DE2053479A1 - Device for measuring the angle of attack of aircraft - Google Patents

Description

Patent Attorneys 3 O. ORf. 13/0

Dipl. Ina. C. Wallach
Dipl.! Ng.-G. a cook
Dr.T. Haibach

8 Munich 2.

Kaufing8, str. 8, T8l.24ö275 ^IZ

Rand Corporate! Ou, Hew'vTerk / tJSA

forrichfcung zur Messimg d ^ ti angle of attack of I.

Luf 1 1 & hx '£ eyes

You can refer to G'töuersystemo fUi 'Luftfahro and, in particular, Vorridtifcungeii for measuring the angle of the air eye from measurements of the litigstiohsau "~ and Hoöhaehaeti s accelerating in help Möiiijußgen the vertical and forward flight speed.? dltiie measurement of the an-sfceX-lwinfceis w.frd so ab- vXtn ^ i -kill them from the a-evodyneunisöhs «. Properties of -

vf & iir ^j Keugiis vnabhäLglg iat * and thus on a multitude

"'tn Flugae'.Itm" drafts sxwv.jnfzm * without the need for extensive and costly tests "- A whiter" aspect of the "present invention includes the need for precise measurements of the aircraft acceleration along the plug orbit direction and of the instantaneous vertical velocity, both of which are free from the deleterious effects of vertical spiral righting control effects.

1098 1 9/0 U6 ./.

The angle of attack of an aircraft is a fundamental flight parameter and is used in guidance and control of an aircraft extremely useful. In the past, the angle of attack was in particular used for guidance and control

augmentation is not u? nfästend used an aircraft, before all thing @ n due d <sr difficulties sin accurate, störungsfr @ ies, AESN angle signal representing over a wide Bereieh of Flu _(5betriefo to get it and the Flugbedingungsn. So far the angle of attack has been measured with the help of probes or flags protruding from the outer surface, which more or less wet the properties of air flow. which had to be filtered or steamed in the Hohfesu Srad © in order to obtain usable control information. Furthermore, such external probes are exposed to accidental damage from dust, ice, treatment on the ground, etc., which causes failure and maintenance problems. rrneprobl © me arise, the probes and wings to make extremely uncomfortable »

Although certain successes have been achieved in measuring the position angle. Without the use of flags or probes, the implementations required long-term longitudinal (and transverse inclination) positions yo ^ srtikalkreisein. For example, s © i on the patent ......... [patent application 19 03 151 ö? 5 of the same. Anmalderin pointed out. The principal disadvantage of data from vertical gyroscopes is that they are subject to inaccuracies due to righting and drifting properties during flights under (accelerated) long-term flight movements. Still iat the reference point? the horizontal plane of the earth instead of the * ⁵ longitudinal plane of the aircraft, which causes errors during U-turns
™ * οα. * Ι 1098 19 / Π U6 ',

BADQftlßlNAL

According to one aim of the invention, a device for measuring the angle of attack of an aircraft is created, the two with their sensitive axes aligned at right angles in the longitudinal plane of the aircraft in the aircraft © Besnhlö.miguiigsmeßralttel, the Jewaila on corresponding roohtwiJslcli ^ ri components of the aircraft in der Lüngsefctne @ Ι: -π! .ί *: € αά € νη address entire accelerations, in this way to deliver Ul m rlretimmende signals, further L! .ffcdafceü "Fü ^; -i ^, äim: U '^J;<rl and Deduction of measurements of the height of the falirtvjlr / l © 3 and the force of gravity relative to the air vehicle gauging means including means for reading one of the readings

söl.O.evaiigung dss FluggeseSiw: indigk «i ':.% hw I uf!" vehicle-representing signals and on this © ii: lgii.r ^:! .em speaking means for the management of your Mossui; *> g des Ati ^' se-XiAiii &gls'. of the aircraft on / travels.

lnig> avenging an andei'ßii target dov K ..-; f; MiclvvT-f; will create a rectilinear for measuring the iinstellw.Vnife.is elrits aircraft !! ©, the svre: l in the hui '^ ff _l hr ^: s; - eug with their sensitive axes at right angles in If :? M'.Tgjsebene des L -ftfalirzeugfes attached.-3 Beijcb.Accelerometer, each speaking to uir * 1 · real square components of the Gesamtbssohleunigung to ^ r ^ eugung of ai 'fi € -n anspreehsn corresponding signals further L ".ftditsn-FUlilermittel to Irzeugang ago with äem Flugbahnilrrtr-J, TIAS aircraft vno de?' Beeohl.eunigung the Fluisgfiiiehvrir-Mfjiceitv of Luftfaiirzrages along its trajectory ttfc öreinstimmenden Signaisn einsohlisSencle-Kc ^ tel VTI ¹ l din signals of the Bt? 3 ^ hlewnigungsmes) 8 (IR and i-uftdatejifUh-Ua * loisprechende means for deriving a au »a Anstßllwi.r.'rl > l & n aircraft

1 Π 9 8 1 9 / Π U 6 ^ΒΛ °

According to a further object of the invention is a Device for measuring the acceleration of an aircraft created along its trajectory that Including accelerometers and gyroscopes Inertia means for deriving a first * signal that corresponds to the acceleration of the aircraft depending on its flight path direction, flight data sensor means to derive a second, with des' 0 @ so speed des Aircraft along its trajectory direction matching signal, fade-out (washout! Filter medium and Means for supplying the two signals. the fade out » (washout) filter media, due to the changing Values of the speed signal derived from the flight data are allowed through and constant state values of the inertia acceleration rank signal a> fog © folookt.

According to a preferred match of the Invention is the aircraft angle of attack from measurements; those that act on the aircraft and are linearly connected to the longitudinal and vertical axes of the aircraft Accelerometers sensed BeoehlQunungen derived « The outputs of these accelerometers are given with the true airspeed uric the altitude change value comprehensive air data information and Kuczseit vertical gyro data combined; to provide an accurate measure of the angle of attack over a wide spectrum of frequencies. Signal filter and attsblend (washout) techniques become more advantageous Used in a way to reduce the effects of sensor mismatch and tolerances caused by vertical gyro alignment properties. the effects of pitch movements on the static pitot source and the Eliminate the effects of wind thrust and gust turbuence. To implement the angle of attack, an exact measurement of the

1 0 9 8 1 _{9/0 U 6 BAD} OWGINAL ./.

Acceleration along the flight path of the aircraft necessary. This measurement is in turn in an unambiguous way using the accelerometer outputs together with the derived measurement of the angle of attack and the true airspeed (or Maeh number) and using Rückführung®- urü Äusblend- (washout) Get techniques "

The measurement of the angle of attack provided by the converter (computer) according to the invention is accurate under all flight operating conditions of a typical transport aircraft and does not depend on the aerodynamic properties of a particular aircraft in the design and is therefore of a large number anwandten plug Sell configurations without extensive Testflugeiohung. Sasätzlieh into account the implementation of all the airframe acting accelerations, including the ₀ that occur when Kurvenflng "ie ^ © ntrifugiüb @ sohX @ unigung with a short turnaround and the Zttitrifugalbeschleunigung because d © r" turn-g © schwinclisksit ύ © η flight path.

Xn practice, the angle of attack of the aircraft to measurement of dtn mutually perpendicular Besohleu- is | nigungsraessern, the Rlehtut da® Fatirtwludes and the change in the Luftfahrzeuggsschwlndigkeit ^ which was here as Fluggeechwlndlglceitis-BesohleimiguHg b <&% ®iohnttb and a determination of the momentary Luffefalirzeugiage, as it is derived from the direction of the visual power system teattmi.

The here for the implementation of the various parameters (angle of attack χ, flight path direction ^ angi vertical speed, Acceleration along the trajectory) used general

10B819 / 0U8 ./.

Technique involves the generation of a variety of measurements the same or similar parameters from generally different sources, each of which has undesired

• - in-

Transmission frequency response properties wY:, & v a multitude of operating or flight conditions tn avX, and the processing of these signals via network with a transmission frequency response in © in such a way that undesirable properties are eliminated and the dynamic sensitivity and accuracy to be striven for is retained Thus a large number of measurements of the flight path acceleration and - ¥ © rtikalgeschwind! Eg-P ity of the aircraft are derived, -ion clensn, one based on measurements derived from inertia means and the other on the outputs of an air data computer processed via filter and fade-out (washout) networks, ale derive a resulting output, which has the desired frequency components of the ^ aremeter for all flight conditions.

The invention is illustrated in the following with reference to a in the drawing preferred embodiment shown in more detail explained. Show in the drawing

^ Fig. 1 diagrams showing the focus on disi of the air- and 2 vehicles acting as compensation vectors representemj

Flg. 3 is a diagram showing the focus of the

Aircraft acting ICräfbe represents, their Imbalance those shown in d @ n FIGS. 1 and 2 Causes accelerations;

Fig. 4, block diagrams showing the preferred embodiment X 5 and 6 examples of the comput © rtachniken for derivation In each case the angle of attack !, the acceleration

,, " 1098 19/0146

along the trajectory and trajectory vortex;

Pig. 7 an integrated block diagram including the sub-computer arrangement.

According to Figs. 1, 2 and? three primary elongation forces act on the center of gravity of an aircraft: the apparent weight (where 0 is the angle of inclination and Jf, μ is the effect of the centrifugal force on a short one

COS Jv M

U-turn)., The aerodynamic force F. (which ™ is the counteraction of the aircraft on the airflow V, which in turn is the resultant of the ¹ perpendicular forces of lift L and drag D) and the power of the engines or thrust T ( which acts along a direction which is defined in relation to the aircraft and which can be assumed here for descriptive purposes as lying along a line which includes an angle dg with the XX or longitudinal axis of the aircraft). Any imbalance in these forces results in an acceleration of the aircraft's center of gravity, depending on the magnitude and direction of these forces. The characters X and 2 show the composition of the ä longitudinal plane acceleration vector Vr of gravity of the aircraft at a short turnaround and under the influence of, for example an increase of the thrust vector ΐ of Figure 5. These components include g is the earth Sehwerkraft. the centrifugal acceleration of a short U-turn g tan 0 t the centrifugal acceleration due to a curved trajectory in the longitudinal plane V ^ y ^ (at right angles to the trajectory acceleration} and the acceleration along the trajectory V ₁ , «

^{109819 / Q1ie}

The angles of the various vectors relative to the horizontal are as follows: the angle of the XX-axis relative to the horizontal is _w (here referred to as the longitudinal inclination of the wing), the angle that the direction of the trajectory of the wing includes with the horizontal If the flight path _{angle $ * w} is the direction of the movement of the airstream V_, and the angle between the direction of the airstream and the XX axis of the aircraft is the angle of attack <h of the fuselage es. The sum of d * and ^ T _w is therefore G ^ and can be given by the well-known equation

can be printed out.

The total acceleration Vj. can be right-angled into two Directions a and a_ are resolved in the aviation are firmly arranged and here each as along the X-X or longitudinal axis and Z-Z, or vertical axis of the aircraft are shown lying down. The primary accelerations described above in the 'Lfingo plane or symmetry plane of the aircraft contribute to the resulting acceleration Vj. and thus to their right-angled components a "and a" according to the following equations

Λ Ζ

^ jj sin (jT _w + <t) + V _T ^ sin * + V _T cos * (2)

^a z " ^cos

These equations can be solved simultaneously to get a to eliminate the changeable. When the variable Vm is eliminated, the equation is given

st a.

ooe * (. _B - ^ f) ₊ a ^ Ca _x -

■■■ - ■ 1 098 1 9 / 0U6

If the variable Vfjfc. is eliminated, results see the equation!

g cos % j ^ ■ g sin O _n

& · Ar% t. * I j>; \ {ö ^ * ■■ ■ ι null J {^ 1

* Coj i / W \ a _x oos 0 / ^kD)

In transport planes, the millions are usually small and their sunanos, for example, hardly exceed 20 °. Therefore, it is fSiSglich to use approximations for small angles for the sine and Cosinusfutiktionen these expressions, preceded the equations (4) and (5) can be written as follows köiinonj d

, g cos © _M

S ^COS

Equation (7) is used for ais Mcich: ia.iai & y.ving according to the present teachings as follows ua ^ Q

^a x

g cos Q _w g J

^a 25 COS '0 ~ ⁺ "^ SS-TfJ? -"

The Flg. However, 4 aeigt a Oine MesBUüig the pitch angle according to the relationship of equation (8) dissipating Ana log computer arrangement "It is understood that the reactions just as well by iiinci general purpose digital Coraputer can be executed / DBEI the data inputs in an appropriate manner were digitized. Me in Fig. 4 blocks shown on the left represent the data sources for the

109819/0146 bad owoinal

angle computers and generally only include the orthogonal accelerometers, a ventilation data computer and a vertical gyro. In Pig. 4, the sources include the output of a Z-axis accelerometer 10, three-twist pitch and roll outputs from a vertical gyro 11, a fixed voltage source 12, the output of an X-axis obsession meter , a Source 14 for data proportional to the acceleration along the plug path and © in © Qu © ll © 15 for data proportional to the flight path angle of the wing. The um- '

^ Settlements of the aircraft acceleration along the flight path or the flight path angle are shown in FIGS. 5 and 6 and are discussed in detail below. The expressions a "and a_ are derived from two linear accelerometers 13 and 10, respectively, attached to the x and z axes of the aircraft, as indicated in FIGS. 1 and 2 of any conventional design and may be of the type shown in U.S. Patent 5,190,128, for example. The accelerometers 10 and 13 can have any angular orientation relative to the aircraft as long as they. are at right angles to each other and in the plane of symmetry

ψ _{K of} the aircraft, ie. the XZ plane »A conventional vertical gyro 11 is used to generate the longitudinal inclination expression ©“ and the transverse inclination expression 0 with the help of conventional three-wire tapers of the Synchroj art, which are attached to the longitudinal inclination and transverse inclination axes of the gyro. Such gyroscopes can be of the general type shown, for example, in U.S. Patent 2,945,380. The three-wire outputs of the vertical gyro are modified in a suitable manner by a signal processing device 16.

109819 / 0H6 ' ^Λ

fied that Scott-T transformers can have voltages that are proportional to the cosine of the pitch angle, the sine of the pitch angle, and the cosine of the roll angle. Electronic dividers 17, 18, 29 and 21 can be of any type and may include, for example, operational amplifiers and pulse width modulation techniques known to those skilled in the art of analog conversion. These devices are used to generate voltage outputs proportional to the quotient of the / variable inputs.

The vertical accelerometer 10 is arranged to deliver an electrical signal proportional to the vertical component a. _{% of} the total acceleration acting on the aircraft, this signal being fed as an input to a summarizing means such as, for example, the amplifier 20. The expressions cos θ ι

_G and * , which influence the pitch

cos 0 cos 0

and bank angle of the aircraft on the entire vertical acceleration component are derived from the vertical gyro 21. The signal processing device 16 derives cos Gq and cos 0 from the Krelsel longitudinal and bank synchros, where cos Q "is divided by cos 0 in divider 17 and I cos 0 as a divisor for the divider that responds to a constant voltage supplied by source 12 The output of the divider 17 suitably modified by the gravitational constant g and the output of the divider 18 are fed as the other inputs to the summing amplifier 20 which provides an output proportional to the equation

cos J3 ⁺ cos ' ß /

109B19 / 0H6 BADOtIQ ₁ NAL

which forms the denominator of equation (8) and which the Vertical component of the aerospace acceleration due to of the aircraft angle of attack, which represents the effects of the aircraft pitch and lateral angle position and gravity balances »

The longitudinal accelerator 10 is arranged in such a way that it supplies an electrical signal which is proportional to the longitudinal direction of the acceleration acting on the aircraft, this signal being fed to an input of another summing amplifier. The effect of the aircraft speed change component, i.e. the aircraft plug speed acceleration on the longitudinal acceleration, is determined by the computer, which has a corresponding input of the summing amplifier 22 connected to it Signal delivers. The effect of the aircraft Plugbahnwinkelss to the longitudinal direction BöschleunigungsmeEHor 3.0 is partially determined by the JFY Computers 13 * nal provides dor elii corresponding Sig to the divider circuit 19 a at its output - orseugt signal proportional to this output with the Sehwerkraftskonstanten g is modified before ei * is fed to amplifier 22 as a further input. The output of the summing amplifier 22 is therefore a signal which forms the numerator of equation (8) and which forms the longitudinal component of the aircraft acceleration due to the aircraft angle of attack, which is compensated for the effects of plug speed changes and flight path angle of the aircraft and gravity .

The counting or. Denominator outputs of amplifiers 22 resp. are fed to the divider 21 to at its output the

108819 / 0U6 * ^A

BAO ORlQINAL

desired measurement of the aircraft angle of attack CK. The angle of attack output of the divider 21 can be separated in a suitable manner with the aid of the isolation amplifier 22 and its gain controlled. This gives you, from primary measurements of the aircraft accelerations in parallel on the longitudinal and vertical axes of the aircraft and measurements of the direction of the airflow, the flight path angle information derived from the flight data and the flight speed acceleration of the aircraft, all these measurements being available from devices within the aircraft structure , together with the information regarding \ to the angular direction of the measured g-forces a ge ~ precise measurement of the angle of attack. It should be noted that the longitudinal inclination output of the vertical gyro 11 does not exactly represent the longitudinal inclination position of the wing (denoted by 0 _{w in this exemplary embodiment)} is, instead of on the Queraehoe of the aircraft. In Pig. 4 the longitudinal inclination output is labeled 0 _Q to distinguish it from the measurement Q ^. It may of course be shown. That is ,,, for normal loading a operating range of pitch and roll positions of transport aircraft cos 9 with sufficient Oenauigkeit equal eos O ^, so that no significant effect on the desired (S * measurement However, with respect to sin Q ^ , it is necessary to process the gyro pitch signal by dividing it by cos 0 in order to obtain the desired accuracy.

Pig. Fig. 5 shows an embodiment according to the invention to generate a signal that is exactly proportional to the acceleration of the aircraft along its plug path

109819 / 0U6 BAD ORIGINAL.

and thus all the ^ data source 14 according to FIG. 4 can be used. The introduction is based on the derivation of redundant measurements of this acceleration expression; one is based on inertia elements, the other on air data. The two measurements are processed with the help of networks with a transmission frequency response in such a way "that undesired prequence characteristics are eliminated and an accurate dynamic response characteristic is then maintained. The mathematical development for the V _T computer is summarized in the following relation"

(9)

In the oleion (9) a , V «, represents the natural speed output of an air data computer and ^ _{1 is} the time constant of an electronic washout

ff * _m ο

So holding with a transfer function ^u 1 , where-

_d T ₁ -S + 1 at S the time differential operator ^ 1st. The Vm expression on the left side of the equation (9) is determined in accordance with the relationships shown in the equations (7) or (8) from accelerometer and gyro data. The blocks shown on the left in FIG. 5 comprise the data sources for the flight path acceleration computer. It can be understood that the sources 10, 11 and 13 and the signal processing device 16 with the corresponding blocks shown in FIG. 4 are identical. The <K -computer data source 22 is to be understood as the entire computer shown in FIG. 4, as can be seen from FIG. The air data computer 24 is completely conventional and forms the measurements d <r * s statl-

109819 / 0U6

BATH ORIGINAL

look around for pressure, dynamic pressure and air temperature in the airspeed V _T proportional electrical values. Such an air data filer can be of the type shown in U.S. Patent 2,936,13k. The divider circuits 25 and 27 serve Iu connection with the buffer amplifier 26 to the avis Verarbeitmiß 0 and C ₀ sin outputs of SignalverarbeitungsgerilVs 16, provide a measurement _of S-sin Qy _au what lf equal ^{g. in 0 is} O.

Oil divider circuit 28 is used in conjunction with the summing amplifier 29 and the multiplier: attitude 30 to the delivery

tion of an output equal to df f., - ', Qf . ^

ι 'οθίΠ2Γ /

The multipliers can also use operational amplifiers and pulse width modulation technology, as was done with the plate networks. The Icrtzfce expression is algebraically summed up in the summing amplifier 31 with the output of the longer!

fen »This makes the V ₁ ^ expression on the“ left side

of equation (9). The expression .-? (Air data) 1 ^ 5 is passed through a fade-out (washout) pilot J2 with a time constant T ₁ . The time skill; T _{1 of} the dimming circuit 32 is f) checked that it is long enough "to filter out irregularities in the aircraft's pitot-ttatitsclien system due to pitch gusts, gusts of wind, wind gusts, etc., but is short enough to prevent from A value of 10 seconds is a typical value that fulfills these requirements in transport aircraft.

Fig. 6 shows a preferred exemplary embodiment

1098 19 / OUe

Computer for generating a signal proportional to the flight path angle of the wing, this expression being required for the conversion of <*> and thus representing the source 15 for, r _w according to FIG. The introduction is based on the derivation of redundant measurements of the vertical speed of the aircraft; one is based on elements of inertia and the other is based on air data. Here again, as in the case of the trajectory acceleration term, the two measurements are processed through a network having a transmission frequency response such that undesirable properties are suppressed and yet uniformity, accuracy and excellent dynamic response characteristics are obtained. The mathematical development is summarized in the relationships (10j bis as follows:

h (10)

From Figs. 1 and 2:

h - (a _x sin 0 _w + a ₂ cos

cos Ou "= cos Qq (12)

sin 0 _w . sin O ₀ cos β

Substituting these expressions into equation (11) and multiplying both sides by oos 0 we get:

h m 8L _x sin O ₀ + a _z cos O ^ cos 0 - g (lj5 ^c }

109819 / 0U6

SAD OAiGINAL

In equation (10), η represents the vertical speed and is an output available from a conventional air data onputer. “Fg is the time constant of an electrical system with a transfer function ” where S is the time differential operator ^.

The expressions introduced into equation (11) for h represent the projection of the entire longitudinal plane acceleration vector V ^ of FIG. 1 onto the true vertical axis of FIG. The equations (12) and (13) show the conversion of the gyro _{pitch angle O Q} into the wing pitch angle O ^. Equation (14) represents the usual expression for small angles for the flight path angle in terms of vertical airspeed and airspeed, while equation (15) converts the flight path angle into the vertical plane so that it corresponds to the aircraft longitudinal plane.

FIG. 6 shows a preferred analog computer solving equations (10) through (15), and uses techniques similar to those used for the angle of attack computer of FIG. 4 and the flight path acceleration computer of FIG. 5 have been described. It should be noted that the signal sources 24, 10, 13, 12 and 11 express the same values as those for the conversion of 6 < and V _T with the addition of the vertical velocity expression from the air data computer and the inclusion of the time constant _{T 2} In in the acceleration. The multiplier 23 * the isolating amplifier 34 and the multiplier 35 supply the a _z * oos Q ^ * oos expression, while the multiplier 36 supplies the * _χ > sin Qg expression and a voltage source 12 supplies the constant g expression, all of them Equation (13 *) · These expressions

109819 / 0H6

are algebraically combined in amplifier 37 to provide the h-term of the solution (10). This expression is fed together with the barometric Vertlkalgesohwlndlgkeltsausdruck ή from the air data computer, ie the output of the amplifier 37 to the Auablend (washout) PiIter 38, which has the transfer function shown, where If _{2 is} on the order of 4 seconds, which is long enough It is designed to fill in the irregularities of the static source for the aerial data, but is clearly short enough not to be affected by the uprightness of the vertical gyro. The output of the filter circuit 38 is thus an interference-free, precise signal proportional to the vertical speed of the aircraft and can be used both for other aircraft systems that require this expression and for the representation of this important parameter in the cockpit. This signal has excellent dynamic properties without being subject to the disturbances and irregularities of the static source, which are associated with pure air data derivatives. Many barometric inertial introductions were adversely affected by brief U-turns that produced spurious and inaccurate results. The implementation of h described here is not burdened with this effect, because compensation is provided by measurements of the angle of inclination and because the X-axis acceleration is included as an addition to the Z-axis acceleration, which is normally used for this purpose in earlier exemplary embodiments were used by barometric inertial indicators for vertical speed.

The solution of the equation (15) is given by the parts set 39, the isolation amplifier 40 and the divider 41 carried out,

109819 / 0U6

«AD ORlQlNAl.

the latter being a bank angle compensation the aircraft results.

In the above, the implementation of the angle of attack ο ** ¹ , the trajectory acceleration V ^ and the plug path angle <f _{w of the} wing was shown separately for purposes of clarity. It will be understood, however, that in practice these implementations are incorporated into a single computer to provide a translation of the value of the aircraft's angle of attack with the auxiliary outputs of V _T , f, and ή that are useful for use in other such measurements required systems are provided. έ

Big. 7 represents the summarized block diagram of the whole computers according to the invention, with corresponding reference characters being used to denote the corresponding blocks in Figs. 4, 5 and 6 to be designated. It can be accepted that the designations attached to the block diagrams clearly show the conversion function and therefore no detailed description of these punctures is required.

Patent claims:

• A

109819/01 46

Claims (1)

Claims I.) Device for measuring the angle of attack of an aircraft, characterized in that two accelerometer means (lo, 13) which are fastened in the aircraft with their sensitive axes aligned at right angles in the longitudinal plane of the aircraft and which each point to the corresponding right-angled components of the address the aircraft acting in the longitudinal direction plane to generate these corresponding signals, also air data sensor means (24) and accelerometer means (10, 13) including means for deriving measurements of the direction of the wind and the direction of gravity relative to the aircraft axes, which Air data sensor means (24) including means for deriving a signal representing the flight speed acceleration (V ^) of the aircraft and means for deriving a measurement of the angle of attack ( (K) of the aircraft responsive to these signals are provided * 2. Device naoh claim 1, characterized in that one of the accelerometer means (13) is attached to the longitudinal axis of the aircraft and is responsive to accelerations of the aircraft along this axis, and that the other (10) of the acceleration measuring means attached to the vertical axis of the aircraft and on the Responds to accelerations of the aircraft along this axis. 1 0 9 8 1 9/0 U 6 BATH ORIGINAL 3 * Device according to claim 1 or 2, characterized in that further means (16, 17, 13, 19) are provided, which respond to the direction of the airstream and the measurement of the direction of gravity, to counteract the effects of the accelerometer signal To compensate for the wind and the force of gravity " 4. Device according to one of the preceding claims, characterized in that the Means for deriving a measurement of the direction of the headwind include flight speed and altitude change value sensing means (15) for producing a measurement of the aircraft flight path angle. 5. Apparatus according to claim 4, characterized in that the flight path angle measurement further includes means (21) responsive to at least one of the acceleration measuring means 6. Apparatus according to claim 4, characterized in that the Flugbahnwinkelmes & ung furthermore has means (21, 22) responsive to both accelerator masses. Device according to claim J, characterized in that the compensation means furthermore gyro means (11) responsive to the position of the aircraft for generating measurements of changes in the effect of the acceleration due to gravity in accordance with changes in aircraft attitude. 8. Device according to one of the preceding claims, characterized in that the BATH 109819 / OUe the airspeed acceleration deriving means comprises air data means (24) responsive to the airspeed of the aircraft and means responsive to one of the acceleration measuring means. Apparatus according to claim 8, characterized in that the flight speed acceleration deriving means includes means responsive to both accelerometer means. 10. Device for measuring the angle of attack of an aircraft «characterized in that that two in the aircraft old their sensitive axes Accelerometer means (10 «13)» fixed vertically aligned in the longitudinal plane of the aircraft each on the corresponding right-angled components of the aircraft in the longitudinal direction acting overall accelerations to generate these respond to corresponding signals «air data sensor means (24) including means for generating the flight path angle of the aircraft and the flight speed acceleration of the aircraft along its flight path corresponding signals and the acceleration meter and computer means responsive to air data sensor signals are provided for deriving a signal proportional to the angle of attack of the aircraft. Π. Device for measuring acceleration of an aircraft along its trajectory, characterized in that acceleration eater (10, 13) and means for diverting including circle means (11) 109819 / 0U6 iAD CWIQlHAL a first signal corresponding to the acceleration of the aircraft along its flight path, flight data fixing means (24) for deriving a second signal corresponding to the speed of the aircraft along its flight path, washout filter means (J2) and means (31) which feed these two signals to the masking filter means (32), changing values of the speed signal derived from the air data being allowed through and values of the broom acceleration signal of the inertia system being blocked in the steady state. 12. Apparatus according to claim 9, characterized in that the acceleration rake include two accelerometers (IG, 13), which are fastened at right angles in the aircraft and on the components of the overall acceleration of the aircraft Address in its longitudinal plane. 13. The device according to claim 9, characterized in that further means (23) are provided which respond to the angle of attack of the aircraft for modifying the acceleration signal in accordance with the angle of attack. | 109819 / 0U6 BATH