DE102020117768B4 - Supersonic inlet with contour bump - Google Patents

Abstract

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 with a lip (6) counter to the main flow direction (2),- with a transverse to the free distance running in the main flow direction (2) between the central body outer contour (4) and an inner contour of the hood (7) in the main flow direction (2) decreases, - 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 direction of inflow (2), and the ramp (5) ends in the main direction of inflow with a contour bulge (8) of the outer contour of the central body (4), onto which a The compression shock (10) emanating from the lip (6) of the hood (7) impinges under at least one design inflow condition of the supersonic inlet (1), characterized in that the outer contour (4) of the central body in the region of the contour bulge (8) compared to a smoothed base line (14) of the outer contour of the central body (4) at a deflection angle α, which corresponds to a coordinate x in the main direction of flow (2) over a first partial bulge length L1 from x=0 to x=L1 according to α(x)/αmax =(1+a1)(x/L1)2−a1(x/L1)3orα(x)/αmax=(x/L1)b1 increases from zero to a maximum deflection angle amax and then after a transition region where α( x) is less than or equal to amax and which extends in the main direction of inflow (2) over a transition length ÜL, where ÜL is 0% to 10% of L1, over a second partial bulge length L2 from x=(L1+ÜL) to x=(L1+ ÜL +L2) according toα(x)/αmax=(1+a1)((L1+U¨L+L2−x)/L)2−a2((L1+U¨L+L2−x)/L2)3orα (x)/αmax=((L1+U¨L+L2−x)/L)b2 falls again towards zero, where - L290 % to 110 % of L1 is, - a1 and a2 independently equal to 0.69 +/- 0.3 and- b1 and b2 independently equal 1.96 +/- 0.3.

Description

TECHNICAL FIELD OF THE INVENTION

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

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

STATE OF THE ART

Missiles equipped with air-breathing engines offer significant advantages over rocket-propelled missiles in terms of range, for example. In supersonic flight in particular, missiles with air-breathing engines require special supersonic intakes that enable the compressed air to be supplied to a ramjet engine in a robust and efficient manner. Supersonic intakes of current missiles are generally designed to be rigid and only function optimally in a narrow range of flight Mach numbers. A complex interaction of flow-physical processes in the rigid inner channel of the supersonic inlets determines the minimum flight Mach number from which the ramjet engine can be adequately supplied with air. The entrance area of the inner channel is particularly critical and vulnerable, in which an initial compression shock, which occurs at a hood of the supersonic inlet and is also referred to as hood shock, impinges on the flow boundary layer on a central body outer contour opposite the hood and thereby typically induces local separation bubbles and further 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 impact/boundary layer interaction, suction or bleed of the boundary layer through openings in the central body outer contour near the point of impact are known, see for example DE 1 198 204 B , U.S. 4,000,869A , U.S. 5,088,660A , U.S. 5,397,077 A and U.S. 7,690,595 B2 . The suction or draining of the boundary layer reduces the tendency of the flow to separate in areas with positive pressure gradients over the outer contour of the central body.

Another known measure for weakening the shock/boundary layer interaction is a kink in the outer contour of the central body in the area of the impact point of the canopy shock, which can completely eliminate the reflected compression shock with flight Mach numbers in the range of a corresponding design flight Mach number. See Y. Zhang et al.: "Influence of Expansion Waves on Cowl Shock/Boundary Layer Interaction in Hypersonic Inlets" J. of Propulsion and Power, Vol. 5, pp. 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 impact point of the cowl impact with the aim of weakening or completely preventing impact reflections, see 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 U.S. 9,429,071 B2 it is known to use upstream vortex generators to reduce the separation bubble caused by impact/boundary layer interaction. Other upstream flow control measures to limit the size of the separation bubble come from the U.S. 3,417,767 A out.

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 are results of a direct numerical Simulation of an impact/boundary layer interaction with adaptive adjustment of the central body outer contour to avoid a separation bubble is described. A resulting bulge in the outer contour of the central body, which is optimal with regard to the change in the separation bubble, is a flat ridge, on the ridge of which the canopy impact strikes, without a separation bubble forming without the bulge at a constant Mach number of 2.

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

From the US 3,450,141A discloses a supersonic intake with an isentropic ramp of a central body upstream of a hood of the supersonic intake. The height of the end of this ramp and thus the height of the entry cross-section of the supersonic inlet is adjustable. Adjoining the ramp below the hood is a moveable inlet surface which is hinged to the ramp such that an outer leading edge of the moveable surface is maintained proximate the trailing edge of the ramp. The isentropic ramp focuses the compressional wave of the incoming flow.

OBJECT OF THE INVENTION

The invention is based on the object of demonstrating a supersonic intake whose outer contour of the central body is shaped in such a way that the formation of a separation bubble as a result of the impact/boundary layer interaction in the impact area of the canopy impact is avoided as completely as possible over different Mach numbers.

SOLUTION

The object of the invention is achieved by a supersonic inlet with the features of independent patent claim 1 . Preferred embodiments of the supersonic inlet according to the invention are defined in the dependent claims 2-10. Patent claim 11 relates to a supersonic aircraft with an engine with a supersonic intake 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 $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{max}}=\left(1+{\text{a}}_1\right)(\left({\text{L}}_1+\ddot{\text{U}}\text{L}+{\text{L}}_2-\text{x}\right){/{\text{L}}_2)}^2-{\text{a}}_2(\left({\text{L}}_1+\ddot{\text{U}}\text{L}+{\text{L}}_2-\text{x}\right){/{\text{L}}_2)}^3$$

oder $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{max}}=(\left({\text{L}}_1+\ddot{\text{U}}\text{L}+{\text{L}}_2-\text{x}\right){/{\text{L}}_2)}^{\text{b}2}$$

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 direction of inflow, a central body having a central body outer contour and a hood beginning with a lip counter to the main direction of inflow. A free distance running transversely to the main flow direction between the central body outer contour and an inner contour of the hood decreases in the main flow direction. The outer contour of the central body has a ramp that begins upstream of the hood and approaches the inner contour of the hood in the main direction of inflow. In the main direction of inflow, the ramp ends with a contour bulge of the outer contour of the central body, which a compression shock emanating from the lip of the hood, ie the so-called hood shock, impinges under at least one design inflow condition of the supersonic inlet. In the area of the contour bulge, the outer contour of the central body is set at a deflection angle α relative to a smoothed base line of the outer contour of the central body, which corresponds to a coordinate x in the main direction of flow over a first partial bulge length L ₁ from x=0 to x=L ₁ $$\mathrm{a}\left(\text{x}\right)/{\mathrm{a}}_{\text{Max}}=\left(1+{\text{a}}_1\right){\left({\text{x/L}}_1\right)}^2-{\text{a}}_1{\left({\text{x/L}}_1\right)}^3$$

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

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 in the main flow direction over a transition length ÜL, where ÜL is 0% to 10% of L ₁ , over a second partial bump length L ₂ from x=(L ₁ +ÜL) to x=(L ₁ +ÜL +L ₂ ) according to $$\mathrm{a}\left(\text{x}\right)/{\mathrm{a}}_{\text{Max}}=\left(1+{\text{a}}_1\right)(\left({\text{<figure-callout id="1" label="L" filenames="DE102020117768B4\_0021.png,DE102020117768B4\_0022.png" state="{{state}}">L</figure-callout>}}_1+\ddot{\text{u}}\text{L}+{\text{<figure-callout id="2" label="L" filenames="DE102020117768B4\_0018.png,DE102020117768B4\_0019.png" state="{{state}}">L</figure-callout>}}_2-\text{x}\right){/{\text{L}}_2)}^2-{\text{a}}_2(\left({\text{<figure-callout id="1" label="L" filenames="DE102020117768B4\_0021.png,DE102020117768B4\_0022.png" state="{{state}}">L</figure-callout>}}_1+\ddot{\text{u}}\text{L}+{\text{<figure-callout id="2" label="L" filenames="DE102020117768B4\_0018.png,DE102020117768B4\_0019.png" state="{{state}}">L</figure-callout>}}_2-\text{x}\right){/{\text{L}}_2)}^3$$

or $$\mathrm{a}\left(\text{x}\right)/{\mathrm{a}}_{\text{Max}}=(\left({\text{<figure-callout id="1" label="L" filenames="DE102020117768B4\_0021.png,DE102020117768B4\_0022.png" state="{{state}}">L</figure-callout>}}_1+\ddot{\text{u}}\text{L}+{\text{<figure-callout id="2" label="L" filenames="DE102020117768B4\_0018.png,DE102020117768B4\_0019.png" state="{{state}}">L</figure-callout>}}_2-\text{x}\right){/{\text{L}}_2)}^{\text{b}2}$$

drops again towards zero. In this case, L ₂ is 90% to 110% of L ₁ ; a ₁ and a ₂ independently equal 0.69 +/- 0.3; and b ₁ and b ₂ independently equal 1.96 +/- 0.3.

The partial bulge lengths L ₁ and L ₂ can both be equal to a half bulge length L, with the transition region being omitted, so that the transition length UL=0. 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 $$\mathrm{\alpha }\left(\text{x}\right)/{\mathrm{\alpha }}_{\text{max}}={\left(2-\text{x}/\text{L}\right)}^{\text{b}},$$

wieder gegen null abfällt, wobei a gleich 0,69 ± 0,3 ist und b gleich 1,96 ± 0,3 ist.Then the outer contour of the central body in the area of the contour bulge is set at a deflection angle α in relation to the smoothed base line of the outer contour of the central body, which corresponds to the coordinate x in the main flow direction from x=0 to x=L $$\mathrm{a}\left(\text{x}\right)/{\mathrm{a}}_{\text{Max}}=\left(1+{\text{a}}_1\right){\left({\text{x/L}}_1\right)}^2-{\text{a}}_1{\left({\text{x/L}}_1\right)}^3$$

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

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

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

falls back to zero where a equals 0.69 ± 0.3 and b equals 1.96 ± 0.3.

In any case, both equations defining the course of the deflection angle α over the first and second partial bulge lengths L _{1 and L 2 result in very similar contour bulges with fixed values of amax and L 1 and L 2 , respectively} . This is not the only reason why the fundamentally different equation for the course of the deflection angle α can also be selected via the second partial bump length L ₂ than via the first partial bump length L ₁ .

With the contour bulge of the supersonic inlet according to the invention, a favorable course of the wall pressure over the outer contour of the central body 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 monotonously with the smallest possible curvature |d ² p/dx ² | on.

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

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

In particular, the maximum deflection angle amax can be chosen to be equal to an acute angle at which the ramp of the outer contour of the central body in the main direction of inflow approaches the inner contour of the hood.

Typically, this acute angle is 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 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 partial bulge 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 outer contour of the central body and the inner contour of the hood transverse to the main direction of inflow 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 outer contour of the central body 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 the same in the H/L relationship κ α _max where κ equals 0.39 ± 0.04 and α _max is in radians.

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

Preferably, the variables a ₁ and a ₂ and b ₁ and b ₂ in the definition of the universal contour bulge of the supersonic inlet according to the invention are fixed 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 partial bulge lengths L ₁ and L ₂ , the contour bulge of the supersonic inlet according to the invention is universal over different Mach numbers of the inflow of the supersonic inlet. If, however, in the case of a supersonic intake, the point of impact of the compression shock emanating from the lip of the hood on the outer contour of the central body moves further with the Mach number of the inflow of the supersonic intake than can be compensated for by a transition area with a limited transition length ÜL, it can make sense to reduce the contour bulge overall in to shift the main direction of flow in such a way that the compression shock always hits the highest point of the contour bulge as precisely as possible. For this purpose, the contour bulge can be displaceable in the main direction of inflow, for example with a displacement device comprising a motor drive. This displacement device can also be designed in such a way that it displaces the contour bulge depending on the current flow conditions of the supersonic inlet in such a way that the compression shock emanating from the lip of the hood after the first partial bulge length at x=L ₁ or in the possibly adjoining transition area, i.e. up to x = L ₁ + ÜL, hitting the contour bump.

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

In concrete terms, the contour bulge in the supersonic inlet according to the invention can be formed on a bulge body which is displaceably guided on a main body of the central body along the base line.

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

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

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

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

The features mentioned in the patent claims and the description are to be understood with regard to their number in such a way that exactly this number or a larger number than the number mentioned is present without the need for an explicit use of the adverb “at least”. So if, for example, an element is mentioned, this is to be understood in such a way that exactly one element, two elements or more elements are present. The features listed in the claims may be supplemented by other features or may be the only features exhibited by the product in question.

character 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.

The invention is further explained and described below with reference to preferred exemplary embodiments illustrated 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 compression shock on a contour bulge of the supersonic inlet according to the invention and the wave structure induced thereby.
• 4 shows the course of the wall pressure over the contour bulge according to 3 along with a history of wall pressure in a reference case.
• 5 Figure 12 is a normalized deflection angle plot according to the contour bump 3 over their first buckling half length L.
• 6 is a plot of a specific course of the contour bulge according to 3 with the normalized deflection angle according to 5 .
• 7 is the closed 6 associated course of the deflection angle of the contour bulge according to 3 .
• 8th is a plot of the normalized wall pressure at the surface of the contour bulge according to 3 with a variation of the inflow Mach number M∞ between 2.0 and 3.0; and
• 9 is a plot of specific curves of further embodiments of the contour bulge.

FIGURE DESCRIPTION

the inside 1 The supersonic inlet 1 shown is designed for a main flow direction 2 . A central body 3 of the supersonic inlet 1 has an outer contour 4 of the central body. The central body outer contour 4 has a two-stage ramp 5 here, which begins upstream of a lip 6 of a hood 7 of the supersonic inlet 1 and ends with a contour bulge 8 . In the area of its downstream step, the ramp approaches an inner contour 9 of the hood 7 at an acute angle 18 to the main flow direction 2 . A compression shock 10 emanating from the lip 6 impinges on the outer contour 4 of the central body at different points depending on an inflow Mach number of an inflow of the supersonic inlet 1 in the main inflow direction 2 . A displacement device 11 for a bulge body 12 on which the contour bulge 8 is formed is controlled such that the contour bulge 8 is arranged in such a way that the compression shock 10 occurs at the highest point of the contour bulge 8 under the respective flow conditions. This and a special shape of the contour bulge 8 prevent a boundary layer 13 of the inflow from detaching from the central body outer contour 4 via a separation bubble due to a shock/boundary layer interaction of the compression shock 10 with the boundary layer 13 . In 1 the bulge body 12 forms the central body outer contour 4 of the central body 3 in an area in which a smoothed base line of the central body outer contour, above which the contour bulge 8 rises, has a constant radius of curvature. The bulge body 12 is guided on a circular path with this radius of curvature and is displaced by the displacement device 11 on this circular path.

At the supersonic inlet 1, which in 2 is shown and which is otherwise fully in accordance with the embodiment 1 corresponds, the displacement device 11 has a conveyor belt 15 with which the bulge body 12 can be moved here over a surface of the central body 3 with the course of the base line 14 . This also serves to always align the contour bulge 8 in such a way that the compression shock 10 strikes the highest point of the contour bulge 8 .

3 outlines the induced wave structure according to the impact of the lip 6 1 or 2 Outgoing compression shock 10 is formed at the highest point of the contour bulge 8 at its bulge half length. Compression waves 16 form upstream and downstream of the impact point of the compression shock 10 and expansion waves 17 emanate from the impact point of the compression shock 10 .

4 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 first rises steadily as the gradient increases and then decreases again to 0. The curvature of the pressure curve is minimal. This provides ideal conditions for avoiding a detachment bubble in the impact area of the compression shock 10 . The situation is different with the pressure curve, which is shown in 4 is shown with a dashed line and an impingement of the compression shock on the in 3 with a base line 14 also drawn in with a dashed line. This pressure profile has strong curvatures and therefore quickly results in the unwanted formation of a detachment bubble in accordance with the boundary layer 13 3 .

around the in 4 shown with a gradient of 0 beginning and ending and having a minimal curvature of the wall pressure, a coordinated curve of the maximum deflection angle α _max normalized deflection angle α of the contour bulge 8 is to be determined. The deflection angle α is the angle of the respective area of the central body outer contour 4 in the area of the contour bulge 8 in relation to the base line 14.

In 5 the normalized deflection angle α/α _max is plotted against the normalized longitudinal coordinate x/L of the contour bulge, where L is the half-length of the contour bulge 8 . The normalized deflection angle α/α _max begins at 0 and then increases with a 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 with regard 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 set equal to the angle 18 at which the ramp 13 to the main flow direction 2 meets the inner contour 9 approaching the hood. The bulge half-length L can depend on a smallest free distance h _e according to 2 and 3 between the central body outer contour 5 and the inner contour 9 of the hood 7 in the range of 0.025 h _e / α _max and 0.75 h _e / α _max .

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

7 shows the corresponding course of the deflection angle α in degrees 6 and 7 that the second half of the contour bulge 8 downstream of its crest, ie from L to 2L is a mirror-symmetrical image of its windward side from 0 to L.

8th shows a normalized pressure distribution p _W /p _Wmax of the wall pressure on the surface of the contour bulge 8 according to FIGS 4 until 7 at flow Mach numbers M _∞ of 2.0 and 3.0, which were predicted using characteristic methods (without boundary layer influence). p _Wmax corresponds to the absolute wall pressure at the trailing edge of the bulge, which is reached at the respective inflow Mach number M _∞ . It is obvious that the same contour bulge 8 according to the desired pressure curve 4 the negligible variations are realized with both considered inflow Mach numbers M _∞ . The contour bulge according to the invention is therefore optimized for different inflow Mach numbers and is therefore universal.

9 12 shows various embodiments of the contour bulge 8, which is not formed into two equal bulge half lengths L, but rather into two partial bulge lengths L ₁ and L ₂ with a transition region of a transition length UL arranged in between. The deflection angle α generally increases or decreases over the partial bulge lengths L ₁ and L ₂ in exactly the same way as over the two half bulge lengths L gel symmetry but also mean a different length of L ₁ and L ₂ . There is also the transition area 19. Here the deflection angle can remain the same up to a ridge 20, as is also present in the previously described embodiments of the contour bulge 8, and then suddenly change its sign at the ridge 20. The amount of the deflection angle α in the transition area 19 must remain ≦α _max . Alternatively, a rounded ridge 21 can be provided in the transition region 19, in which the deflection angle α steadily falls to zero and then increases again in magnitude after its sign change. In principle, it is also possible for a flattened ridge 22 with a deflection angle of zero to be provided in the transition region 19 . (In 9 the transition length ÜL is shown too long in comparison to the partial buckling lengths L ₁ and L ₂ . The transition length ÜL is no more than 10% of L ₁ and also no more than 10% of L ₂ , even if L ₁ and L ₂ differ. However, this difference between L ₁ and L ₂ also remains ≤ 10% of L ₁ .)

Reference List

1

supersonic enema

2

main flow direction

3

central body

4

central body outline

5

ramp

6

lip

7

Hood

8th

contour bump

9

inner contour

10

compression shock

11

shifting device

12

bump body

13

boundary layer

14

baseline

15

conveyor belt

16

compression wave

17

expansion wave

18

angle

19

transition area

20

ridge

21

rounded ridge

22

flattened ridge

a

deflection angle

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 with a lip (6) counter to the main flow direction (2), - with a transverse the free distance running in the main flow direction (2) 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 direction of inflow (2), and - the ramp (5) in the main direction of inflow having a contour bulge (8) on the outside of the central body contour (4) on which a compression shock (10) emanating from the lip (6) of the hood (7) impinges under at least one design inflow condition of the supersonic inlet (1), characterized in that the central body outer contour (4) in the area of the contour bulge ( 8) is set at a deflection angle α relative to a smoothed base line (14) of the outer contour (4) of the central body, which corresponds to a coordinate x in the main direction of inflow (2) over a first partial bulge length L ₁ from x=0 to x=L ₁ $$\mathrm{a}\left(\text{x}\right)/{\mathrm{a}}_{\text{Max}}=\left(1+{\text{a}}_1\right){\left({\text{x/L}}_1\right)}^2-{\text{a}}_1{\left({\text{x/L}}_1\right)}^3$$

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

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

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

falls back 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 +/- 0.3 are. 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 direction of flow (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 partial bulge length L ₁ : $$0.025{\text{H}}_{\text{e}}/{\mathrm{a}}_{\text{Max}}\le {\text{L}}_1\le 0.75E{H}_{\text{e}}/{\mathrm{a}}_{\text{Max}}$$

, where h _e is the smallest free distance between the central body outer contour (4) and the inner contour (9) of the hood (7) transverse to the main direction of flow (2). Supersonic inlet (1) according to one of the preceding claims, characterized in that the first partial bulge length L ₁ is at least as large as a length of a separation bubble in the main direction of inflow, which extends under the at least one design inflow condition above the smoothed base line (14) of the central body outer contour (4th ) forms without the contour bump (8). Supersonic inlet (1) according to any 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) by a displacement device (11) comprising a motor drive. Supersonic inlet (1) after claim 7 , characterized in that the displacement device (11) is designed such that the contour bulge (8) depending on the current flow conditions of the supersonic inlet (1) is shifted in such a way that the compression shock (10) emanating from the lip (6) of the hood (7) occurs in the transition area or, if ÜL is equal to 0% of L ₁ , at the first bulge half length x=L ₁ meets the contour bump (8). Supersonic inlet (1) after claim 7 or 8th , characterized in that the displacement device (11) is designed in such a way that it controls the contour bulge (8) depending on signals from a first pressure sensor in the area of the first half-length of the bulge and a second pressure sensor in the area of the second half-length of the bulge in such a way that a pressure difference between wall pressures is of the central body outer contour (4), which are detected by the first pressure sensor and by the second pressure sensor, is minimized. Supersonic inlet (1) according to one of Claims 7 until 9 , characterized in that the contour bulge (8) is formed on a bulge body (12) which is guided displaceably on a main body of the central body (3) along the base line (14). Supersonic aircraft with an engine with a supersonic intake (1) according to one of the preceding claims.

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