DE102020117768A1 - Supersonic inlet with contour bulge - Google Patents

Abstract

In the case of a supersonic inlet (1) with a main flow direction (2), a central body (3) having a central body outer contour (4) and a hood (7) beginning opposite to the main flow direction (2) with a lip (6), one takes a transverse to the main flow direction (2) running free distance between the central body outer contour (4) and an inner contour of the hood (7) in the main flow direction (2). The central body outer contour (4) has a ramp (5) which begins upstream of the hood (7) and which approaches the inner contour (9) of the hood (7) in the main flow direction (2); and the ramp (5) ends in the main inflow direction with a contour bulge (8) of the central body outer contour (4), onto which a compression shock (10) emanating from the lip (6) of the hood (7) under at least one design inflow condition of the supersonic inlet (1) hits. The central body outer contour (4) is positioned in the area of the contour bulge (8) against a smoothed base line (14) of the central body outer contour (4) at a deflection angle α, which is with a coordinate x in the main flow direction (2) over a first bulge neck length with an increasing gradient rises from zero to a maximum deflection angle amax and then falls back to zero over a second half length L of the bulge with a falling gradient that results in a smooth curve of the wall pressure with minimal curvature over the contour bulge (8).

Description

TECHNICAL FIELD OF THE INVENTION

The invention relates to a supersonic enema. In particular, the invention relates to a supersonic inlet with a main flow direction and the further feature of the preamble of independent claim 1.

A supersonic inlet is understood here to mean an inlet for gas, in particular air, which is designed for a relative flow speed of the gas in the main direction of flow above the speed of sound in the gas.

STATE OF THE ART

Missiles that are equipped with air-breathing engines, for example, offer significant advantages over missiles with rocket propulsion in terms of their range. In supersonic flight in particular, missiles with air-breathing engines require special supersonic inlets that enable the compressed air to be fed robustly and efficiently to a ramjet. Supersonic inlets of current missiles are usually designed to be rigid and only function optimally in a narrow range of flight mach numbers. A complex interaction of physical flow processes in the rigid inner duct of the supersonic inlets determines the minimum Mach number from which the ramjet engine is supplied with sufficient air. The entrance area of the inner channel is particularly critical and susceptible, in which an initial compression shock, which occurs on a hood of the supersonic inlet and is also referred to as hood shock, hits the flow boundary layer on a central body outer contour opposite the hood and thereby typically induces local detachment bubbles and other compression shocks . Mitigating this interaction is one of the most important tasks in the development of new air-breathing engines.

As a measure to weaken the so-called shock / boundary layer interaction, suction or bleed of the boundary layer through openings in the outer contour of the central body near the point of impact are known, see for example DE 1 198 204 B , U.S. 4,000,869 A , U.S. 5,088,660 A , U.S. 5,397,077 A and US 7 690 595 B2 . The suction or discharge of the boundary layer reduces the tendency of the flow to detach in areas with positive pressure gradients over the outer contour of the central body.

Another known measure for weakening the impact / boundary layer interaction is a kink in the outer contour of the central body in the area of the impact point of the canopy joint, which can completely eliminate the reflected compression shock with flight Mach numbers in the range of a corresponding design flight Mach number. See also Y. Zhang et al .: "Influence of Expansion Waves on Cowl Shock / Boundary Layer Interaction in Hypersonic Inlets" J. of Propulsion and Power, Vol. 30, No. 5, pages 1183-1191, September-October 2014 .

Yet another known measure for weakening the impact / boundary layer interaction is based on the application of a flat bulge in the area of the point of impact of the hood impact with the aim of attenuating or completely preventing impact reflections, see FIG Y. Zhang et al .: "Control of Cowl-Shock / Boundary-Layer Interactions by Deformable Shape-Memory Alloy Bump", AIAA J., Vol. 57 No. 2, pages 696-705, February 2019 .

From the US 9 429 071 B2 For example, it is known to use upstream vortex generators to reduce the separation bubble caused by the impact / boundary layer interaction. Other upstream flow control measures to limit the size of the separation bubble go from the U.S. 3,417,767 A emerged.

In V. Shinde et al .: "Control of Transitional Shock Wave Boundary Layer Interaction using Surface Morphing", AIAA Scitch 2019 Forum, AIAA 2019-1895, https://doi.org/10.2514/6.2019-1895 the results of a direct numerical simulation of an impact / boundary layer interaction with adaptive adaptation of the outer contour of the central body to avoid a separation bubble are described. A resulting bulge in the outer contour of the central body that is optimal in terms of the change in the detachment bubble is a flat ridge, the ridge of which is hit by the bonnet joint without a detachment bubble without the bulge forming at a constant Mach number of 2.

From the CN 107 091 158 A there is known a supersonic inlet with a two-dimensional bulge in an inlet of the supersonic inlet. Air discharge holes or slots are arranged in inwardly concave areas on the windward and leeward sides of the bulge, and an air discharge cavity below the bulge is divided into two sub-cavities, both of which can be opened and closed with the aid of electromagnets.

From the U.S. 3,450,141 A a supersonic inlet with an isentropic ramp of a central body upstream of a hood of the supersonic inlet is known. The height of the end of this ramp and thus the height of the inlet cross section of the supersonic inlet is adjustable. A movable inlet surface is attached to the ramp below the hood, which is hinged to the ramp such that an outer leading edge of the movable surface is held in the vicinity of the rear edge of the ramp. The isentropic ramp focuses the compression wave of the incoming flow.

OBJECT OF THE INVENTION

The invention is based on the object of providing a supersonic inlet whose central body outer contour is shaped in such a way that the formation of a detachment bubble due to the impact / boundary layer interaction in the area of impact of the hood impact is avoided as completely as possible over various Mach numbers.

SOLUTION

The object of the invention is achieved by a supersonic inlet with the features of independent claim 1. Preferred embodiments of the supersonic inlet according to the invention are defined in the dependent claims 2 to 10. Claim 11 relates to a supersonic aircraft with an engine with a supersonic inlet according to the invention.

DESCRIPTION OF THE INVENTION

oder $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{max}}={\left({\text{x/L}}_1\right)}^{\text{b1}}$$

von null bis auf einen maximalen Ablenkwinkel amax ansteigt und der dann nach einem Übergangsbereich, in dem α(x) kleiner oder gleich amax gilt und der sich in der Anströmungshauptrichtung über eine Übergangslänge ÜL erstreckt, wobei ÜL 0 % bis 10 % von L₁ beträgt, über eine zweite Beulenteillänge L₂ von x=(L₁+ÜL) bis x=(L₁+ÜL+L₂) gemäß $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{max}}=\left(1+{\text{a}}_2\right){\left(\left({\text{L}}_1+\text{L}+{\text{L}}_2-\text{x}\right)/{\text{L}}_2\right)}^2-{\text{a}}_2{\left(\left({\text{L}}_1+\text{L}+{\text{L}}_2-\text{x}\right)/{\text{L}}_2\right)}^3$$

oder $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{max}}=(\left({\text{L}}_1+\text{L}+{\text{L}}_2-\text{x}\right){/\text{L})}^{\text{b2}}$$

wieder gegen null abfällt. Dabei beträgt L₂ 90 % bis 110 % von L₁; a₁ und a₂ sind unabhängig voneinander gleich 0,69 +/- 0,3; und b₁ und b₂ sind unabhängig voneinander gleich 1,96 +/- 0,3.A supersonic inlet according to the invention has a main flow direction, a central body having a central body outer contour and a hood beginning with a lip opposite to the main flow direction. In this case, a free distance running transversely to the main flow direction decreases between the central body outer contour and an inner contour of the hood in the main flow direction. The outer contour of the central body has a ramp which begins upstream of the hood and which approaches the inner contour of the hood in the main direction of flow. In the main inflow direction, the ramp ends with a contour bulge of the central body outer contour, which is hit by a compression shock emanating from the lip of the hood, ie the so-called hood shock, under at least one design flow condition of the supersonic inlet. In the area of the contour bulge, the central body outer contour is at a deflection angle compared to a smoothed base line of the central body outer contour α employed with a coordinate x in the main direction of flow over a first bulge part length L ₁ from x = 0 to x = L ₁ according to $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{Max}}=\left(1+{\text{a}}_1\right){\left({\text{x / L}}_1\right)}^2-{\text{a}}_1{\left({\text{x / L}}_1\right)}^3$$

or $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{Max}}={\left({\text{x / L}}_1\right)}^{\text{b1}}$$

from zero to a maximum deflection angle amax and which then rises after a transition area in which α (x) is less than or equal to amax and which extends over a transition length ÜL in the main flow direction, where ÜL is 0% to 10% of L ₁ , over a second buckle part length L ₂ from x = (L ₁ + ÜL) to x = (L ₁ + ÜL + L ₂ ) according to $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{Max}}=\left(1+{\text{a}}_2\right){\left(\left({\text{L.}}_1+\text{L.}+{\text{L.}}_2-\text{x}\right)/{\text{L.}}_2\right)}^2-{\text{a}}_2{\left(\left({\text{L.}}_1+\text{L.}+{\text{L.}}_2-\text{x}\right)/{\text{L.}}_2\right)}^3$$

or $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{Max}}=(\left({\text{L.}}_1+\text{L.}+{\text{L.}}_2-\text{x}\right){/\text{L.})}^{\text{b2}}$$

falls back to zero. L _{2 is} 90% to 110% of L ₁ ; a ₁ and a ₂ are independently equal to 0.69 +/- 0.3; and b ₁ and b ₂ are independently equal to 1.96 +/- 0.3.

The dent part lengths L ₁ and L; can both be equal to a bulge half length L, the transition area being omitted, so that the transition length ÜL = 0 applies. Furthermore, a ₁ = a ₂ = a and b ₁ = b ₂ = b can apply.

oder $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{max}}={\left(\text{x/L}\right)}^{\text{b}},$$

von null bis auf einen maximalen Ablenkwinkel amax ansteigt und dann von x=L bis x=2L gemäß $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{max}}=\left(1+\text{a}\right){\left(\text{2}-\text{x/L}\right)}^2-\text{a}{\left(\text{2}-\text{x/L}\right)}^3,$$

oder $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{max}}={\left(\text{2}-\text{x/L}\right)}^{\text{b}},$$

wieder gegen null abfällt, wobei a gleich 0,69 ± 0,3 ist und b gleich 1,96 ± 0,3 ist.Then the central body outer contour is in the area of the contour bulge with respect to the smoothed base line of the central body outer contour at a deflection angle α employed, which corresponds to the coordinate x in the main flow direction from x = 0 to x = L $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{Max}}=\left(1+\text{a}\right){\left(\text{x / L}\right)}^2-\text{a}{\left(\text{x / L}\right)}^3,$$

or $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{Max}}={\left(\text{x / L}\right)}^{\text{b}},$$

increases from zero to a maximum deflection angle amax and then from x = L to x = 2L according to $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{Max}}=\left(1+\text{a}\right){\left(\text{2}-\text{x / L}\right)}^2-\text{a}{\left(\text{2}-\text{x / L}\right)}^3,$$

or $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{Max}}={\left(\text{2}-\text{x / L}\right)}^{\text{b}},$$

drops again towards zero, where a is equal to 0.69 ± 0.3 and b is equal to 1.96 ± 0.3.

In any case, both result in the course of the deflection angle α The equations defining the first and second part length L ₁ and L ₂ , respectively, with fixed values of amax and L ₁ and L _2, very similar contour bumps. It is not only for this reason that the fundamentally different equation for the course of the deflection angle can also be used via the second bulge part length L ₂ α can be selected as over the first dent part length L ₁ .

With the contour bulge of the supersonic inlet according to the invention, a favorable profile of the wall pressure over the central body outer contour is achieved. This favorable course of the wall pressure begins at x = 0 with dp / dx = 0 and ends at x = (L ₁ + ÜL + L ₂ ) with dp / dx = 0. In between, the gradient dp / dx of the wall pressure increases monotonically with the least possible curvature | d ² p / dx ² | at.

Without a transition area, so that the transition length ÜL = 0, the contour bulge of the supersonic inlet according to the invention has a ridge at the bulge half length L, ie a discontinuity in the course of the central body outer contour with a change in the sign at the deflection angle α on. Nevertheless, the desired steady course of the wall pressure results across this ridge. This steady course of the wall pressure results in a particularly low sensitivity of the flow into the supersonic inlet with regard to the formation of detachment bubbles due to the impact / boundary layer interaction under the action of the compression shock emanating from the hood. With the contour bulge of the supersonic inlet according to the invention, the steady course of the wall pressure is also achieved over a wide range of Mach numbers. In this respect, the contour bulge is universal.

If ÜL> 0 applies to the transition length, the contour bulge of the supersonic inlet according to the invention shows when the sign of the deflection angle changes α not necessarily a sharp ridge, but in particular a transition area that is rounded or flattened compared to such a sharp ridge. As a result of this transition area, the desired effect of the contour bulge of the supersonic inlet according to the invention is less sensitive to migration of the point of impact of the compression shock emanating from the hood on the contour bulge with regard to the desired constant course of the wall pressure.

The maximum deflection angle amax can in particular be selected to be an acute angle at which the ramp of the central body outer contour approaches the inner contour of the hood in the main flow direction.

Typically, this acute angle in conventional supersonic inlets and correspondingly the maximum deflection angle α _max in the supersonic inlet according to the invention in radians in a range from 0.15 or from 0.17 to 0.35 or up to 0.4, the narrower limits of this range correspond to about 10 ° to 20 °.

In the case of the supersonic inlet according to the invention, the bulge part lengths L ₁ and L ₂ are in a typical range from 0.025 h _e / α _max to 0.75 h _e / α _max . Here, h _{e is} the smallest free distance between the central body outer contour and the inner contour of the hood transversely to the main flow direction or the smallest height of the inner channel of the supersonic inlet in the area of the contour bulge. The resulting maximum height H of the contour bulge above the smoothed base line of the central body outer contour is between 0.01 h _e and 0.3 h _e , since this height H and the bulge half length L with the maximum deflection angle α _{max are} practically equal in the relationship H / L κ α _max , where κ is 0.39 ± 0.04 and α _{max is given} in radians.

Another approach to specifying a typical bulge part length L ₁ or L ₂ or bulge half length L in the supersonic inlet according to the invention is that the bulge part length L ₁ or L ₂ or bulge half length L is at least as long as a detachment bubble in the main direction of flow, which is without the contour bulge under the at least one design flow condition above the smoothed baseline of the central body outer contour.

The variables a ₁ and a ₂ as well as b ₁ and b ₂ in the definition of the universal contour bulge of the supersonic inlet according to the invention are preferably set to the values of 0.69 and 1.96 more precisely than ± 0.3. Specifically, the tolerance ranges can be limited to ± 0.2 or even ± 0.1.

With regard to its course over its two bulge part lengths L ₁ and L ₂ , the contour bulge of the supersonic inlet according to the invention is universal over various Mach numbers of the flow to the supersonic inlet. If, however, in the case of a supersonic inlet, the point of impact of the compression shock emanating from the lip of the hood on the outer contour of the central body with the Mach number of the inflow to the supersonic inlet moves further than can be compensated for by a transition area with a limited transition length ÜL, it can make sense to place the contour bulge in the main direction of the inflow so that the shock wave always hits the highest point of the contour bulge as precisely as possible. For this purpose, the contour bulge can, for example, be displaceable in the main direction of flow using a displacement device comprising a motorized drive. This displacement device can also be designed in such a way that it moves the contour bulge depending on the current one The flow condition of the supersonic inlet shifts so that the compression shock emanating from the lip of the hood hits the contour bulge after the first part length of the bulge at x = L ₁ or in the possibly adjacent transition area, i.e. up to x = L ₁ + ÜL.

Furthermore, the displacement device can be designed to control the contour bulge depending on signals from a first pressure sensor in the area of the first bulge half length and a second pressure sensor in the area of the second bulge half length in such a way that a pressure difference between wall pressures over the central body outer contour, from the first pressure sensor and from the second pressure sensor are detected, is minimized. When this minimization is achieved, the desired steady pressure increase with minimal curvature of the pressure curve is present.

Specifically, in the case of the supersonic inlet according to the invention, the contour bulge can be formed on a bulge body which is guided on a main body of the central body so as to be displaceable along the base line.

A supersonic aircraft according to the invention with one engine has a supersonic inlet according to the invention.

Advantageous further developments of the invention emerge from the patent claims, the description and the drawings.

The advantages of features and of combinations of several features mentioned in the description are only exemplary and can come into effect alternatively or cumulatively without the advantages necessarily having to be achieved by embodiments according to the invention.

With regard to the disclosure content - not the scope of protection - of the original application documents and the patent, the following applies: Further features can be found in the drawings - in particular the illustrated geometries and the relative dimensions of several components to one another and their relative arrangement and operative connection. The combination of features of different embodiments of the invention or of features of different patent claims is also possible, deviating from the selected back-references of the patent claims, and is hereby suggested. This also applies to features that are shown in separate drawings or mentioned in their description. These features can also be combined with features of different patent claims. Features listed in the claims can also be omitted for further embodiments of the invention, but this does not apply to the independent claims of the granted patent.

The number of features mentioned in the claims and the description are to be understood in such a way that precisely this number or a greater number than the specified number is present without the need for the explicit use of the adverb “at least”. For example, when an element is mentioned, it is to be understood that there is exactly one element, two elements or more elements. The features listed in the patent claims can be supplemented by other features or be the only features that the respective product has.

Figure list

• 1 zeigt eine erste Ausführungsform eines erfindungsgemäßen Überschalleinlaufs im Längsschnitt.
• 2 zeigt eine zweite Ausführungsform des erfindungsgemäßen Überschalleinlaufs im Längsschnitt.
• 3 zeigt das Auftreffen eines Verdichtungsstoßes auf eine Konturbeule des erfindungsgemäßen Überschalleinlaufs und die dabei induzierte Wellenstruktur.
• 4 zeigt den Verlauf des Wanddrucks über der Konturbeule gemäß 3 zusammen mit einem Verlauf des Wanddrucks in einem Referenzfall.
• 5 ist eine Auftragung eines normierten Ablenkwinkels der Konturbeule gemäß 3 über deren ersten Beulenhalblänge L.
• 6 ist eine Auftragung eines konkreten Verlaufs der Konturbeule gemäß 3 mit dem normierten Ablenkwinkel gemäß 5.
• 7 ist der zu 6 zugehörige Verlauf des Ablenkwinkels der Konturbeule gemäß 3.
• 8 ist eine Auftragung des normierten Wanddrucks an der Oberfläche der Konturbeule gemäß 3 bei einer Variation der Anströmmachzahl M_∞ zwischen 2,0 und 3,0; und
• 9 ist eine Auftragung von konkreten Verläufen weiterer Ausführungsformen der Konturbeule.

In the following, the invention is further explained and described with reference to preferred exemplary embodiments shown in the figures.

• 1 shows a first embodiment of a supersonic inlet according to the invention in longitudinal section.
• 2 shows a second embodiment of the supersonic inlet according to the invention in longitudinal section.
• 3 shows the impact of a shock wave on a contour bulge of the supersonic inlet according to the invention and the wave structure induced thereby.
• 4th shows the course of the wall pressure over the contour bulge according to 3 together with a course of the wall pressure in a reference case.
• 5 is a plot of a normalized deflection angle of the contour bulge according to 3 over the first half-length of the bump L.
• 6 is a plot of a specific course of the contour bulge according to 3 with the normalized deflection angle according to 5 .
• 7th is that to 6 associated course of the deflection angle of the contour bulge according to 3 .
• 8th is a plot of the normalized wall pressure on the surface of the contour bulge according to 3 with a variation of the inflow number M _∞ between 2.0 and 3.0; and
• 9 is a plot of specific courses of further embodiments of the contour bulge.

FIGURE DESCRIPTION

The in 1 illustrated supersonic inlet 1 is for one main flow direction 2 educated. A central body 3 of the supersonic enema 1 has a central body outer contour 4th on. The central body outer contour 4th has a two-step ramp here 5 on that upstream of a lip 6 a hood 7th of the supersonic enema 1 begins and with a contour bump 8th ends. In the area of its downstream step, the ramp approaches at an acute angle 18th to the main flow direction 2 to an inner contour 9 the hood 7th at. One from the lip 6 outgoing shock wave 10 meets a flow of the supersonic inlet depending on an inflow number 1 in the main flow direction 2 at different points on the central body outer contour 4th on. A sliding device 11 for a bump body 12th on which the contour bump 8th formed, is controlled so that the contour bump 8th each is arranged so that the shock wave 10 under the respective flow conditions to the highest point of the contour bulge 8th occurs. This and a special shape of the contour bulge 8th prevents a boundary layer 13 the oncoming flow due to an impact / boundary layer interaction of the compression shock 10 with the boundary layer 13 over a separation bubble away from the central body outer contour 4th replaces. In 1 forms the body of the bump 12th the central body outer contour 4th of the central body 3 in an area in which a smooth baseline of the central body outer contour, over which the contour bump is located 8th rises, has a constant radius of curvature. The bump body 12th is guided on a circular path with this radius of curvature and is operated by the displacement device 11 moved on this circular path.

At the supersonic enema 1 who is in 2 is shown, and otherwise completely according to the embodiment 1 corresponds, the shifting device 11 a conveyor belt 15th on with which the bump body 12th here over a surface of the central body 3 with the course of the baseline 14th is movable. This also serves to reduce the contour bump 8th always align so that the shock wave 10 to the highest point of the contour bump 8th hits.

3 outlines the induced wave structure when it hits the lip 6 according to 1 or 2 outgoing shock wave 10 to the highest point of the contour bulge 8th is formed at the length of their bulge. Upstream and downstream of the point of impact of the shock wave 10 compression waves are formed 16 from and from the point of impact of the shock wave 10 expansion waves go 17th out.

4th shows the to 3 associated course of the wall pressure over the central body outer contour 5 with a solid line. Starting from a gradient of 0, the wall pressure rises steadily, initially as the gradient increases and then as the gradient falls back to 0. The curvature of the pressure curve is minimal. This creates ideal conditions for avoiding a detachment bubble in the area of impact of the shock wave 10 given. The situation is different with the pressure curve shown in 4th is shown with a dashed line and the impact of the shock wave on the in 3 baseline also drawn with a dashed line 14th results. This pressure curve has strong curvatures and therefore quickly results in the unwanted formation of a separation bubble in the boundary layer 13 according to 3 .

To get the in 4th To realize the course of the wall pressure shown starting and ending with a gradient of 0 and exhibiting a minimal curvature, a course of the deflection angle normalized to the maximum deflection angle α _max is coordinated therewith α the contour bump 8th to be determined. Where is the deflection angle α the angle of the respective area of the central body outer contour 4th in the area of the contour bulge 8th opposite the baseline 14th .

In 5 the normalized deflection angle α / α _{max is} plotted over the normalized longitudinal coordinate x / L of the contour bulge, where L is the half-length of the contour bulge 8th is. The normalized deflection angle α / α _max starts at 0 and then increases with the gradient increasing from 0 at x / L equal to 1 to α / α _max equal to 1.

For a specific dimensioning of the contour bulge in relation to the maximum deflection angle α _max and the bulge half length L in relation to a specific supersonic inlet 1 the maximum deflection angle α _{max can be} equal to the angle 18th be set under which the ramp is 13 to the main flow direction 2 to the inner contour 9 approaches the hood. The bump half length L can be depending on a smallest free distance h _e according to 2 and 3 between the central body outer contour 5 and the inner contour 9 the hood 7th can be selected in the range of 0.025 h _e / α _max and 0.75 h _e / α _max .

6 shows that from the course of the normalized deflection angle according to 5 for a bulge half length of 5 mm and a maximum deflection angle α _max of 10 ° resulting coordinates of the contour bulge 8th .

7th represents the corresponding course of the deflection angle α in degrees. Show 6 and 7th that the second half of the contour bump 8th downstream of their crest, ie from L to 2 L, a mirror-symmetrical image of their wind side from 0 to L is.

8th shows a normalized pressure distribution p _W / p _{Wmax of} the wall pressure on the surface of the Contour bump 8th according to the 4th to 7th for inflow numbers M _∞ of 2.0 and 3.0, which were predicted using the characteristic method (without boundary layer influence). p _Wmax corresponds to the absolute wall pressure at the rear edge of the bulge, which is achieved with the respective inflow number M _∞ . It is obvious that the same contour bump 8th the desired pressure curve according to 4th the negligible variations are realized with both of the inflow numbers M _∞ taken into account. The contour bulge according to the invention is therefore optimized for different inflow numbers and to that extent universal.

9 shows different embodiments of the contour bulge 8th which is not formed in two identical bulge half lengths L, but in two partial bulge lengths L ₁ and L ₂ with a transition area of a transition length UL arranged in between. The deflection angle increases over the length of the bulge part L ₁ and L ₂ α basically exactly the same or the deflection angle falls α basically exactly the same as over the two bulge half lengths L. A deviation of the course of the deflection angle from an exact mirror symmetry can also mean a different length of L ₁ and L ₂ . There is also the transition area 19th . Here the deflection angle can be up to a ridge 20th as he did in the previously described embodiments of the contour bulge 8th is present, stay the same and then along the ridge 20th suddenly change its sign. The following applies to the amount of the deflection angle α in the transition area 19th that it remains ≤ α _max . Alternatively, in the transition area 19th a rounded ridge 21st be provided in which the deflection angle α falls steadily to zero and then increases again in magnitude after its sign change. In principle, it is also possible that in the transition area 19th a flattened ridge 22nd is provided with a deflection angle of zero. (In 9 the transition length ÜL is shown too long compared to the buckle part lengths L ₁ and L ₂ . The transition length ÜL is not more than 10% of L ₁ and also not more than 10% of L ₂ , even if L ₁ and L ₂ differ. This difference between L ₁ and L _{2 also} remains ≤ 10% of L _1. )

List of reference symbols

1

Supersonic enema

2

Main flow direction

3

Central body

4th

Central body outer contour

5

ramp

6th

lip

7th

Hood

8th

Contour bump

9

Inner contour

10

Shock wave

11

Sliding device

12

Bump body

13

Boundary layer

14th

Baseline

15th

Conveyor belt

16

Compression wave

17th

Expansion wave

18th

angle

19th

Transition area

20th

ridge

21st

rounded ridge

22nd

flattened ridge

α

Deflection angle

QUOTES INCLUDED IN THE DESCRIPTION

This list of the documents listed by the applicant was generated automatically and is included solely for the better information of the reader. The list is not part of the German patent or utility model application. The DPMA assumes no liability for any errors or omissions.

Patent literature cited

• DE 1198204 B [0004]
• US 4000869 A [0004]
• US 5088660 A [0004]
• US 5397077 A [0004]
• US 7690595 B2 [0004]
• US 9429071 B2 [0007]
• US 3417767 A [0007]
• CN 107091158 A [0009]
• US 3450141 A [0010]

Non-patent literature cited

• Y. Zhang et al .: "Influence of Expansion Waves on Cowl Shock / Boundary Layer Interaction in Hypersonic Inlets" J. of Propulsion and Power, Vol. 30, No. 5, pages 1183-1191, September-October 2014 [0005]
• Y. Zhang et al .: "Control of Cowl-Shock / Boundary-Layer Interactions by Deformable Shape-Memory Alloy Bump", AIAA J., Vol. 57 No. 2, pages 696-705, February 2019 [0006]
• V. Shinde et al .: "Control of Transitional Shock Wave Boundary Layer Interaction using Surface Morphing", AIAA Scitch 2019 Forum, AIAA 2019-1895, https://doi.org/10.2514/6.2019-1895 [0008]

Claims (11)

Supersonic inlet (1) with - a main flow direction (2), - a central body (3) having a central body outer contour (4) and - a hood (7) beginning opposite to the main flow direction (2) with a lip (6), - one being transverse to the main flow direction (2) running free distance between the central body outer contour (4) and an inner contour of the hood (7) decreases in the main flow direction (2), - the central body outer contour (4) having a ramp (5) beginning upstream of the hood (7) , which approaches the inner contour (9) of the hood (7) in the main inflow direction (2), and - the ramp (5) ends in the main inflow direction with a contour bulge (8) of the central body outer contour (4) onto which one of The compression shock (10) emanating from the lip (6) of the hood (7) strikes the supersonic inlet (1) under at least one design flow condition , characterized in that the central body outer contour (4) in the area of the contour The bulge (8) is set against a smoothed base line (14) of the central body outer contour (4) at a deflection angle α which, with a coordinate x in the main flow direction (2) over a first bulge part length L ₁ from x = 0 to x = L ₁ according to $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{Max}}=\left(1+{\text{a}}_1\right){\left({\text{x / L}}_1\right)}^2-{\text{a}}_1{\left({\text{x / L}}_1\right)}^3$$

or $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{Max}}={\left({\text{x / L}}_1\right)}^{\text{b1}}$$

from zero to a maximum deflection angle amax and then after a transition area in which α (x) is less than or equal to amax and which extends over a transition length ÜL in the main flow direction (2), where ÜL 0% to 10% of L ₁ is, over a second buckle part length L ₂ from x = (L ₁ + ÜL) to x = (L ₁ + ÜL + L ₂ ) according to $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{Max}}=\left(1+{\text{a}}_2\right){\left(\left({\text{L.}}_1+\text{L.}+{\text{L.}}_2-\text{x}\right)/{\text{L.}}_2\right)}^2-{\text{a}}_2{\left(\left({\text{L.}}_1+\text{L.}+{\text{L.}}_2-\text{x}\right)/{\text{L.}}_2\right)}^3$$

or $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{Max}}={\left(\left({\text{L.}}_1+\text{L.}+{\text{L.}}_2-\text{x}\right)/\text{L.}\right)}^{\text{b}2}$$

again drops to zero, where - L _{2 is} 90% to 110% of L ₁ , - a ₁ and a _{2 are} independently equal to 0.69 +/- 0.3 and - b ₁ and b _{2 are} independently equal to 1, 96 are +/- 0.3. Supersonic inlet (1) after Claim 1 , characterized in that the maximum deflection angle amax is equal to an acute angle at which the ramp (5) of the central body outer contour (4) approaches the inner contour (9) of the hood (7) in the main flow direction (2). Supersonic inlet (1) after Claim 1 or 2 , characterized in that the maximum deflection angle amax in radians is in a range from 0.15 to 0.4. Supersonic inlet (1) according to one of the preceding claims, characterized in that the following applies to the first part length L ₁ : $$0.025{\text{H}}_{\text{e}}/{\mathrm{\alpha }}_{\text{Max}}\le {\text{L.}}_1\le 0.75{\text{H}}_{\text{e}}/{\mathrm{\alpha }}_{\text{Max}}$$

, where h _{e is} the smallest free distance between the central body outer contour (4) and the inner contour (9) of the hood (7) transversely to the main flow direction (2). Supersonic inlet (1) according to one of the preceding claims, characterized in that the first bulge part length L _{1 is} at least as large as a length of a detachment bubble in the main flow direction which, under the at least one design flow condition, is above the smoothed base line (14) of the central body outer contour (4 ) without the contour bulge (8). Supersonic inlet (1) according to one of the preceding claims, characterized in that - a ₁ and a _{2 are} independently equal to 0.69 +/- 0.2 and - b ₁ and b _{2 are} independently equal to 1.96 +/- 0 , 2 are. Supersonic inlet (1) according to one of the preceding claims, characterized in that the contour bulge (8) can be displaced in the main flow direction (2) with a displacement device (11) comprising a motor drive. Supersonic inlet (1) after Claim 7 , characterized in that the Displacement device (11) is designed in such a way that it displaces the contour bulge (8) depending on the current flow conditions of the supersonic inlet (1) so that the compression shock (10) emanating from the lip (6) of the hood (7) in the transition area or if ÜL is equal to 0% of L ₁ , meets the contour bulge (8) at the first half length x = L ₁ . Supersonic inlet (1) after Claim 7 or 8th , characterized in that the displacement device (11) is designed such that it controls the contour bulge (8) depending on signals from a first pressure sensor in the area of the first bulge half length and a second pressure sensor in the area of the second bulge half length so that a pressure difference between wall pressures over the central body outer contour (4), which are detected by the first pressure sensor and the second pressure sensor, is minimized. Supersonic inlet (1) after one of the Claims 7 to 9 , characterized in that the contour bulge (8) is formed on a bulge body (12) which is displaceably guided along the base line (14) on a main body of the central body (3). Supersonic aircraft with an engine with a supersonic inlet (1) according to one of the preceding claims.

Images