Take-off and Landing Devices[36]

General remarks As has been mentioned in Sec. 8-1, the take-off and landing devices on the wing serve to increase the maximum lift coefficient. A great variety of arrangements are utilized to increase the maximum lift. The older kinds of take-off and landing devices consist of flaps and balance tabs attached to the wing trailing edge or the wing nose (Fig. 8-3). More recently, devices have frequently been used that increase the lift through boundary-layer control by suction or ejection. A brief account of this method has been given in Sec. 2-5-3. A comprehensive survey of the various methods for the increase in maximum lift is included in Lachmann [28].

The effect of take-off and landing devices on the lift characteristic c^(a) of a

wing is presented schematically in Fig. 8-16. Curve 1 gives the values without flap deflection. Curve la shows the increase in the coefficient cLmax by boundary-layer control at the wing nose. Curve 2 gives the values with flap deflection, and curve 2а again the increased values of cLmax through boundary-layer control at the nose. Curves 3 and За give the corresponding data when, in addition, the boundary layer at the flap nose is controlled as well. The summary report about theoretical and experimental studies on boundary-layer control by Carriere et al. [8] should be mentioned. Earlier, a paper on the properties of flap wings was given by Young [55].

Flaps The simplest method of increasing C£,max is the deflection of a cambered flap as shown in Fig. 8-1 la. This effect is obtained because the flap deflection increases the effective camber of the wing, resulting in a lift augmentation that may be considerable. As an example, Fig. 8-1 la shows cL against the angle of attack for several flap deflections. The increase in С£тах depends on the flap chord ratio Xf, the highest values are usually obtained for Xf = 0.20-0.25 [7].

A quite simple landing device in terms of design is the split flap as shown in Fig. 8-3e. This is a flat plate lying against the lower side of the wing and turning about its forward edge. The lift curves cL(a) of Fig. 8-17b for several flap angles rjf are similar to those of the cambered flap (compare Fig. 8-17a). The effectiveness of the split flap is, according to Gruschwitz and Schrenk [19], due not only to an increased camber but also to a reduction of the static pressure on the suction side of the profile. In Fig. 8-18, the pressure distribution is shown for a wing with

Figure 8-16 Effect of flap deflection and boundary-layer control on the lift of a flap wing (schematic). Explana­tions in the text.

Figure 8-18 Pressure distribution on a wing with deflected split flap, from [19]. Curve 1, without flap deflection. Curve 2, with flap deflection.

deflected split flap. Because of the flow around the sharp trailing edge of the deflected plate, a strong low-pressure range is formed in the wake of the flap, having an effect up to the upper side of the wing.

Basically, the cLmax value increases with Reynolds number. In Fig. 8-19, the results on the effect of the Reynolds number on the value of cLmax are given, both for a wing without a flap and one with a 60° deflection of a split flap. Young [54] reports on the separation characteristics of flap wings. Flaps extending over only a portion of the wing span will be treated in Sec. 8-3.

The effectiveness of the simple cambered flap is limited by the flow separation occurring at large deflection щ right behind the flap nose. By boundary-layer control at the station of greatest danger. of separation, the lift-increasing effect of the cambered flap can be improved, as shown schematically in Fig. 8-16. Boundary-layer control by suction or ejection requires a considerable design and construction effort and will be discussed later in more detail. On the other hand, the slotted flap as shown in Fig. 8-3&, first suggested by Betz [6] and by Lachmann [27], represents a simple design for natural boundary-layer control. The slotted flap functions in such a way that the air, flowing through the slot from the lower to the upper side, carries the boundary layer, formed on the wing, into the free flow before separation can occur. Starting at the flap nose, a new boundary layer forms that can again grow over a larger distance before separation.

The maximum lift coefficient cLmax depends on the separation processes at the main wing in front of the flap as discussed in detail in Sec. 2-5-1. The most unfavorable flow conditions occur shortly behind the profile nose of the wing and at large angles of attack, а « a(c£,max). Here, the pressure increase that follows the

Figure 8-19 Change of maximum lift coeffi­cient with Reynolds number for a wing without and with a split flap. Flap chord ratio Kf =0.20, flap angle 17f= 60°, from [7].

suction peak usually leads to boundary-layer separation at the wing leading edge (see Fig. 2-44). By boundary-layer control, similar to that of the trailing-edge flap, separation can be shifted to larger angles of attack. The extension of the linear range of the <?x(a) curve of Fig. 8-16 leads to a considerable additional lift gain.

Another effective arrangement for the increase of the maximum lift is the slat (flap before the wing leading edge) as shown in Fig. 8-3/, whose characteristics have already been discussed in Sec. 2-5-3. A polar curve of it is given in Fig. 2-53. Figure

8- 20 shows the lift coefficient plotted against the angle of attack for a wing without and with a slat. In agreement with profile theory, the slat does not generate a noticeable change of the profile camber, because this would cause a parallel shift of the cL(ct) curves without and with slat. Because of natural boundary-layer control, the maximum lift coefficient of a wing with a slat is reached at very large angles of attack.

An effect similar to that of the slat is produced by the so-called nose flap, first proposed by Kruger [44]. Here, the increase of a(cLmax) results from a different effect, namely, the shape of the profile nose, responsible for the separation process, which is changed favorably by the flap deflection (see also Fig. 2-44).

In addition to the conventional landing devices on the trailing edge discussed so far, the double-section wing as shown in Fig. 8-3c and the Fowler flap as shown in Fig. 8-3<i must be mentioned. The former is a simpler design of the slotted flap. The latter consists of a flap that is driven out rearward and deflected. A simultaneous camber and area increase is thus accomplished.

Frequently, several landing devices are utilized in combination to establish a maximum lift that is as large as possible. As an example, Fig. 8-21 gives the lift coefficient of the profile Go 819 with a slat and a double-section flap against the angle of attack. The favorable effect on the boundary layer of the flow through the slot between the slat and the main wing is clearly indicated by comparison with the measurement when the nose slot is closed. In this latter case, the cLmayi values for

Figure 8-20 Lift coefficient Ci(a) of a wing with slat, from [48]. Profile Clark Y, Reynolds number Re — 6 • 10s. Curve 1, without slat. Curve 2, with slat.

Figure 8-21 Lift coefficient c^(q) for the profile Go 819 slat and double-slot flap, from Wuest [53].

all measured flap angles are lower by Acimax ~ 0.6; also, the flow separation leads to a larger lift drop than for the open nose slot.

Comprehensive data on the maximum lift coefficient of wings with and without landing devices are given in [32, 33, 46].

Suction In an effort to increase further the maximum lift of wings, suction was studied quite early (see Betz [4]).

The suction intensity is defined by a dimensionless suction coefficient as


Here Q is the volume removed per unit time, A is the wing area, and Ux is the incident flow velocity. The maximum lift can be increased considerably by slot suction. Comprehensive tests on this method were conducted by Schrenk [4]. The most effective method, particularly for thick profiles, was found to be slot suction with a flap wing. Lift coefficients up to about cL = 4 may be obtained, as shown in Fig. 8-22 for a thick profile with flap and suction. Here the coefficients of suction are about Cq = 0.01-0.03 and the suction pressures cp =(p —p^/q™ — ~2 to —4, where Q stands for the total flow volume removed, p for the pressure in the suction slot, and qm = (p00/2)£/’« for the dynamic pressure of the incident flow. The effect of suction lies in its keeping the flow essentially attached to the flap. The greatest danger of separation is near the flap nose. If the decelerated boundary layer at this

Figure 8-22 Lift coefficients of flap wings with slot suction, from Schrenk.

station is removed strongly enough by suction, the flow over the entire trailing-edge flap may be kept attached. After favorable wind tunnel results had been obtained, for flap profiles with suction, the Aerodynamische Versuchsanstalt Gottingen (AVA) conducted the first flight tests of the suction effect in the early 1930s. The possible gain in lift for fully attached flap flow {cq = Cq£) over the lift of uncontrolled •flow (cq = 0) may be seen in Fig. 8-23. This diagram shows cL as a function of flap deflection at several angles of attack of the wing. Note that the lift for potential flow is reached when the suction is just strong enough for complete prevention of separation. Arnold [4] studied the computation of the required amount cql. More recently, both slot suction and continuous suction through perforated walls have been applied, the latter at the trailing-edge flap as well as at the wing nose. Further developments of suction procedures have been summarized by Regenscheit [36] and Schlichting [36].

The continuously distributed suction has been studied theoretically by Schlichting and Pechau [38]. Flight tests by Schwarz [38] and by Schwarz and Wuest [38] confirm the feasibility of nose suction.

Ejection The boundary layer may be controlled by ejection as well as by suction for increased maximum lift. This method has been applied most successfully to the wing with a trailing-edge flap. By tangential ejection of a thin jet of high velocity at the nose of the deflected flap, flow separation from the flap can be prevented and the lift can be increased. Critical for the effectiveness of ejection is, according to Williams [51], the dimensionless momentum coefficient

= QjOjVj (8-20)


where the index / refers to the conditions in the jet and the index •» to those of the incident flow.

Comprehensive studies on the lift increase of flap wings with ejection have been conducted by Thomas [43]. In Fig. 8-24, a typical result of these measurements is given, namely, the gain in the lift coefficient AcL against the momentum coefficient Cj for several flap angles rif. The curves AcL versus Cj clearly show two ranges: first, a very steep increase at small momentum coefficients; and second, a considerably smaller increase at large momentum coefficients. The first range is that of boundary-layer control. It extends to the momentum coefficient that just suffices to produce complete flow attachment back to the flap trailing edge, thus completely preventing separation. The second range of considerably smaller lift gain with the momentum coefficient is the range of supercirculation. Here, the “hard jet” (of very high momentum) acts similarly to an extended mechanical flap.

In Fig. 8-25, the lift coefficient of a wing at fixed flap deflection is plotted against the angle of attack for several momentum coefficients су. The ejection has a similar effect as an increased camber (flap deflection). Flow separation sets in at smaller angles of attack, however, than without ejection. Inspecting Fig. 8-16 shows that an additional lift gain can be generated by combination with a boundary-layer

Figure 8-23 Lift increase due to slot suction at the trailing-edge flap for completely attached flap flow, from Arnold. (——————– ) Measurements with­out suction. ( ) Measurements with


Figure 8-24 Flap wing with ejection, lift increase леї vs – momentum coefficient Cj for various flap angles г?/ at constant angle of attack a — —5°, from Thomas.

Figure 8-25 Lift coefficient of a wing with ejection over the trailing-edge flap, from Williams, profile f/c = 0.08. Flap deflection rj/ = 45°, flap chord ratio hf = 0.25.

control at the wing nose, either by suction or by ejection (see Gersten [15]). Even when the flow is completely attached, a further increase in lift may be accomplished by stronger ejection on the flap. This is the result of supercirculation and the jet reaction force. This problem area has been summarized by Poisson – Quinton [34] and by Williams [51]; see also [28]. Levinsky and Schappelle [29] developed a method aimed at maintaining potential flow through tangential ejection on flap wings.

Jet flaps Effects very similar to those generated by a solid trailing-edge flap are obtained by ejecting a high-speed jet under a certain angle V/ near the wing trailing edge. This method, illustrated in Fig. 8-26, is termed a jet flap. The vertical component of the reaction force of the jet is supplemented by an induced lift that may be many times larger than the jet reaction (supercirculation). This effect has been studied by many experiments [34, 52].

In Fig. 8-26, the theories of Spence [42] and Jacobs [9] are compared with experiments on a symmetric profile with jet flap. The figure shows the dependence of the lift slopes dcLfda and dcLJdrjj on the momentum coefficient Cj as defined by Eq. (8-20). Here the momentum coefficients Cj are much larger than in Fig. 8-25. Up to values of about c;- = 0.1, the jet acts on the boundary layer; for larger values of Cj it essentially causes the circulation to increase (supercirculation). Either lift, slope increases strongly with increasing Cj. For Cj = 4, the lift slope dci/da has about twice the value of that without ejection (cy = 0). The agreement of theory

Figure 8-26 Profile with jet flap, com­parison of theory and experiment for lift slopes bcj^jda and Эс^/Эру. Theory from Spence and Jacobs. Measurements from Dimmok [52]. (®) rjy = 31°. O) py = 58°.

and experiment is good. Helmbold [20] studied the theory of the wing of finite span with jet flap.

A comprehensive wing theory for the wing of finite span with jet flap has been developed by Das [9]. An example of this theory and a comparison with experiments is given in Fig. 8-27 for a swept-back wing with a jet flap spanning the entire trailing edge. Agreement between theory and experiment is good. Murphy and Malmuth [9] report on the computation of the aerodynamics of the jet flap wing in transonic flow. The jet flap wing near the ground has been studied by Lohr [30]. The aerodynamic problems of the maximum lift have been summarized by Schlichting [37]. Questions of the practical application of the jet effect to the generation of high lift on wings with and without flap are discussed in the summarizing paper of Korbacher [25].

Air brakes, spoilers The aerodynamic effect of air brakes has been investigated repeatedly (see Arnold [3]). In particular, various positions of the brakes on the lower and upper sides of the wing have been studied. Figure 8-28 shows the result of three-component measurements for a wing with air brakes over the entire span. The polar curves illustrate the very large drag increase. Compared with the wing alone, the drag coefficient is about 20 times larger.

Devices of a similar kind mounted only on the upper side of the wing are also, termed spoilers. By extending them on only one side of the wing, they can be used

Figure 8-27 Lift coefficient of a swept-back wing with jet flap; com­parison of theory and measurement from Das [9]. Aspect ratio.1 = 3.5, sweepback angle уз = 45°, jet angle 7}/ = 30°.

Figure 8-28 Three-component measure­ments on a rectangular wing with air brake, from Reller [3]. Aspect ratio л = 5.1; flaps extend over the entire span. WO, wing without flap; S, flap on suction side; P, flap on pressure side.

for control about the vertical and longitudinal axes. The flow separation from the wing caused by the spoiler leads to a strong, one-sided lift loss and thus to a rolling moment. Wing tunnel test results on spoilers and a few computations on the effect of the spoiler are found in [11, 21, 23, 50].

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