2,900 meters up in the air.

14. Madeira Airport, Portugal

It’s pretty easy to see why it’s so dangerous with an elevated (and artificial) runway.

15. Los Angeles (LAX)

Los Angeles International has been ranked multiple times throughout the years as the most dangerous airport in the U.S.

16. Ice Runway, Antarctica

One of three major airstrips used to haul supplies and researchers to Antarctica’s McMurdo Station. This was developed to allow passenger Boeing jets to land instead of taking up room on cargo planes.

17. Kai Tak Airport, Hong Kong

Once called “the mother of all scary airports,” it was shut down in 1998 after pilots couldn’t stand it anymore.

18. Juancho E. Yrausquin Airport, Saba Island

Only fit for the smallest of planes, and even they often come VERY close to falling over the edge.

19. Gustaf III Airport, Saint Barthélemy

This runway is so small that planes can only carry a maximum of 20 people.

20. Gisborne Airport, New Zealand

This airport is unique. Unique because it has a regular freight train rolling through it every day!

21. Don Mueang International Airport, Thailand

Yes, that IS an 18 hole golf course in the middle of the runways.

22. Courchevel Altiport, France

Serving the uber rich who want to ski the Alps, this incredibly short runway has a ski mogul in the middle of it.

23. Dammam King Fahd International Airport, Saudi Arabia

King Fahd International is the largest airport in the world in terms of landmass, sprawling over three hundred square miles of desert.

24. Congonhas Airport, Brazil

Located just 5 miles from the center of downtown São Paulo. Apparently pilots need strong doses of alcohol after their flight to calm themselves down.

25. Barra International Airport, Scotland

Barra International Airport is the only airport on the planet where scheduled flights use a beach as the runway.
Refrence :

Aircraft Records

Aircraft Records

All Aircraft Records

This article gives yearly aviation records under 5 headings: airspeed, range, ceiling, gross take-off weight, and engine power.

Airspeed Records

Range Records

Ceiling Records

T/O Weight Records

Engine Power Records

 

 

Year Airspeed Range Ceiling T/O Weight Engine power
1905 60.91 km/h (37.85 mph)
USA
Wilbur Wright
Flyer III
October 5, 1905
38.95 km (24.2 miles)
USA
Wilbur Wright
Flyer III
October 5, 1905
15 m (50 ft)
USA
Wilbur Wright
Flyer III
September 28, 1905
388 kg (855 lb)
USA
Wright Brothers
Flyer III
37 kW (50 hp)
France
Léon Levavasseur
Antoinette
1907 25 m (82 ft)
France
Louis Blériot
Blériot VI
September 17, 1907
522 kg (1,151 lb)
France
Voisin Brothers
Voisin-Farman No 1
40 kW (54 hp)
France
Renault
VB Renault
1908 64.79 km/h (40.26 mph)
France
Henry Farman
Voisin-Farman No 1
October 30, 1908
124.69 km (77.48 miles)
USA
Wilbur Wright
Wright A
December 31, 1908
110 m (361 ft)
USA
Wilbur Wright
Wright A
December 18, 1908
544 kg (1,200 lb)
USA
Wilbur Wright
Wright A
59 kW (79 hp)
France
Gobron-Brille
Gobron
1909 76.96 km/h (47.82 mph)
France
Louis Blériot
Blériot XII
August 28, 1909
234.21 km (145.53 miles)
France
Henry Farman
HF.1 No III
December 1, 1909
453 m (1,486 ft)
France
Hubert Latham
Antoinette VII
December 1, 1909
620 kg (1,367 lb)
France
Louis Blériot
Blériot XII
66 kW (89 hp)
France
Gobron-Brille
Gobron
1910 110 km/h (68.2 mph)
France
Alfred Leblanc
Blériot XI
October 29, 1910
584.7 km (363.34 miles)
France
Maurice Tabuteau
Maurice-Farman
December 30, 1910
3,100 m (10,170 ft)
France
Georges Legagneux
Hubert Latham
Antoinette VII
December 1, 1909
1,322 kg (2,950 lb)
Great Britain
Samuel Cody
Cody Michelin Cup
132 kW (177 hp)
France
Clerget
Gobron
1911 141.2 km/h (87.73 mph)
France
Édouard Nieuport
Nieuport Nie-2 N
June 21, 1911
740 km (460 miles)
France
Armand Gobé
Nieuport
December 24, 1911
3,910 m (12,828 ft)
France
Roland Garros
Blériot XI
September 4, 1911
1,350 kg (2,976 lb)
France
Léon Levavasseur
Antoinette Monobloc
132 kW (177 hp)
France
Clerget
Double Clerget 4W
1912 174.1 km/h (108.18 mph)
France
Jules Védrines
Deperdussin Monocoque
September 9, 1912
1,010.89 km (628.14 miles)
France
Géo Fourny
Maurice Farman
September 11, 1912
5,610 m (18,405 ft)
France
Roland Garros
Morane-Saulnier
December 11, 1912
147 kW (197 hp)
France
Clerget
Clerget
1913 203.8 km/h (126.66 mph)
France
Maurice Prévost
Deperdussin Monocoque
September 29, 1913
1,021.19 km (634.54 miles)
France
A Seguin
Henry Farman
October 13, 1913
6,120 m (20,079 ft)
France
Georges Legagneux
Nieuport II N
December 28, 1913
4,080 kg (8,995 lb)
Russia
Igor Sikorsky
Russian Knight
162 kW (217 hp)
France
Salmson Canton-Unné
(CU) 2M7
1914 216.5 km/h (134.54 mph)
Great Britain
Norman Spratt
RAF SE.4
June 1914
1,900 km (1,180.61 miles)
Germany
Werner Landmann
Albatros
June 28, 1914
8,150 m (26,739 ft)
Germany
Heinrich Oelerich
DFW
July 14, 1914
4,800 kg (10,582 lb)
Russia
Igor Sikorsky
Ilya Muromets A
168 kW (225 hp)
Great Britain
Sunbeam
1915 6,350 kg (14,000 lb)
Great Britain
Handley Page Ltd
Handley Page 0/100
169 kW (227 hp)
France
Louis Renault
Renault 12A
1916 12,129 kg (26,739 lb)
Germany
Staaken
R.VI
296 kW (296 hp)
France
Louis Renault
Renault 12F
1917 12,955 kg (28,561 lb)
Germany
Staaken
R.VII
298 kW (400 hp)
USA
Packard & Hall Scott
Liberty
1918 262.4 km/h (163.06 mph)
USA
Roland Rohlfs
Curtiss Wasp
August 19, 1918
8,808 m (28,897 ft)
USA
Rudolph Schroeder
Bristol F.2B
November 18, 1918
15,900 kg (35,053 lb)
Germany
Staaken
R.XIVa
522 kW (700 hp)
Italy
Fiat
Fiat A.14
1919 307.5 km/h (191.1 mph)
France
Joseph Sadi-Lecointe
Nieuport-Delage 29v
December 16, 1919
3,032 km (1,884 miles)
Great Britain
Alcock and Brown
Vickers Vimy
June 15, 1919
10,549 m (34,610 ft)
USA
Roland Rohlfs
Curtiss Wasp
September 18, 1919
20,263 kg (44,672 lb)
Great Britain
WG Tarrant Ltd
Tarrant Tabor
1920 313 km/h (194.49 mph)
France
Joseph Sadi-Lecointe
Nieuport-Delage 29v
December 12, 1920
1921 330 km/h (205.22 mph)
France
Joseph Sadi-Lecointe
Nieuport-Delage
September 26, 1921
26,000 kg (57,319 lb)
Italy
SAI Caproni
Caproni Ca 60
625 kW (838 hp)
France
Marcel Riffard
Breguet-Bugatti 32A
1922 361 km/h (224.28 mph)
USA
William Mitchell
Curtiss R-6
October 18, 1922
4,052 km (2517.8 miles)
USA
Oakley G. Kelly and John A. Macready
Fokker T-2
October 6, 1922
1923 430 km/h (267.16 mph)
USA
Alford J Williams
Curtiss R2C-1
November 4, 1923
5,300 km (3,293 miles)
USA
Smith and Richter
De Havilland DH.4B
August 28, 1923
11,145 m (36,565 ft)
France
Joseph Sadi-Lecointe
Nieuport-Delage
October 30, 1923
1924 448 km/h (278.47 mph)
France
Florentine Bonnet
Bernard Ferbois V2
December 11, 1924
746 kW (1,000 hp)
Great Britain
Napier
Cub
1925
1926 5,396 km (3,352.92 miles)
France
Costes and Rignot
Breguet 19 GR
October 19, 1926
1927 479 km/h (297.83 mph)
Italy
Mario de Bernardi
Macchi M.52
November 4, 1927
6,294 km (3,911 miles)
USA
Chamberlin and Levine
Bellanca
June 6, 1927
11,710 m (38,418 ft)
USA
CC Champion
Wright Apache
July 25, 1927
1928 513 km/h (318.57 mph)
Italy
Mario de Bernardi
Macchi M.52bis
March 30, 1928
7,665.3 km (4,763.81 miles)
Italy
Arturo Ferrarin and Carlo del Prete
SIAI-Marchetti S.64
July 5, 1928
1929 583 km/h (362 mph)
Italy
Giuseppe Motta
Macchi M.67
August 22, 1929
8,029.4 km (4,989.26 miles)
France
D. Costes and P.Codos
Breguet 19
December 17, 1929
12,739 m (41,795 ft)
Germany
Willi Neuenhofen
Junkers W 34
May 26, 1929
56,000 kg (123,457 lb)
Germany
Dornier
Dornier Do X
1,119 kW (1,500 hp)
Great Britain
Rolls Royce
Rolls Royce R
1930 8,188.8 km (5,088.28 miles)
Italy
Maddalena and Cecconi
SIAI-Marchetti S.M.64bis
June 2, 1930
13,157 m (43,166 ft)
USA
Apollo Soucek
Wright Apache
June 4, 1930
1931 655 km/h (406.94 mph)
Great Britain
GH Stainforth
Supermarine S.6B
September 29, 1931
10,371 km (6,444.27 miles)
France
Le Brix and Doret
Dewoitine D-33
June 10, 1931
2,280 kW (3,058 hp)
Italy
Fiat
Fiat AS.6
1932 10,601.5 km (6,587.45 miles)
France
Bossoutrot and Rossi
Blériot 110
March 26, 1932
13,404 m (43,976 ft)
Great Britain
Cyril F Uwins
Vickers Vespa
September 16, 1932
1933 682 km/h (423.82 mph)
Italy
Francesco Agello
Macchi-Castoldi MC-72
April 18, 1933
13,660 m (44,816 ft)
France
Gustave Lemoine
Potez 50
September 28, 1933
1934 709 km/h (440.68 mph)
Italy
Francesco Agello
Macchi-Castoldi MC-72
October 23, 1934
14,432 m (47,352 ft)
Italy
Renato Donati
Caproni Ca 113 AQ
April 11, 1934
1935 14,575 m (47,818 ft)
USSR
Vladimir Kokkinaki
Polikarpov TsKB-3 (I-15)
November 21, 1935
1936 15,223 m (49,944 ft)
Great Britain
S.R. Swain
Bristol 138
September 28, 1936
1937 16,440 m (53,937 ft)
Great Britain
MJ Adam
Bristol 138
1938 11,650 km (7,239 miles)
Japan
Takahashi and Sekine
Gasuden Koken
May 16, 1938
17,083 m (56,046 ft)
Italy
Mario Pezzi
Caproni 161bis
October 22, 1938
1939 755 km/h (469.22 mph)
Germany
Fritz Wendel
Messerschmitt Me-209 V1
April 26, 1939
12,936 km (8,038 miles)
Italy
Tondi, Degasso, Vignoli
Savoia-Marchetti S.M.75
August 1, 1939
544 kgf thrust (1,200 lbf thrust)
Germany
Pabst von Ohain
Heinkel HeS 3B
1940
1941 1,004 km/h km/h (623.85 mph)
Germany
Heini Dittmar
Messerschmitt Me 163A
October 2, 1941
748 kgf thrust (1,650 lbf thrust)
Germany
Walter HWK
R11
1942 63,500 kg (140,000 lb)
USA
Boeing
Boeing XB-29 Superfortress
910 kgf thrust (2,006 lbf thrust)
Germany
Junkers Motoren
Jumo 004 B
1943 75,500 kg (166,447 lb)
Germany
Junkers
Junkers Ju-390
1,700 kgf thrust (3,748 lbf thrust)
Germany
Walter HWK
109-509 A-2
1944 94,339 kg (207,981 lb)
Germany
Blohm & Voss
Blohm & Voss V238 V1
2,000 kgf thrust (4,410 lbf thrust)
Germany
Walter HWK
109-509 C
1945
1946 18,081 km (11,235 miles)
USA
Thomas D Davies
Lockheed P2V-1 Neptune
October 1, 1946
140,614 kg (310,000 lb)
USA
Convair
Convair XB-36
2,722 kgf thrust (6,000 lbf thrust)
USA
Reaction Motors Inc
XLR 11-RM-5
1947 1,434 km/h (891.07 mph)
USA
Charles Yeager
Bell X-1
November 6, 1947
181,437 kg (400,000 lb)
USA
Hughes Aircraft Co
H-4 Hercules
1948 1,540 km/h (957 mph)
USA
Charles Yeager
Bell X-1
March 26, 1948
19,507 m (64,000 ft)
USA
Charles Yeager
Bell X-1
May 26, 1948
2,740 kgf thrust (6,041 lbf thrust)
USSR
Klimov
Klimov VK-1
1949 37,165 km (23,093 miles)
USA
James Gallagher
Boeing B-50A
March 2, 1949
21,916 m (71,902 ft)
USA
Frank Everest
Bell X-1
August 8, 1949
2,948 kgf thrust (6,500 lbf thrust)
Great Britain
Rolls Royce
Rolls Royce Avon 100
1950 3,969 kgf thrust (8,750 lbf thrust)
USA
Pratt & Whitney
Pratt & Whitney J48-P-5
1951 1,997 km/h (1,240.89 mph)
USA
William Bridgeman
Douglas D-558-2
August 7, 1951
24,230 m (79,494 ft)
USA
William Bridgeman
Douglas D-558-2
August 15, 1951
4,400 kgf thrust (9,700 lbf thrust)
USA
Pratt & Whitney
Pratt & Whitney J57-P-3
1952 190,509 kg (420,000 lb)
USA
Boeing
B-52A Stratofortress
1953 2,655 km/h (1,650 mph)
USA
Charles Yeager
Bell X-1A
December 12, 1953
25,376 m (83,253 ft)
USA
Marion Carl
Douglas D-558-2
August 21, 1953
8,700 kgf thrust (19,181 lbf thrust)
USSR
Mikulin
Mikulin AM-3D
1954 27,566 m (90,440 ft)
USA
Arthur Murray
Bell X-1A
August 26, 1954
9,500 kgf thrust (20,945 lbf thrust)
USSR
Mikulin
Mikulin AM-3M
1955 10,000 kgf thrust (22,047 lbf thrust)
USSR
Lyulka
Lyulka AL-7F TRD-31
1956 3,370 km/h (2,094 mph)
USA
Milburn Apt
Bell X-2
September 27, 1956
38,376 m (125,907 ft)
USA
Iven Kincheloe
Bell X-2
September 7, 1956
204,117 kg (450,000 lb)
USA
Boeing
B-52C Stratofortress
1957 39,147 km (24,325 miles)
USA
Archie Old Jr
B-52B Stratofortress
January 18, 1957
12,251 kgf thrust (27,008 lbf thrust)
USSR
Lyulka
Lyulka AL-21F
1958 221,353 kg (488,000 lb)
USA
Boeing
B-52G Stratofortress
1959
1960 3,534 km/h (2,196 mph)
USA
Joseph Albert Walker
North American X-15
August 4, 1960
41,605 m (136,500 ft)
USA
Robert Michael White
North American X-15
August 12, 1960
1961 6,587 km/h (4,093 mph)
USA
Robert White
North American X-15
November 9, 1961
1962 6,605 km/h (4,104 mph)
USA
Joseph Walker
North American X-15
June 27, 1962
95,936 m (314,750 ft)
USA
Robert White
North American X-15
July 17, 1962
1963 107,960 m (354,200 ft)
USA
Joseph Albert Walker
North American X-15
August 22, 1963
1964 249,476 kg (550,000 lb)
USA
North American
XB-70A Valkyrie
14,742 kgf thrust (32,500 lbf thrust)
USA
Pratt & Whitney
Pratt & Whitney JT11D-20B
1965 250,000 kg (551,147 lb)
USSR
Antonov
Antonov An-22 Anteus
1966
1967 7,297 km/h (4,534 mph)
USA
Pete Knight
North American X-15
October 3, 1967
1968 348,359 kg (768,000 lb)
USA
Lockheed
C-5A Galaxy
20,000 kgf thrust (44,095 lbf thrust)
USSR
Kuznetsov Design Bureau
Kuznetsov NK-144
1969 21,296 kgf thrust (46,950 lbf thrust)
USA
Pratt & Whitney
Pratt & Whitney JT9D-7AW
1970 351,534 kg (775,000 lb)
USA
Boeing
Boeing 747-200B
1971 377,842 kg (833,000 lb)
USA
Boeing
Boeing 747-200F
22,861 kgf thrust (50,400 lbf thrust)
USA
General Electric
General Electric CF6-50C
1972
1973 23,496 kgf thrust (51,800 lbf thrust)
USA
General Electric
General Electric CF6-50E
1974
1975
1976 24,145 kgf thrust (53,230 lbf thrust)
USA
Pratt & Whitney
Pratt & Whitney JT9D-59B
1977
1978 24,190 kgf thrust (53,330 lbf thrust)
USA
Pratt & Whitney
Pratt & Whitney JT9D-7Q1
1979
1980 24,875 kgf thrust (54,840 lbf thrust)
USA
Pratt & Whitney
Pratt & Whitney JT9D-7Q2
1981
1982 25,401 kgf thrust (56,000 lbf thrust)
USA
Pratt & Whitney
Pratt & Whitney JT9D-7H1
1983
1984
1985 404,994 kg (892,859 lb)
USSR
Antonov
Antonov An-124 Condor
26,762 kgf thrust (59,000 lbf thrust)
USA
General Electric
General Electric CF6-80C2A1
1986 40,213 km (24,987 miles)
USA
Dick Rutan and Jeana Yeager
Voyager
December 23, 1986[1]
1987 27,896 kgf thrust (61,500 lbf thrust)
USA
General Electric
General Electric CF6-80C2A5
1988
1989 404,994 kg (892,859 lb)
USSR
Antonov
Antonov An-225 Mriya
1990 27,928 kgf thrust (61,570 lbf thrust)
USA
Pratt & Whitney
Pratt & Whitney 4060A
1991
1992
1993
1994
1995
1996
1997
1998
1999
2000
2001
2002
2003
2004 112,000 m (367,454 ft)
USA
Brian Binnie
Scaled Composites SpaceShipOne
October 4, 2004[2]
2005 41,467 km (25,766 miles)
USA
Steve Fossett
Virgin Atlantic GlobalFlyer
March 3, 2005[3]
2006
2007

Reference : https://en.wikipedia.org/wiki/Aircraft_records

What Is the Most Fuel-Efficient Airplane?

What Is the Most Fuel-Efficient Airplane?

There is no greater concern among pilots and airplane owners today than the cost of fuel. Prices vary widely from airport to airport, but $5 is often on the low end and $7 a gallon is not the top. And in many instances jet fuel costs more than avgas, a reversal of traditional pricing. Fuel costs for most airplane owners have doubled in the past year and nobody can predict future trends, but higher prices sure seem more likely than lower.

So now we all want to own and fly the most fuel-efficient airplane, but what is it? There is no single answer because as with all desirable characteristics in an airplane we always must trade one attribute for another. If fuel were really the only driving factor in finding the most efficient aircraft, powered parachutes and motorgliders would win hands down. Or everybody in a jet would cram into a piston-powered airplane because jets can’t match the fuel efficiency of a reciprocating engine driving a propeller. What we really want is not the most fuel-efficient airplane possible, but the most thrifty one that suits our mission.

The aviation industry uses a metric called specific range to measure fuel efficiency. Specific range is the number of miles — normally a fraction except for piston singles — that an airplane flies through the air per pound of fuel consumed. For example, a piston airplane with a true airspeed of 150 knots while burning 12 gallons per hour (72 pounds) would have a very good specific range of 2.08. A business jet cruising at 440 knots true burning 1,200 pounds per hour (pph) has a specific range of 0.37, good for a jet.

Specific range can be calculated for the cruise condition, as I have done above. But a more useful measure is specific range for the entire trip. When fuel consumed for start, taxi, climb, approach and landing are all measured against the block speed, the specific range will be worse than for cruise.

Though we buy fuel in volume — by the gallon or liter — and the tank’s capacity is determined by volume, engines burn fuel and air in mass, so measuring fuel consumption in pounds is more useful. At standard temperature of 15° C (59° F) a gallon of avgas weighs about 6 pounds. A gallon of jet-A fuel at standard temperature weighs in at about 6.7 pounds. Avgas density changes very little over the normal temperature range, but jet fuel density can vary by several percent, so a gallon of really cold jet-A is going to take up less space in the tanks than a warm gallon, but the useful work of the fuel will still be measured by the pound.

By using the specific range calculation it’s easy to compare airplanes to one another or to measure the efficiency of power settings. For example, let’s say that a powerful piston single can cruise at 140 knots while burning 72 pounds per hour (pph), which equals 12 gallons per hour. That yields a specific range of 1.94. If you increase the power and fuel flow to 132 pph and the true airspeed climbs to 180 knots, the specific range is down to 1.36. The airspeed increased by about 22 percent, but the fuel efficiency decreased by about 30 percent.

As you can see, the fuel efficiency of a piston airplane is largely in the hands of its pilot. More speed equals less efficiency. The greatest specific range airspeed for a piston airplane is so slow that few of us would ever contemplate it unless we absolutely had to stretch the fuel all the way back to shore. In general, an indicated airspeed at about the best rate of climb speed is also the most fuel-efficient speed in level cruise. The heavier the airplane, the higher its maximum efficiency speed, so to achieve true maximum range you must continuously slow down as fuel burns off to maintain the same angle of attack and thus drag.

This amazing gain in fuel efficiency at lower speeds was driven home to me flying to Oshkosh this past summer. Controllers told me to pull the indicated airspeed back to 120 knots to stay in trail of slower IFR traffic ahead. My Baron was burning about 30 gph to indicate 170 knots, but it took about 16 gph to hold the 120 indicated. Fuel flow was nearly halved to cut airspeed by less than a third.

Pilots of turbine airplanes actually have less control over the fuel efficiency of their flights because there are so many variables, first among them being air traffic control. Turbine engines are at their least efficient down low where the air is dense. As the airplane climbs and the air thins, the turbine produces less power and thus consumes less fuel, but the drag of the thinning air on the airplane decreases faster than the power from the engine drops, so the airplane speeds up and the fuel flow goes down. There is an optimum altitude for every turbine airplane at its present weight and any level lower than that optimum decreases fuel efficiency. In crowded airspace the pilots of turbine airplanes seldom are cleared for unrestricted climb to the optimum altitude, so the airplane doesn’t come close to matching its potential specific range. And on the ground at idle a turbine burns a surprisingly large percentage of its optimum cruise fuel flow, so takeoff delays really cut into fuel efficiency in a jet compared to a piston engine.

The other big variables for turbine airplane pilots are air temperature and weight. When air temperature aloft is warmer than the international standard atmosphere (ISA) the air is less dense so the engines have less mass — fewer pounds — of air to compress and burn. There is also a slight decrease in drag because of the lower air density, and engine fuel flow goes down some, but in this case engine power falls off faster than the drag benefit, so at warm temperatures a turbine-powered airplane specific range suffers.

The weight of a jet also greatly impacts its specific range. When heavy the highest altitude that the wing can support effectively is lower, so fuel flow is higher, drag greater, and the airplane travels a shorter distance on the same amount of fuel. Because the weight of necessary fuel and payload changes with almost every flight, the specific range of a jet, or less so a turboprop, changes, too.

Jets have long had large swings in fuel efficiency at various airspeeds with the longest range cruise speeds typically much slower than high-speed cruise, often 50 to 100 knots slower. Higher airspeeds in any airplane create more drag, but there is a particular issue at jet speeds that is linked to the speed of sound, or Mach 1. At speeds below Mach 1 the air behaves, well, like air. The molecules of air bounce around and get out of the way of the advancing airplane. But when airflow approaches or exceeds the speed of sound it begins to compress and behave more like a fluid and drag shoots way up. At the point where airflow hits Mach 1 a shock wave forms, separating the faster and slower moving air streams, and that wave is like sticking a speedbrake up into the air stream. Drag at the point of the wave skyrockets and gobs of power is required to overcome the increase.

No operational civilian jet can exceed Mach 1 in level flight or is approved to go that fast even in a dive over the United States, but the air flowing over the wing and fuselage of typical jets does reach the speed of sound. The reason is that as the jet moves through the air at say Mach .80, or 80 percent of the speed of sound, air flowing over the wing and fuselage must accelerate to pass over the airplane. This is called local airflow, and depending on the shape of the wing and fuselage, the local airflow can accelerate to Mach 1 at normal cruise speed and a shock wave forms even though the airplane is not supersonic. That speed is called the “critical Mach” of the wing because that is where the drag jumps up and fuel efficiency goes into rapid decline.

Aerodynamicists have made many advances in jet wing design to control the shock waves formed in the local airflow. Wing sweep makes the air believe the airfoil section of the wing is thinner so it accelerates less to pass over and under the wing. The “super critical” wing design has a thicker and flatter forward section with a concave cusp on the underside trailing edge of the wing. The flat section delays the formation of a shock wave, and the cusp helps move it back and minimize its intensity.

The newest wing designs, such as those on the Falcon 7X and Gulfstream G650 that is in development, control shock wave formation so well that the difference between maximum range airspeed and high-speed cruise has shrunk. Gulfstream predicts that its new G650 wing will be so aerodynamically efficient that maximum range will be achieved at a Mach .85 cruise speed, which is actually high-speed cruise for most jets. Gulfstream engineers say the G650 will fly 7,000 nm with IFR reserves in still air at Mach .85, and no further at any slower airspeed.

Another big factor in fuel efficiency is the size of the airplane. A bigger fuselage has to shove more air out of the way so it creates more drag and thus requires more power. As with the wing, the shape of the fuselage can reduce drag somewhat, and its length to diameter ratio — the fineness ratio — can also help at higher airspeeds, but there is no getting around the fact that more cabin space is going to require more fuel to push through the air. Passengers want the most room possible for comfort, but it must be paid for at the fuel pump.

As you can see, everything we all want in an airplane — largest possible cabin, fastest possible speed and longest range — all work against fuel efficiency. So even at today’s fuel prices we make trades. Higher fuel cost may force some to trade down, but it doesn’t alter the desirability of speed, comfort and range.

However, I think it’s useful to examine some airplanes that were designed with fuel efficiency high on the list, or perhaps, even at the top of the list. Among those are the earlier Mooneys, particularly the 201 M20J model, and the Piaggio P180 Avanti that is discussed fully on page 46. Among the jets that offer good efficiency and range are the Falcon 50/900, Learjet 30s and Gulfstreams.

By all accounts Al Mooney strove for efficiency even though he did his fundamental design work in the 1950s when fuel was relatively cheap and plentiful. His goal was to design an airplane that delivered the most speed with the least powerful engine, and he was successful. Mooney chose the then new to general aviation laminar flow airfoil to help control drag, and designed a fuselage and cabin big enough for people, but with no extra room, particularly in cabin height. The first wings were made from wood because the plywood skin could be formed into the very smooth shape needed to optimize laminar airflow. When the wing was changed to aluminum Mooney used flush rivets to retain the smoothness.

The Mooney series reached its height of fuel efficiency versus performance in the 1970s with development of the 201. The goal was to fly at least one mile per hour for each horsepower available from the engine. The 201 has a 200 hp Lycoming four-cylinder IO-360, and it managed 201 mph true airspeed with everything just right.

The country had suffered two oil shocks brought on by supply disruptions from the Middle East, and the topic of fuel availability and price was paramount in aviation. Mooney owners had always bragged about how little fuel they used, but the quirks of the airplane kept it in a niche. The 201 with its new sloped windshield, greatly improved cowling and, most importantly, its vastly better cabin furnishings broke out of the niche and thousands were sold. The airplane cruises at about 160 knots on 60 pph (10 gallons) fuel flow for an excellent specific range of 2.66.

As memory of the oil crisis faded, and some excellent and very powerful engines became available from Continental, Mooney introduced models with ever more power and higher cruise speed. Top airspeed is what pilots valued most over the past 10 to 15 years, and Mooney delivered with its Acclaim at the top of the piston single lineup for max cruise speed. If you pull the power back in an Acclaim or Ovation to 60 pph, you probably won’t go as fast as the 201 because the new airplanes are heavier. And it would be silly to cruise that slowly in the powerful Mooneys because you paid upfront for the speed, and you are wearing out the airplane while traveling fewer miles. We pilots tend to think in terms of hours, but it’s miles flown over the ground that count and cost.

We are also seeing steady improvement in the fuel efficiency of jets. The improvements are coming from both the engines and aerodynamics. And jets are now available in a greater range of sizes and performance categories, so you can match the mission and its required fuel to the airplane more efficiently. The aerodynamics at piston airplane airspeeds are so mature that we won’t see the same degree of efficiency gains in those airplanes, but small improvements are constant.

The drive for fuel efficiency in jets was more a desire for longer range and there are two very important ways to increase range — lighter weight and lower drag, and the ability to fly higher.

The Falcon 50 is an example of a jet where emphasis was placed on weight and drag and it achieved very long range in the 1970s when it was designed. The three engines, supercritical airfoil and very sophisticated structural design controlled empty weight and drag, yielding intercontinental range. The Learjet 35 and 36, and the Gulfstreams, all emphasized climb capability and could go directly to 41,000 feet or higher at maximum weight. At those flight levels fuel flows and drag were reduced and range increased.

Now nearly every recently designed business jet can reach 41,000 feet and above, so that has leveled the playing field and put the emphasis back on the lightest possible structure, most efficient wing and lowest drag shape, and a number of new jets are capable of astonishing nonstop trips.

So, to answer the question, what is the most fuel-efficient airplane? It’s the one that delivers the most speed, comfort and range that you demand with the highest specific range. A pilot’s life remains one of tradeoffs, and fuel efficiency is now a big item on the list, but the other choices are still there, too.

Reference : http://www.flyingmag.com/what-most-fuel-efficient-airplane

All Engine Type Of Jets

All Engine Type Of Jets

7-6-2016 2-46-16 AM

The basic idea of the turbojet engine is simple. Air taken in from an opening in the front of the engine is compressed to 3 to 12 times its original pressure in compressor. Fuel is added to the air and burned in a combustion chamber to raise the temperature of the fluid mixture to about 1,100°F to 1,300° F. The resulting hot air is passed through a turbine, which drives the compressor. If the turbine and compressor are efficient, the pressure at the turbine discharge will be nearly twice the atmospheric pressure, and this excess pressure is sent to the nozzle to produce a high-velocity stream of gas which produces a thrust.

Substantial increases in thrust can be obtained by employing an afterburner. It is a second combustion chamber positioned after the turbine and before the nozzle. The afterburner increases the temperature of the gas ahead of the nozzle. The result of this increase in temperature is an increase of about 40 percent in thrust at takeoff and a much larger percentage at high speeds once the plane is in the air.

The turbojet engine is a reaction engine. In a reaction engine, expanding gases push hard against the front of the engine. The turbojet sucks in air and compresses or squeezes it. The gases flow through the turbine and make it spin. These gases bounce back and shoot our of the rear of the exhaust, pushing the plane forward.

7-6-2016 2-47-10 AM

A turboprop engine is a jet engine attached to a propellor. The turbine at the back is turned by the hot gases, and this turns a shaft that drives the propellor. Some small airliners and transport aircraft are powered by turboprops.

Like the turbojet, the turboprop engine consists of a compressor, combustion chamber, and turbine, the air and gas pressure is used to run the turbine, which then creates power to drive the compressor.

Hungarian, Gyorgy Jendrassik who worked for the Ganz wagon works in Budapest designed the very first working turboprop engine in 1938. Called the Cs-1, Jendrassik’s engine was first tested in August of 1940; the Cs-1 was abandoned in 1941 without going into production due to the War. Max Mueller designed the first turboprop engine that went into production in 1942.

7-6-2016 2-48-59 AM

A turbofan engine has a large fan at the front, which sucks in air. Most of the air flows around the outside of the engine, making it quieter and giving more thrust at low speeds. Most of today’s airliners are powered by turbofans. In a turbojet all the air entering the intake passes through the gas generator, which is composed of the compressor, combustion chamber, and turbine. In a turbofan engine only a portion of the incoming air goes into the combustion chamber.

The remainder passes through a fan, or low-pressure compressor, and is ejected directly as a “cold” jet or mixed with the gas-generator exhaust to produce a “hot” jet. The objective of this sort of bypass system is to increase thrust without increasing fuel consumption. It achieves this by increasing the total air-mass flow and reducing the velocity within the same total energy supply.

7-6-2016 2-49-50 AM

This is another form of gas-turbine engine that operates much like a turboprop system. It does not drive a propellor. Instead, it provides power for a helicopter rotor. The turboshaft engine is designed so that the speed of the helicopter rotor is independent of the rotating speed of the gas generator. This permits the rotor speed to be kept constant even when the speed of the generator is varied to modulate the amount of power produced.

7-6-2016 2-50-38 AM

The most simple jet engine has no moving parts. The speed of the jet “rams” or forces air into the engine. It is essentially a turbojet in which rotating machinery has been omitted. Its application is restricted by the fact that its compression ratio depends wholly on forward speed. The ramjet develops no static thrust and very little thrust in general below the speed of sound. As a consequence, a ramjet vehicle requires some form of assisted takeoff, such as another aircraft.

It has been used primarily in guided-missile systems. Space vehicles use this type of jet.