Indicated to Calibrated Airspeede Airpeed Graphed Agains Indicated

Airspeed

Since the airspeed is below the trim airspeed (decreasing lift and drag) at position 4, the airplane system will automatically respond so as to regain the equilibrium condition by pitching the nose downward.

From: International Geophysics , 2014

Aircraft Instruments

W.B. Ribbens , in Encyclopedia of Physical Science and Technology (Third Edition), 2003

XVIII Air Data System

Measurement of critical flight variables such as airspeed and altitude have long been important in aircraft. In lower performance general aviation aircraft, such measurements are still today performed by stand-alone pneumatic-mechanical instruments that respond as required to static, dynamic, or total pressure. In high-performance (and high cost) general aviation, transport, and military aircraft, these and other variables must be computed to relatively high accuracy and must be available in a computer-based instrument where these variables can be combined in known functional relationships to evaluate and optimize aircraft performance.

An air data system provides calculations of flight variables, including calibrated airspeed, true airspeed, equivalent airspeed, Mach number, free-stream static pressure and outside air temperature, air density, pressure altitude, density altitude, angle of attack, and side slip angle. The static pressure p _s is the absolute pressure of the still air at any point in the atmosphere. An approximate measurement of static pressure can be obtained by means of a port along the side of the fuselage (called the static port). Total pressure p_t is the pressure sensed in a tube that is open at the front, closed at the rear, and directed toward the free-stream air velocity vector.

The various airspeeds are derived by computation from measurements of total pressure, static pressure, and absolute air temperature T. Impact pressure q _c is defined as

${\text{q}}_{\text{c}}={\text{p}}_{\text{t}}-{\text{p}}_{\text{s}},$

which for subsonic flight is given by

${\text{q}}_{\text{c}}={\text{p}}_{\text{s}}{\left[1+\left(\frac{\gamma -1}{\gamma }\right)\frac{\rho }{2{\text{p}}_{\text{s}}}{\text{V}}^2\right]}^{\frac{\gamma }{1-\gamma }}-{\text{p}}_{\text{s}},$

where	ρ   =   local air density (slug/ft3)
	V  =   true airspeed (ft/sec)
	γ   =   ratio of specific heats for air   =   1.4

Air density can be obtained from local air static pressure and temperature

$\rho =\frac{{\text{p}}_{\text{s}}}{\text{g}\text{R}\text{T}},$

where	R  =   53.3   ft per degree Kelvin
	g  =   acceleration of gravity

The true airspeed can be obtained by solving the q _c equation for V. Mach number M is the ratio of V to the local speed of sound a:

$\text{M}=\frac{\text{V}}{\text{a}},$

where

$\text{a}=\sqrt{\gamma \text{g}\text{R}\text{T}}.$

Calibrated airspeed is the value that would be obtained from the q _c equation if measurements were made at sea level on a standard day at which ρ   =   ρ₀  =   0.002378 slug per cubic feet.

The above measurements of speed also require a measurement of local outside air temperature. A temperature sensor mounted on the outside of an aircraft measures a temperature that is higher than the still air temperature due to friction and compressibility. The measured temperature T _M is given by

${\text{T}}_{\text{M}}=\text{T}\left[1+\frac{\gamma -1}{2}\eta {\text{M}}^2\right],$

where η is the empirically determined constant for the sensor.The air data system solves this equation for T from measurements of T _m and M.

Altitude measurements are derived from the measurement of static pressure and a standard model for the atmosphere in the following equations:

$\begin{array}{l}\frac{\text{d}{\text{p}}_{\text{s}}}{\text{d}\text{h}}=-\text{g}\rho \hfill \\ \overline{\text{T}}(\text{h})={\text{T}}_{\text{o}}-{\alpha }_{\text{a}}\text{h}\hfill \end{array}$

where	h  =   true altitude
	α a   =   lapse rate for the atmosphere
	= 0.003566   °F/ft

Integrating the above equation in conjunction with the air density equation yields the standard atmosphere model for pressure vs altitude:

$\text{p}={\text{p}}_{\text{o}}{\left(1-\frac{\alpha \text{h}}{{\text{T}}_{\text{o}}}\right)}^{\frac{1}{\alpha \text{R}}}.$

Altitude can be computed using this relationship and measurements of static pressure and outside air temperature. Angle of attack and side slip angle are measured using rotary position sensors connected to movable vanes that are mounted on the surface of the aircraft as illustrated in Fig. 16. The rotary position sensor is essentially a potentiometer having a movable vane attached to the rotary shaft. The movable vane aligns with the air velocity vector. The angle between this vane and the aircraft longitudinal axis is the angle of attack. Consequently, the output voltage from the potentiometer is an essentially linear function of angle of attack.

FIGURE 16.

The vane assembly is mounted flush with the aircraft surface via a mounting flange. The plane of symmetry of the flange is in the plane of symmetry of the aircraft for side slip measurements and is in a horizontal plane for angle of attack measurements.

A block diagram of an air data system is depicted in Fig. 17.

FIGURE 17. Air data system block diagram.

Sensors for measuring pressure are available in a variety of technologies that often incorporate a diaphragm that seals a closed chamber and is coupled to a displacement sensor. Another class of pressure sensor involves fabricating the diaphragm from doped silicon whose resistivity varies with stress due to a piezo-resistive property. Stress dependent resistance is readily converted to a measurement of pressure for a given diaphragm configuration via a bridge circuit or the like.

Temperature sensors often consist of a small coil of wire whose resistance varies with temperature. Alternatively, a semiconductor slab can also provide a temperature dependent resistance. Water ingestion and icing can cause significant errors that must be minimized by design.

The calculated values for the various air data variables are used in cockpit instrumentation via a suitable display such as an analog or digital display or a cathode ray tube or solid sate equivalent.

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Aircraft Speed and Altitude

William Gracey , ... Tony Whitmore , in Encyclopedia of Physical Science and Technology (Third Edition), 2003

II.D True Airspeed Equations

From Eq. (10) , the true airspeed V is seen to be a function of the dynamic pressure q and the air density ρ as follows:

(18) $\text{V}=\sqrt{2\text{q}/\rho }$

As shown by the following equations, the dynamic pressure is a function of the static pressure p and the Mach number M:

(19) $\text{q}=\gamma \text{p}{\text{M}}^2/2$

and the air density ρ is a function of the static pressure and the air temperature T:

(20) $\rho ={\rho }_0\frac{\text{p}{\text{T}}_0}{{\text{p}}_0\text{T}}$

Since Mach number is derived from measurments of q _c and p, the measurement of these pressures can be combined with the measurement of T to provide indications of true airspeed, as discussed earlier. In contrast to the airspeed indicator, which indicates true airspeed only at sea level (under standard conditions), the true-airspeed indicator indicates true airspeed independent of altitude. Table I gives values of true airspeed V in knots for values of calibrated airspeed V _c in knots and pressure altitude H in geopotential feet.

TABLE I. True Airspeed V for Values of Calibrated Airspeed V_c and Values of Pressure Altitude H ^a

H	V c (knots)
(geopotential ft)	100	200	300	400	500	600	700	800	900	1000
0	100.0	200.0	300.0	400.0	500.0	600.0	700.0	800.0	900.0	1000
5000	107.7	215.0	321.6	427.4	532.2	635.8	740.3	847.3	955.2	1064
10,000	116.2	231.6	345.4	457.2	566.8	674.5	785.0	900.5	1018	1136
15,000	125.8	250.0	371.5	489.4	603.8	716.3	835.2	960.9	1089	1218
20,000	137.2	270.5	400.1	524.4	643.4	763.0	892.4	1030	1170	1310
25,000	148.7	293.4	431.5	562.0	686.6	816.2	958.0	1109	1263	1418
30,000	162.4	318.9	465.9	602.6	735.4	877.5	1034	1201	1370	1541
35,000	178.0	347.4	503.5	646.9	791.6	948.7	1122	1307	1494	1682
40,000	199.1	385.6	553.7	708.9	871.5	1049	1245	1454	1666	1878
45,000	223.7	429.1	610.0	782.4	967.0	1169	1392	1629	1869
50,000	251.0	476.4	671.6	865.7	1076	1306	1559	1827
55,000	281.3	527.3	740.3	960.3	1199	1460	1747
60,000	314.9	581.8	817.9	1068	1340	1636	1961
65,000	351.8	640.4	906.0	1191	1499	1835
70,000	394.8	709.9	1013	1338	1690	2073
75,000	440.3	785.3	1130	1501	1901
80,000	489.4	870.3	1263	1684	2139
85,000	540.2	962.9	1408	1885
90,000	596.2	1071	1576	2111
95,000	656.2	1193	1766
100,000	722.2	1330	1979

From Gracey, W. (1981). "Measurement of Aircraft Speed and Altitude," Wiley–Interscience, New York.

a

True-airspeed values V are given in knots.

The airspeed indicator, true-airspeed indicator, and Machmeter measure speed with respect to the air mass. Since the air mass can move with respect to the ground, the measurement of ground speed, the speed of basic importance to air navigation, must be derived from ground navigational aids or an inertial guidance system.

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Reuse, Recycling, and Resource Recovery

In Environmental Engineering (Fourth Edition), 2003

Problems

14.1

Estimate the critical speed for a trommel screen, 2 m in diameter.

14.2

What air speed is required in an air classifier to suspend a spherical piece of glass 1 mm in diameter? (See Eq. 9.4; assume ρ _s = 2.65 g/cm³, C _D = 2.5, μ = 2 × 10⁻⁴ P, and ρ = 0.0012 g/cm³.)

14.3

Refer to Fig. 14-16. By replotting the curves, estimate the Rosin–Rammler constants n and x _c . Compare these with the data in Table 14-1.

Figure 14-16. Feed and product curves in shredding, for Problems 14.3 and 14.4

14.4

If the Bond work index is 400 kWh/ton, what is the power requirement to process 100 ton/h of the refuse described in Fig. 14-16?

14.5

Estimate the heating value of MSW based on the ultimate analysis shown in Table 12-1.

14.6

A power plant burns 100 ton/h of coal. How much air is needed if 50% excess air is used?

14.7

An air classifier performance is:

	Organics (kg/h)	Inorganics (kg/h)
Feed	80	20
Product	60	10
Reject	20	10

Calculate the recovery, purity, and efficiency.

14.8

Using the data in Example 14.6, what percent of the heating value can be saved by reducing the moisture content of refuse to 0?

14.9

In Example 14.6, what fraction of the coal is "wasted" owing to the loss of heat in the stack gases (if steam is produced)?

14.10

What is the "wasted" coal in Problem 14.9 heat equivalent in barrels of oil? The heating value of a barrel of oil is 6 × 10⁶ kJ.

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Fluid Biomechanics

David E. Alexander , in Nature's Machines, 2017

In fact, flying squirrels and flying lizards often avoid equilibrium gliding. By starting out fast, with a quick drop to build up airspeed, they can decelerate while constantly increasing their angle of attack and plan to stall just as they reach their intended landing spot. (Their low aspect ratio wings can reach unusually high angles of attack before stalling ( Alexander, 2002).) Pulling off such a precise maneuver requires a sophisticated ability to judge distances and fine aerodynamic control, but most such purely gliding animals seem to perform it routinely.

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V/STOL Airplanes

Barnes W. McCormick , in Encyclopedia of Physical Science and Technology (Third Edition), 2003

III.B Aerodynamics of High Lift

The lift of an airplane wing is proportional to the product of the air mass density ρ, the wing area S , and the square of the airspeed V. The constant of proportionality is called the wing lift coefficient and is denoted by C _L. This relationship for the lift L can be written

$\text{L}=\frac{1}{2}\rho {\text{V}}^2\text{S}{\text{C}}_{\text{L}}$

Here, the quantity $\frac{1}{2}$ ρV ² is referred to as the dynamic pressure.

The wing lift coefficient C _L is dependent on the wing geometry and its angle relative to the airflow. For a given wing shape, C _L increases linearly with this angle, denoted as the angle of attack, up to a maximum value C _Lmax , beyond which the flow separates from the upper surface of the wing, causing a loss of lift. Here, the wing is said to be stalled. In steady, level flight, the lift of the wing must equal the weight W of the airplane. Thus, it follows that the slowest speed at which an airplane can fly and be supported by the wing lift, the stalling speed V _s, is given by V _s =(2W/ρSO _Lmax )^1/2. To achieve good STOL performance, an airplane must therefore produce a high C _Lmax . Typically, an ordinary airfoil will produce a maximum lift coefficient of ∼1.6. To increase C _Lmax , movable surfaces known as flaps are fitted to the leading and trailing edges of the airfoil. Various configurations of leading and trailing edge flaps are shown in Fig. 11. Some of the flaps increase the chord of the airfoil when they are extended. An elaborate flap system, such as a double-slotted trailing edge flap combined with a Krueger leading edge flap, can achieve C _Lmax values of ∼3.0.

In order to achieve C _Lmax values higher than those possible with unpowered flap systems, one must utilize power. Here, the airflow around the airfoil is energized so as to delay flow separation or increase the pressure difference (or both), and hence the lift, from the upper surface of the airfoil to the lower. The product of the airfoil chord and the mean velocity over the upper surface minus that over the lower surface is referred to by the aerodynamicist as the circulation Γ. Stated precisely, the circulation is defined as the closed line integral about the airfoil of the velocity vector tangent to the direction of integration. A well-known theorem, the Kutta-Joukowski theorem, states that the lift per unit span on an airfoil is related to the circulation by

$\text{L}=\rho \text{V}\Gamma$

The jet flap and other similar high-lift devices increase the circulation around an airfoil to values higher than can be achieved without power. This action caused by the trailing sheet of high momentum air was defined earlier as circulation control.

A parameter that characterizes the performance of a jet flap is the momentum coefficient C _μ. This coefficient is equal to the momentum in the jet per unit wing area divided by the free-stream dynamic pressure:

${\text{C}}_{\mu }={\text{m}}_{\text{j}}{\upsilon }_{\text{j}}/\text{q}\text{S}$

Here, m _j is the mass flow rate from the jet having a velocity of v _j, S is the planform area of the wing, and q represents the dynamic pressure.

The lift performance of the elliptically shaped airfoils discussed previously and being used on the X-wing CCR rotor is presented in Fig. 12. The bottom curve shows the lift coefficient as a function of C _μ for a fixed jet-flap deflection angle. Here, the behavior is similar to an ordinary jet flap. For the other curves, the jet is free to separate at the point determined by the jet momentum. As a result, the deflection angle of the jet increases with C _μ, giving a faster rate of increase of C _L with C _μ as compared with the jet-flap case. Observe that C _L values more than double those attainable without power are achieved with relatively low values of C. Of course, this increase in the lift coefficient requires some expenditure of power as well as some additional weight for the system to duct the air from the compressor. The power can be determined from the flux of kinetic energy in the jet and a knowledge of the compressor efficiency and losses in the air ducts.

FIGURE 12. Graph of lift performance of elliptical airfoils with circulation control.

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Aircraft Performance and Design

Francis Joseph Hale , in Encyclopedia of Physical Science and Technology (Third Edition), 2003

VII.C Propulsive Efficiency

Although the piston–prop has the lowest specific fuel consumption of all the air breathers, it is also the heaviest and has the largest drag, and its best-range airspeed is 24% lower than that of a comparable turbojet. Furthermore, the propeller efficiency drops off sharply at relatively low Mach numbers as the propeller tip speeds approach the sonic velocity. The turbojet produces its thrust by expanding all of the turbine exhaust gases through a nozzle, and it has the highest specific fuel consumption but is the lightest. It also has the fewest moving parts and the lowest drag, and its propulsive efficiency improves with airspeed. The turbofan and the turboprop are basically turbojets in which part of the exhaust gases is used to drive an ungeared multibladed ducted fan (the turbofan) or a propeller connected to the turbine drive shaft through a gearbox (the turboprop). The ratio of the mass of the cold air passing through the fan (or the propeller) to the mass of the hot air passing through the burners and turbines is called the bypass ratio. If the bypass ratio is zero, the turbofan becomes a pure turbojet; current operational bypass ratios are of the order of 5–6. As the bypass ratio increases, specific fuel consumption decreases and the turbofan begins to take on the characteristics of a turboprop with the exception that the ducted fan efficiency is essentially independent of the airspeed. Although it is not customary to refer to the bypass ratio of turboprops, it is of the order of 50.

Although turboprops have a specific fuel consumption that approaches that of the piston–prop and are four times lighter, their cruise speed has been limited by the degradation in the propeller efficiency at higher airspeeds. The fuel shortage in the mid-1970s and the subsequent increase in prices sparked the development of two ultrahigh bypass (UHB) engines, one known as a propfan and the other as an unducted fan (UDF), which in appearance resemble each other. They each have 2 rows of counter-rotating blades (6 or more in each row) that are approximately 12   ft in diameter; the blades do not resemble conventional propeller blades, being highly swept and having variable camber. The propfan is essentially a turboprop with a gearbox that reduces the high rpm of the turbine to that of conventional propellers; the rpm and the pitch of the blades can be varied. The unducted fan, on the other hand, is essentially a turbofan with a bypass ratio of the order of 28; it has no gearbox and operates at a constant rpm (that of the turbine), although the pitch of the blades can be varied through feathering and into reverse. With the duct removed, it is necessary to give greater consideration to the fan efficiency.

Although the propeller efficiency has been extended, it still drops off at the higher subsonic Mach numbers. Furthermore, the best-range airspeed of aircraft using both of these UHB engines is still lower than that of the pure turbojet and turbofan so that at competitive airspeeds the UHB lift-to-drag ratio will also be lower. In spite of these off-design point degradations at the high subsonic Mach numbers of current airliners, the specific range of UHB-equipped aircraft is sufficiently higher to make such an aircraft commercially attractive at Mach 0.8.

Although both types of engines were flown on demonstrator aircraft and design studies for commercial aircraft with UHB engines were started, current efforts seem to have ceased.

The thermal efficiency of gas turbine engines is strongly dependent on the turbine inlet temperature, which is primarily limited by the physical properties of the turbine blades. There are ongoing investigations into new materials and alloys, into techniques for forming turbine blades and fastening them to the hub, and for controlling the clearance between the blade tips and the hubs. As the turbine inlet temperature increases, so does the exhaust temperature, with a subsequent increase in acoustic noise, which must be considered in light of the drive for quieter engines.

In concluding this section, mention should be made of the increasing use of derated turboprops to replace piston (reciprocating) engines; derating means that the propeller selected cannot absorb all of the sea-level power developed by the engine. As the engine shaft power decreases with altitude, the propeller thrust power will remain constant until the maximum engine power available decreases to that value. The altitude at which the engine power and propeller thrust power are equal is analogous to the critical altitude of a turbocharged piston–prop and can be as high as 20,000   ft; the derated turboprop behaves like a turbocharged piston–prop but with a much lighter weight and less complexity.

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The Exxon Valdez Oil Spill

Stanley D. Rice , ... Jeffrey W. Short , in Long-term Ecological Change in the Northern Gulf of Alaska, 2007

Early Surveys (1989–1993)

The massive beach cleanup effort required comprehensive monitoring to allocate cleanup resources and to evaluate efficacy. The initial location and extent of oiled shorelines was documented by low-altitude, low-airspeed color videotaping ( Teal, 1991). These surveys defined the general area to be evaluated and monitored by ground-based "shoreline cleanup assessment teams" (SCAT). The SCAT teams included at minimum a geomorphologist, an ecologist, and an archaeologist, who comprehensively inspected the entire shoreline within the potential affected region during beach walks or from boats when beach access was impractical (Neff et al., 1995). The SCAT teams received similar training and employed uniform criteria and forms for recording observations (Owens, 1999). These observations provided the basis for segregating the shoreline into contiguous series of beach segments up to 2.5 km long that were bounded by readily identifiable landmarks in the field. Visual assessment of oiling intensity on each of these segments was carefully and consistently documented during the spring and summers of 1989 through 1992 (Fig. 5.11). Subsurface oil was also monitored after 1989 by excavation of thousands of pits to assess persistence.

Figure 5.11. Visual assessment of shoreline oiling, completed in fall 1989 by the Alaska Department of Environmental Conservation, after the first summer's cleanup efforts (Gundlach et al., 1990).

The SCAT survey results showed that surface oil on PWS beaches dispersed quickly, but corresponding results for oiled beaches outside the sound have not been reported. The cumulative length of visibly oiled beach segments inside PWS decreased from 783 km in 1989 to 10 km by 1992. The cumulative area of oiled beach was substantially less than what might be inferred from these results because the beaches were often not entirely coated by oil, especially after 1989. Very approximate estimates of oiled beach area may be calculated from the oiled shoreline lengths and percentages of intertidal area covered by oil given by Neff et al. (1995), which indicate a decline from 240 ha in 1989 to about 20 ha by 1990, and less than 1 ha by 1992. A total of 2.7 ha of oiled beach area was reported from the 1993 SCAT survey (Gibeaut and Piper, 1998). Subsurface oil appeared to decline more slowly than surface oil. The surface area of beaches contaminated by moderate to very heavy subsurface oiling exceeded 5 ha in 1991, and decreased by nearly 70% from 1991 to 1992 on beaches surveyed both years, but the total extent of subsurface oiling was not reported for either year (Neff et al., 1995). In 1993, the total beach area contaminated by subsurface oil was estimated as 3.4 ha, and this oil was suspected to be more firmly associated with the sediments, and hence less easily dispersed compared with prior years (Hayes et al., 1991; Gibeaut and Piper, 1998). The amount of oil remaining on PWS beaches in 1992 was estimated to be about 2% of the volume spilled, or about 817 m³ (Wolfe et al., 1994).

There were no long-term appreciable accumulations of subtidal oil. Although fluxes of oil were substantial to the shallow subtidal during the first year following the spill (much of it aided by scouring from the cleanup efforts), these did not usually result in accumulations because the subtidal sediments were continually resuspended and dispersed after initial oil deposition (Short et al., 1996). Some of the oil formed clay–oil flocs were of near-neutral buoyancy, promoting dispersion over a wide area of the northern Gulf of Alaska (Bragg and Yang, 1995). Oiled sediments did accumulate in the shallow subtidal adjacent to heavily oiled beaches when trapped in basins, for example, oil collected behind a terminal sill at the mouth of Northwest Bay. Subtidal sediment PAH concentrations on the order of 1000 ppb, and decreasing with depth, occurred in 1990. In succeeding years, these concentrations declined rapidly (O'Clair et al., 1996). A resurvey of subtidal sediments during 2001 failed to detect hydrocarbons from the Exxon Valdez near any of the five beaches that had been heavily oiled (Short et al., 2003).

The rapid loss of oil from PWS beaches determined by the SCAT surveys implied that the remaining oil would soon disappear (Boehm et al., 1995). It was suggested that clay–oil floc formation would lead to rapid natural removal of subsurface oil (Bragg and Yang, 1995), although the low concentrations of clay-sized sediments in the surface seawaters of PWS (≤1 ppm) likely constrained the significance of this process (Payne et al., 1989, 2003). Comparison of the volume of oil remaining on these beaches in fall 1992 with the estimated volume that originally beached in 1989 (∼42% of the spill volume, or ∼17,200 m³) (Wolfe et al., 1994), and assuming first-order dispersion kinetics during the 3.5-year interval, leads to an instantaneous dispersion rate of –0.87 yr⁻¹, or 58% less oil every year. If this were true, only about 60 m³ of oil would have remained by the fall of 1995. Unfortunately, such predictions were overly optimistic.

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Airplanes, Light

Darrol Stinton , in Encyclopedia of Physical Science and Technology (Third Edition), 2003

II.C.6 Aerobatic Airplanes

Aerobatics have come to the fore as an internationally competitive sport. Maneuvers have been developed that are far beyond anything previously conceived, and specialized airplanes are designed to fly them. Jose Luis de Aresti, a Spanish aerobatic pilot, devised a system of assessment in which aerobatic maneuvers are analyzed in families, from 1–9. Armed with the basic families it is possible to combine the maneuvers into sequences, the most demanding of which are flown in various international aerobatic contests. The Aresti system is only the foundation of contest rules, which can modify the Aresti concept considerably.

During aerobatic contests the object is to remain within a box of airspace of defined limits. This requirement results in the need for the following:

1.

High structural strength

2.

The ability to maneuver at comparatively low airspeeds (too much kinetic energy through too much speed makes it difficult to stay within the box)

3.

Crisp responses to effective controls and the ability to fly either way up or on one's side with ease

4.

Plenty of power per unit weight

5.

Ample wing area for the weight of the airplane, resulting in relatively low wing loading (i.e., weight per unit area)

6.

Quietness for obvious social reasons (although few airplanes achieve it)

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Measuring Wind and Indoor Air Motions

Dario Camuffo , in Microclimate for Cultural Heritage (Third Edition), 2019

Hot Wire and Hot Film Anemometers

The airstream cooling power is used in hot wire or hot film anemometry. The principle is to measure the current required keeping constant the temperature of an overheated wire. Alternatively, to measure the change of temperature of a wire heated at constant current. As the wire is very thin, the probe is very small and sensible, fast response, particularly suitable for microclimatic studies and to monitor low-speed air movements as well as short-term fluctuations. The sensor is very fragile and must be used with great care and is more suitable for indoor studies.

Miniature hot wire (or hot film) anemometry is appreciated for its simple use and low cost (Fig. 20.19). The size of the sensor is of the order of one or few millimetres in length and the diameter is of the order of 5 μm. The hot wire measures airspeeds above 10  cm   s^{−   1} and the time constant is of the order of 0.001   s. A lower threshold, i.e. 5   cm   s^{−   1}, is obtained with a nickel thin film deposited by sputtering on a spherical glass sensor, with a diameter of 3   mm (Dantec, 1996). The relatively larger mass increases the time constant to 0.08   s and the overheating generates a convective motion that interferes with the air movement at low airspeeds. This interference determines the lower limit of reliable measurements, which is around 3   cm   s^{−   1}.

Fig. 20.19. A hot wire anemometer (to the right) was used inside the Giotto Chapel, Padua to monitor the vertical flow (W) of warm air (red arrows) ascending over the incandescent lamps (L) and the cold return flow (cyan arrows) along the frescoed walls (to the left). The observed flow along the frescoes has speed fluctuating between 0 and 32   cm   s^{−   1}. From Camuffo and Schenal (1982) (see credits).

As a single wire responds to the velocity component perpendicular to it, a variety of probes exist, mounted either single or coupled orthogonally in a plane or three-dimensionally, suspended between the tips of a fork-like support, for detecting one, two or three components of the airstream. Some probes are inserted into a cylindrical shield (a tube) in order to measure the stream component along the cylinder axis. However, the edges of the tube disturb the flow field and generate departures, instability or even turbulence. It is convenient to remove this shield and insert the bare probe into the airstream, with the wire normal to the flow direction.

The physical principle (DISA, 1976; Doebelin, 1990; Tropea et al., 2007) is the thermal loss of an overheated wire operating as resistance sensor. The heat loss not only depends on the airspeed but also on a number of parameters such as air temperature and pressure. If only the airspeed changes, or the influence of the other parameters is compensated through other sensors and suitable electronic units, the output gives the airspeed. The characteristic transfer function is in first approximation composed of an exponential and a square root function, but the signal can be linearized, so that the processed output is simply proportional to the airspeed.

Two different circuits are available for these kinds of sensors: the constant-temperature and the constant-current anemometer. The constant-temperature type consists of a Wheatstone bridge and a servo amplifier and the sensor acts as the active arm of the bridge. The current through the wire is adjusted to keep the wire temperature constant and is a measure of the flow velocity. The constant-current type has the sensor powered by a constant current supplied by a generator having high internal resistance in order to be independent of any resistance changes in the bridge. The wire reaches an equilibrium temperature that is determined by the heat exchanges with the airstream. The heat generated is the product of the electrical resistance, the square of the current intensity and the wire temperature. Briefly, the airspeed is measured in terms of the electrical resistance. In practice, constant-temperature anemometers are preferable and are effectively popular for their easy use, fast response and low cost.

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Cloud Dynamics

Robert A. HouzeJr., in International Geophysics, 2014

8.10.2 Effects of Microbursts on Aircraft

Fujita's conceptual model has been used as an explanation for many aircraft accidents. A pilot must make quick critical adjustments in flying through the wind pattern of the microburst. For example, in taking off through a microburst (Figure 8.52 ), the aircraft experiences an increase of headwind as it accelerates down the runway. Then, the aircraft lifts off in the increasing headwind and begins to climb (position 1). Near position 2, it encounters the microburst downdraft, and climb performance is decreased. By position 3, the headwind is lost. Consequently, airspeed is decreased, and lift and climb performance are further reduced. Added to this is the increased downdraft at the microburst center. By position 4, all available energy is needed to maintain flight, as the tailwind continues to increase. However, there is no source on which the aircraft can draw to increase its potential energy (climb). A large aircraft is typically configured ("trimmed") such that thrust, drag, lift, and weight are all in equilibrium. Thus, no pilot input is required for the aircraft to maintain a set trajectory. Since the airspeed is below the trim airspeed (decreasing lift and drag) at position 4, the airplane system will automatically respond so as to regain the equilibrium condition by pitching the nose downward. In the illustration, the pilot intervenes and compensates for this effect. Should the pilot not fully compensate, a more radical descent could occur. The descent rate continues to increase as the airplane passes through position 5. Depending on the strength of the event, encounter altitude, aircraft performance margin, and how quickly the pilot recognizes and reacts to the hazard, the high descent rate may be impossible to arrest before the plane crashes into the ground. ⁴⁷ Similar difficulties occur when attempting to land in a downburst.

Figure 8.52. Idealization of an aircraft taking off through a microburst. After lifting off in an increasing headwind at 1, the aircraft begins to lose the headwind and enter the downdraft at 2, experiences stronger headwind loss at 3 with subsequent arrest of climb, begins descending in increasing tailwind at 4, and experiences an accelerating descent rate through 5.

From Elmore et al. (1986). Republished with permission of the American Meteorological Society.

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