CC BY
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AERODYNAMIC AIRFOIL AT CRITICAL ANGLES OF ATTACK
Cedomir Lj. Kostica, Bosko P. Rasuob
a General Staff of the Serbian Army, AF&AD, 204.vbr, Batajnica, Republic of Serbia, e-mail: cedokostic@gmail.com, ORCID iD: ©http://orcid.org/0000-0001-9940-3560 b University of Belgrade,Faculty of Mechanical Engineering, Belgrade, Republic of Serbia, e-mail: brasuo@mas.bg.ac.rs, ORCID iD: ©http://orcid.org/0000-0002-0912-6844
DOI: 10.5937/vojtehg64-8048
FIELD: Mechanical Engineering ARTICLE TYPE: Review Paper ARTICLE LANGUAGE: English
Summary:
Aircraft construction experts must not neglect the behavior of aircraft in extreme or closely extreme flight conditions, such as flights at critical angles of attack, where a normal flight can be easily converted into a stall.
This paper highlights the essential factors that influence the behavior of aircraft in flight at critical angles of attack. Based on the available experimental results and estimations, the performances of airfoils were analysed depending on air flow conditions (categorized according to Mach and Reynolds numbers), airfoil shapes, dynamics of the transition of angles of attack, description of the flow around airfoils with increasing the angle of attack upon reaching a critical value, and the effect of roughness of airfoil surfaces at critical angles of attack.
The paper gives a physical interpretation of a lift decrease and a stall. It minutely describes the origin of flow separation and categorizes airfoil sections by type of separation and their behavior during the flow at the critical angle of attack.
Based on modern aerodynamics, this paper aims to show and explain the issues and the most important characteristics of the flow past the body at critical angles of attack and give practical recommendations for airfoil design. As such, it may be of interest to pilots and engineers as well as to educational and research institutions.
Key words: aerodynamic, performance of airfoils, fluid dynamics, airfoil shape, critical angles of attack.
Introduction 5
oo
A value of the lifting force of the wing (of a body in flow) obtained at
the moment when further increase of the angle of attack (a) does not ^
increase it any more is called the maximum lift force of the body for given o
flow conditions (Gretchikhin, 2014). The angle of attack at which ^
separation of the laminar flow occurs on the curverepresents the point of o
falling of the CL/a curve angle (chart CL/a Figurel). jS
Figure 1 - Change of the lift coefficient Cl when increasing the angle of attack a Рис. 1 - Изменение коэффициента подъемной силы Cl при увеличении угла атаки a Slika 1 - Promena koeficijenta uzgona Cl sa povecanjem napadnog ugla a
A value of the angle attack at which the lift coefficient has the maximum value represents the angle of the maximum lift. A further increase of the attack angle causes that the lift force decreases slowly in the beginning, then more steeply. For certain values of the angle of attack (акр), the lift force is lost.
The maximum lift depends on various factors:
- Airfoil shape (of a body in flow);
- The flow conditions (modeled by Reynolds and Mach numbers);
- The shape of the wing (of a body in flow), slenderness, the position of the wings in relation to the flow and the aircraft fuselage (the angle of the arrow, curl wings);
- Aerodynamic additions to wings;
- The engine; and
- Flight dynamics of the body.
The value of the maximum lift coefficientis a complex function that depends on the geometric size of the airfoiland the Reynolds number; therefore, it cannot be calculated analytically. The wind tunnel testing has led to the empirical equations that apply to a specific area of the Reynolds number, so that usual airfoils for approximate calculations can be determined from the equation:
Cz max = 1.73 - 2.4
> -14 ( d L+5« ( d,
The above formula applies to the value of Re=3106, while for other values of the Reynolds number it is necessary to make corrections.
An approximate value of the critical angle of attack is obtained from the equation:
«max =(11 + «) + 50 fy) + 701
V J max
As shown in Figure 2, Cix is from 1.5 to 1.6, i.e. less than 10% of the theoretically possible values (Cone,1985).
Figure 2 - Potential maximum lift force coefficient Рис. 2 - Потенциальное максимальное значение коэффициента подъемной силы Slika 2 - Potencijalni maksimalni koeficijent uzgona
With aerodynamic accessories (slats and flaps) and an additional impact of jet (engine), the values of about 30% of the theoretical value are obtained.
The physical interpretation of a stall
An increase in the angle of attack observed along the airfoil leads to the following:
- Large vacuum near the leading edge;
- Strong positive pressure gradient between the negative peak and the trailing edge;
- Increase of the boundary layer along the upper side of the airfoil.
An increase of the angle of attack of the flow at the airfoil cross-
section leads to two critical points in the stream where the separation of flow can be expected:
1. At the leading edge, where a streamline must flow around the attack "nose'', which corresponds to a loss of momentum;
2. In the vicinity of the trailing edge, where there is a noticeable increase in the thickness of the boundary layer.
The loss of lift due to the separation of flow occurs in one of these points or at both points at the same time. In Figure 3, the CL-a diagram (Kostic, 2010) shows three specific ways of flow separation.
C,
1.4 -
1.2
1.0
0.8
0.6
0.4
0.2
аеропрофил:
NACA 63-018(одва)а№е од излазне ивице NACA 63~012(05Baja№e на нападноj ивици)
NACA 63-009(комбиновани одва)а»е стру}а№а) симетричан клин(одва)а на Hen^Hoj ивици.
дугачак мехур)
12
20
Figure 3 - Types of the separation flow of the A-profile, Re= 5.8 • 10 Рис. 3 - Виды срыва потока на А-профиле,Re=5,8• 106 Slika 3 - Tipovi otcepljenja strujanja A-profila, Re=5,8406
The selection included the airfoils whose CL-a curve values coincide up to a~60, after which there is a characteristic discrepancy.
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a
0
4
8
Separation from the leading edge - elongated bubble shapes
Separation from the leading edge isgradually formed on thinner airfoils with relatively sharp leading edges, as a form of an elongated laminar separation bubble;
Adouble-fold curve, characteristic for flow separation at the leading edge in a shape ofelongated bubbles, ca be noticed. Lift rises, almost linearly, then gently falls and remains constant for values a ~ 8.5-10°, then the lift force continues to decline slightly, followed by an increase.
transition
turbulent wake
constant pre s suis
-1
indicates the sepaiation —V
transition
, oj=e CHU Crf=<№7 Ch=0j85 cpi
+1
Figure 4 - Example of stalling a thin, sharp airfoil Рис. 4 - Пример сваливания тонкого стреловидногокрыла Slika 4 - Primer prevlacenja tankog, ostrog aeroprofila
The reason for this behavior of airfoils lies in the short distance between the stagnation point (when a>0, it is located on the underside) and the leading edge, so that the boundary layer in this zone is very thin.
The Reynolds number is low for the given case.We take into account the existence of a strong negative pressure gradient between the stagnation point (where Ср=+1) and the leading edge (where probably Ср=-10). The airflow over the sharp leading edge is laminar in the boundary layer until abrupt flow separation (McCullough, 1955). Namely, the usual transition to turbulent flow inside the bubble is not typical for this case. After flow separation, as shown in Figure 4, the boundary layer next to a newly formed bubble becomes significant space created by mixing with the fluid within the
separation bubble. The result of this mixing is the reduced thickness of the separation bubble as the flow moves away from the leading edge to the place where it again joins the boundary layer of the upper airfoil.
The presence of bubbles is recognized through pressure distribution along the airfoil, Figure 4, presenting a constant area of the negative pressure coefficient, where, if the bubble did not exist, there would occur a sudden change in pressure (peak on the curve), which in this case does not exist.
Figure 5 - Lift, drag and the separation bubble characteristic for the NACA 64A006 airfoil a) Lift with discontinuity at Cl=0.56 b) Section drag showing sudden increase at Cl=0.56 c) Length of the separation bubble from the leading edge Рис. 5 - Подъемная сила, сопротивление и область отрыва
а) Разрыв непрерывности подъемной силы при Cl=0,56 b) Резкий скачок коэффициента сопротивления при Cl=0,56 c) Длина отрывного пузыря упередней кромки Slika 5 - Uzgon, otpori, separacioni mehur za aeroprofil NACA 64A006 a) Diskontinuitet uzgona pri Cl=0,56 b) Naglo povecanje koeficijenta otpora pri Cl=0,56 c) Duzina mehura otcepljenja od napadne ivice
In the diagram CL-a, CD, x/c (bubble) for the NACA 64А006 airfoil shown in Figure 5, it can be noted that in the section «6% x/c for CL=0.56 acurve discontinuity appears, showing the moment of the creation of initial separation and a bubble, which is reflected in a sudden increase of the drag coefficient.
With a further increase of the angle of attack, the area affected by the separation bubble grows proportionally to a critical moment which
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corresponds to the maximum lift, after which it extends along the entire length of the upper airfoil, followed by the reduction in lift and transit to a stall. In the previous case, the separate bubble extends to 60% of the airfoil chord (x/c), then the increasing of a leads to the 75% of the airfoil chord (x/c).
Separation from the leading edge - short bubble shapes
Short bubble shapes occur as sudden separation when flow passes airfoils with a sharp leading edge, or at flow past a low curved airfoil;
For airfoils characterized by an abrupt onset of a small flow bubble, it is characteristic, as shown in the diagram for the NACA 63-012 airfoil (Figure 3), that the lift force increases almost linearly up to the maximum and then falls sharply.
Separation from the trailing edge
Separation from the trailing edge is characterized by slight gradual separation. It starts from the trailing edge and moves towards the leading edge, proportionally to the increase of the angle of attack. The NACA 63-018 airfiol, Figure 3, with a rounded leading edge, is a typical representative of the airfoil group in which separation occurs at the trailing edge. It is characterized by an almost linear increase of lift, then a gradual transition to lift loss and complete flow separation. The length at which a separation bubble is formed depends primarily on the airfoil shape. In the case of flow past the airfoil, Figure 4, the length extends up to 60% of the airfoil chord, while with further increase of the angle of attack separation occurs starting from the airfoil trailing edge. A stall occurs after the moment when the separation begins from the trailing edge and comes to a point of 25% of the leading edge.
For airfoils used for subsonic flows, the length of the separation bubble varies from 15% to 25% of the length of the airfoil chord.
If a rounded leading edge is used, and the airfoil thickness and curvature are increased, the risk of separation bubble formation on the leading edge is eliminated.
The next critical zone for flow separation due to an increased angle of attack is near the trailing edge.
Figure 6 (Pinkerton, 1996) introduces the distribution of pressure along the airfoil at the time immediately prior to reaching the maximum coefficient of lift (CLX).
The distribution of the pressure coefficient depends on the shape of the airfoil, Reynolds number, angle of attack, relative thickness, curvature of the center line and the relative curvature of the leading airfoil edge.
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Figure 6 - Characteristics of the lift force of the NACA 4412 airfoil and the distribution of —-
Cl, Cn, Cm ~
Рис. 6 - Характеристики подъемной силы профиля крыла NACA 4412 ,
и распределение Cl, Cn, Cm Slika 6 - Karakteristike uzgona aeroprofila NACA 4412 i raspodelaCbCw,Cm £
s o K
Let us consider the values of the local maximum speed using the equation:
w/V + 2t/c + 0.2 CL/(r/t) (3)
where r/t is the ratio of the radius of the leading edge and the airfoil thickness (typically between 0.5 and 2.0 t/c). A possible maximum local velocity w, the maximum dynamic pressure q and the minimum pressure coefficient can be obtained from this equation.
Both parts of the right side of equation (3) produce a positive pressure gradient while approaching the trailing edge of the airfoil. Under the influence of this gradient, the boundary layer increases significantly and the trailing edge has a total thickness of 5, which can be written as:
5c = 5Cf+ kfCO"
where Cf~ 0.03 is the coefficient of friction resistance, n=2, the coefficient k of the order of 10.
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After the boundary layer thickening, it reaches the critical moment when its movement against pressure cannot continue and flow separation occurs in the accumulated boundary layer. As a result, there has been a cessation of circulation increase with the increase of the angle of attack and the lift reaches its maximum.
Figure 6 shows the NACA 4412 airfoil in the CL,Cm-adiagram. Before reaching its maximum, the lift coefficient diagram diverges from the straight line, initiating flow separation from the trailing edge. An increase of the angle of attack through CLX results in a gradual lift decrease, without discontinuity, down to the angle of attack after which a complete separation flow occurs, which corresponds to the scenario when the separation from the trailing edge is extended to the leading edge.
Airfoils with circular sections
Let us consider the behavioral characteristics of the lift of the airfoil section with a circular shape, with a curvature of 10%, for three values of the Reynolds number, shown in Figure 7 (Kostic, 2010).
When the airfoil lower sideis flat, the degree of curvature is equal to 0.5 of the airfoil slenderness. The maximum thickness located at 50% of the length of the chord gives airfoils with the prominent wedge-shaped trailing edge at the site of potential vacum, which easily leads to initial separation. Symmetric thickening with the maximum curvature in the middle of the airfoil, combined on the side with lower pressure, increases sensitivity towards separation flow while approaching the trailing edge.
It can be noticed that, under the sameflow conditions, the behavior of the airfoil depends on the Reynolds number, therefore:
1. When the value of Re=105, the laminar type of separation occurs. The point with the minimum pressure corresponds to the half of the airfoil tendon. Under these conditions, the minimum drag coefficient is 0.02, which confirms the separation of flow.
2. When the value of Re=3105, at the angle of attack corresponding to the lift coefficient of 0.575, which is identical to the theoretical CLsym=0.115(f/c)%.
The value of the drag coefficient (CD5=0.008=min) clearly confirms that this is a completely clean, pressed flow, which represents in this case optimum.
Increasing the Reynolds number over Re=3105 gradually reduces the lift coefficient for a minimum resistance value and for Re=5106 the lift coefficient is CLO=0.46. In the separation layer, turbulent flowappears.
tuibuJfiïit baimdary layer optyriiaiiy
liiïujw sepaià+iOM
CI
I—1 I—1
0J&
as ■ fff
I f/w'
■10 . J 0 10 20 3Î:
ш - J
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Figure 7 - Characteristics of lift and the mechanism of stalling an airfoil with 10% of curvature Рис. 7 - Характеристики подъемной силы и механизм обтекания крыла с 10% -ой
кривизной профиля Slika 7 - Karakteristike uzgona i mehanizam prevlacenja aeroprofila sa 10% zakrivljenosti
For the observed cases, Figure 7, the graph of the lift coefficient decreases in all three cases. The angle of zero lift at Re=105h as a value of -30; while an increase of the Reynolds number to Re=3105 results in a decrease of the angle of zero liftto the value of -60. However, at the value of Re=5106, the angle of zero lift increases to -50.
This variation of the angle of zero lift value depends on the coefficient of the resistance to friction from the lower value of 105 for laminar flow separation, through the part with turbulent separation for which the minimum drag coefficient (optimum) is valid, to the value where turbulent separation is dominant, i.e. a value of 5106.
The curvature of the leading edge of the observed airfoils is possible to reduce the drag coefficient, so the curvature of r/c«1-2% gives a reduction in the drag coefficient of CDS=0.038 atCDs=0.017 for Re=4106.
Figure 8 - Optimum lift coefficient as a function of the airfoil camber, t/c = 6-12% Рис. 8- Оптимальный коэффициент подъемной силы хорды профиля крыла,
t/c=6-12%
Slika 8 - Optimalni koeficijenti uzgona u funkciji srednje linije aeroprofila, t/c=6-12%
The diagram in Figure 8 shows the influence of the airfoil camberand the position of the airfoil maximum thickness to the values of the lift coefficient. The diagram is obtained using the equation CLsym=0.115(fc)% and compared with the graphs obtained by equations experiments, where it can be concluded that the optimal position of the maximum thickness of subsonic symmetric airfoils is between 30-40% of the mean airfoil chord.
Maximum lift depending on the shape and Re
Changing the boundary layer and the drag coefficient is directly dependent on the Reynolds number and roughness of the flowsurface. The loss of lift of flow around the body directly depends on the body's shape. The transition from the laminar flow of the boundary layer into the turbulent one can be caused by an external factor (turbulence of the flow field, turbulence inside the wind tunnel, etc.) or by increased surface roughness of the leading airfoiledge.
CLX as a function of the shape of the leading edge
Figure 9 shows the dependence of the maximum lift coefficient (Cu) on the leading edge radius (r/c) of symmetric airfoils.
Figure 9 - Maximum lift of symmetric profiles, depending on the leading edge radius r/c Рис. 9 - Максимальная подъемная сила симитричного профиля крыла в зависимости от радиуса закругления его передней кромкиг/с Slika 9 - Maksimalni uzgon simetricnog aerodinamickog profila, u zavisnosti od radijusa
napadne ivice r/c
The radius which produces the largest lift coefficient is between 1.5 and 2.0% of the tendon (r/c=1.5-2.0%) for the airfoils with t/c=6-18%.
For NACA 0012, NACA 0018 and NACA 63-012 airfoils shown in Figure 10 (Hoerner, Borst, 1985), it is evident that, with an increase in the r/c, the value of Clx grows. From a value of r/c>1.6%, the gradient of the curve decreases.
It should be considered that, as the curvature of the leading edge of the airfoils increases, the drag coefficient also increases, as shown in Figure 10 (Kostic, 2010), which displays the values of CL and CD for the airfoils of relative thickness of t/c=6% and r/c=0.81% and r/c=0.24% depending on the change in the angle of attack. It is noticeable that the drag coefficient increases as the angle of attack increases.
Airfoils with rounded leading edges have a greater coefficient of lift and a lower drag coefficient at the same angles of attack, so it is only for a«18° that the ratio is CL/ CD<1.
r/c=0,024%
NACA 65A006
20 с«"0)
1,0 Cl
Figure 10 - Lift and drag t/c = 6% of airfoilswith different leading edges, Re = 5 ■ 10 Рис. 10 - Подъемная сила и сопротивление t/c = 6% профиля крыла с различными
передними кромками, Re=5105
Slika 10 - Uzgon i otpor t/c=6% aeroprofila razlicitih napadnih ivica, Re=5105
It is noticeable that the values of the CDmin are of approximately same values, and that flow separation at lower values of the angle of attack has resulted in a rapid increase of the drag coefficient for the airfoil with a sharper leading edge.
The mechanism of separation of flow around a thin airfoil can be considered the same as the one in airfoils with sharp leading edges, especially for the values of Re<105 (von Karman, 2001).
Figure 11 shows the dependence of the curvature of the leading edge of the airfoil (r/c) and Re. It can be seen that the methods of flow separation can be grouped depending on the values (r/c) and Re. For the curvature of the leading edge from 0.25% to 1.60%, it can be concluded that, at Re<106, separation flow corresponds to the model of thin airfoil separation, while for Re>2• 106 separation flow corresponds to the model of thin airfoil separation and trailing edge separation.
Figure 11 - Stall domains due to the leading edge radius r/c and Re Рис. 11 - Модель срыва потока в зависимости от закругленности передней кромки
r/c и значения Re
Slika 11 - Model otcepljenja strujanja u zavisnosti od zaobljenosti napadne ivice r/c
i vrednosti Re
A typical example of application of sharp leading edges is seen when creating ship propellers, where the value of the Reynolds number is below the critical value of Re<105.
The curvature of the airfoil leading edge sectors
Also, Clx is a function of the curvature of the airfoil leading edge sector (camber line of the airfoil), f/c, and the shape of the leading edge, i.e. r/c. Around the thin airfoil leading edge, a flat plate, flow bends from the lower side through the leading edge to the upper side. At larger Reynolds numbers, laminar separation and the bubble effect of the boundary layer can be reduced.
Figure 12 shows the change in the lift coefficient of airfoils with sharp leading edges for different values of the curvature of the center line of the airfoil in the conditions of flow of similar Reynolds numbers.
З.п - Ch
leading edge separation
a) thin sirroil
standard thicknesses aii foil
vu. -t-jï.c-.: :
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Figure 12 - The maximum lift coefficient of an airfoil with a sharp leading edge, as a function of the degree of curvature of the camber line of the airfoil Рис. 12 - Максимальный коэффициент подъемной силы крыла с острой передней кромкой, как функция степени округленности средней линии профиля Slika 12 - Maksimalni koeficijent uzgona aeroprofila ostre napadne ivice, kao funkcija stepena zakrivljenosti srednje linije aeroprofila
It can be said that there is no difference in the examples shown in terms of the location of the maximum curvature of the camber line of the airfoil for the maximum lift between 30% and 50% of the chords of the airfoil. This is illustrated by graphically presenting the CLX/(f/c) relations, obtained by testing airfoils in the wind tunnel, (Doenhoff, 2015) Figure 13.
Figure 13 - The maximum lift coefficient, relative maximum lift coefficient, the relative thickness of airfoil 6-9% and 12-24% for Re, f = 4 • 106 Рис. 13- Максимальная подъемная сила, относительный максимальный коэффициент относительной толщины профиля6-9% и 12-24%, для Re, f = 4 • 106 Slika 13 - Maksimalni koeficijent uzgona, relativni maksimalni koeficijent uzgona, relativne debljine aeroprofila 6-9% i 12-24%, za Re,f = 4 • 106
After the adjustment to A = 5 in accordance with the equation (Kostic, 2010):
(f/c)%>»(1 .4+0.2)Cl=1.6%Cl
the characteristics of thin airfoils, with t/c=6 and 9%, are shown in Figure
13, which shows that, for the degree of curvature of the camber line of the airfoil between -1% and +1, the maximum lift coefficient corresponds to the critical angle of attack of separation flow (of 300).
Practical use of thin airfoils and the curvature of sections in aviation is limited by structural conditions (providing the necessary wing capacity for the required airfoil thickness), but they are applicable in the construction of turbine blades and fan blades.
The coefficients of the maximum lift decrease with the increase of the maximum relative thickness of airfoils of 12% or more are shown in Figure 13. The maximum lift coefficients grow as a function of a relative airfoil curvature, to the airfoils with the highest tested curvatures, those of 10%.
The curvature of the nose (leading sections of airfoils)
The study of the NACA 63012 airfoil, Figure 14, at Re=5(10)6 and M=0.18, leads to the conclusion that at Clx = 1.36 separation flow occurs suddenly, with strong shocks and shaking for angles corresponding to the values immediately prior to the occurrence of separation. This can also be anticipated by observing the diagram in Figure 8.
Increasing the radius of the leading edge nose of 1.1% to 2% of the length of the camber aerodynamic chord, makes CLxgrow to a value of 1.50, while a sudden stall is still present.
To avoid a sudden stall, the airfoil leading section was folded, Figure
14. The values of CLx grow again, but a sudden stall is still present. Separation occurs following the model of a thin airfoil with a separation bubble which does not change from the initial shape to the moment when it suddenly increases and extends towards the trailing edge. The distribution of the pressure coefficient shows that separation occurs at
Cpmin 10.
Figure 14 - Modifications to the leading edge, NACA 63-012, Re = 4.9 • 106 Рис. 14 - Модификация передней кромки, NACA 63-012, Re= 4,9 • 106 Slika 14 - Modifikacije napadne ivice, NACA 63-012, Re=4,9 • 106
Changing the curvature and the radius of the airfoil leading edge increases the value of the maximum lift approximately, but it does not change the separation mechanism. As a sudden entry intoa stall is unfavorable, it is recommended that a combination of the relative thickness and the airfoil curvature is used to transfer the mechanism of sudden separation occurrence in the attack sector to gradual separation which is a characteristic of the separation from the trailing edge.
The critical Reynolds number
When choosing airfoils (any aerodynamic form), it is important to know the values of the Reynolds number for which they can be used. The critical Reynolds number at which turbulent flow occurs in the boundary flow field of the upper wing surface for thin airfoils is Re=105, Figure 15 (Kostic, 2010). It is noticeable that airfoils with a smaller curvature have higher values of the critical Rewith an increase in CLX.
Clx
AVAJoukovski Interpol 9%,t=1,2 NACA64-210 A=6, 1=1,0 NACA64-00,t=1,8 DVL9%-0,83, A=5,t=1 RAF-34Han ивица (10/15)%,t=2 RAF-34,1=2,3
Re
Rekrit
ради]ус носа нападне ивице
r/c(%)
Figure 15 - Maximum lift of rectangular Figure 16 -Critical Re as a function of the
wings as a function of Re, t/c = 8-10% Рис. 15 - Максимальная подъемная сила правоугольного крыла как функция Re,
t/c = 8-10% Slika 15 - Maksimalni uzgon pravougaonih krila kao funkcija Re, t/c = 8-10%
radius at the leading edge of the airfoil Рис. 16 - Критический Re, как функция радиуса закругления передней кромки Slika 16 - Kriticni Re, kao funkcija radijusa napadne ivice
Airfoils such as the NACA 0009 have the same critical way of changing the image of the flow at higher values of Re. The transition from laminar to turbulent flow occurs at ten times higher values of Re than in airfoils with higher relative thickness and the radius of the leading edge.
The radius of the leading edge with the increase decreases the values of the critical Re, as shown in Figure 16 (von Karman, 2001). Airfoils with smaller curvature have a higher value ofthe critical Re with an increase in CX. For larger Reynolds numbers (106 to 107), they represent aircraft at low speed during landing and takeoff. TheC^ values are shown in the diagrams above.
The characteristics of airfoils for Reynolds numbers in this range are:
- Flat planes and airfoils with thickness up to 3% have the same approximate maximum lift coefficient, with separation characteristic for thin airfoils;
- Thin airfoils (t/c=5-10%) are characterized by increasing Clx as their camber line curvature increases from 0% to 2% (Figure 16). Loss of lift varies from sudden to gradual, depending on the shape and dimensions of the airfoil and on flow past an airfoil conditions;
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As it can be seen in Figures 16 and 17, airfoils with a relative thickness between 8 and 12% and the curvature of the camber line between 0% and 2% have a strong tendency towards increasing the maximum lift. Increasing the angle of attack leads to the characteristic flow around the leading edge and to a gentle transition from laminar to turbulent flow.
For airfoils with larger curvature and with aprominently rounded leading edge, flow continues around the leading edge, while the maximum lift coefficient is in a function of the angle of attack at which the laminar boundary layer extends towards the trailing edge. CLX depends on the coefficient of friction along the profile of the upper side of the airfoil. For Reynolds numbers whose value exceeds 107, the drag coefficient increases as a result of an increase in the thickness of the boundary layer.
It is noticeable that, depending on the tendencies of Clx change, different types of separation flow occur:
- Airfoils with separation flow of the thin airfoil type are characterized by a fairly constant value of the maximum lift coefficient;
- Airfoils with rapid separation flow tend to increase CLX with an increase of the Reynolds number;
- Airfoils in which the separation bubble is formed towards the airfoil trailing edge with increasing Re grows slowly, so that the maximum lift coefficient decreases for cases with larger curvature of the camber line.
Theoretical analysis
In the early theoretical studies of the boundary layer around the leading edge of the airfoil, the maximum lift was analyzed as a function of the Reynolds number and turbulence flow for thin and curved airfoils.The theory was applicable for Reynolds numbers at which there occurs separation flow characteristic for separation in the leading edge characterized by a laminar separation bubble.
For the case where separation is in relation to the trailing edge, the maximum lift coefficient is the result of the balance between kinetic and pressure forces. Potential energy is represented by an increase in pressure between the point of minimum pressure and the trailing edge. For sections with a pronounced curvature and a rounded leading edge, separation can be determined as a function of the kinetic energy inside the boundary layer. Therefore, while the kinetic energy inside the boundary layer is a function of the Reynolds number and grows in line with the decline in the value of the frictional drag, it can be said that:
where n=0.5(1/6)=1/12 and 1/6 is the equivalent ratio during turbulent flow. On the basis of the aforementioned, the above equation can be applied to the value of Re<107.
The relative thickness of airfoils
For Reynolds numbers of 4(10)6,Figure 17 shows the dependence of the maximum lift coefficient and the relative thickness of the airfoil.
Clx
RAF-89 (25/1,96) F=2, A=6 AVA 0025-40, f=1, A=5
Re
10-
10'
10
Figure 17 - Airfoil maximum lift, as a function of t/c, at Re = 4 • 10 Рис. 17 - Максимальная подъемная сила крыла, как функция t/c, при Re = 4 • 106 Slika 17 - Maksimalni uzgon aeroprofila, kao funkcija t/c, pri Re = 4 • 106
The maximum lift coefficient is rising sharply in line with the relative curvature of the camber line of the airfoil between 0% and 2%, and for the relative thickness of t/c=5-8%. For values of t/cover 8% and //cover 4%, the airfoilis characterized by stalling as in thin airfoil. The peaks of the curves can be obtained for values t/c«10%. The source in literature stated that during the experiment, for the values of t/c=12%, airfoils showed more efficiency, especially at high subsonic speeds.
Airfoils with t/c=(0 до 10)% are most common, and thus most tested, representing a compromise between the aerodynamic and structural demands of an object in an air stream.
Depending on structural needs, the relative thickness of airfoils increases; consequently, the value of t/c«18% is characteristic for the root of the wing of subsonic aircraft. Over the value of t/c=18%, the maximum lift coefficient gradually declines; therefore, a stall is gradual and predictable.
2
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IvS
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Itr-J
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very rough surface
smooth surface
very rough surface
Figure 18- Maximum lift of airfoils of series 63 and 64, Re = 6 • 10 , as a function of the relative thickness and curvature of the airfoil, for A = 5-6 Рис. 18 - Максимальная подъемная сила крыла серии 63 и 64, Re = 6 • 106, как функция толщины аэродинамического профиля и кривизны хорды, для А= 5-6 Slika 18 - Maksimalni uzgon aeroprofila serija 63 i 64, Re = 6 • 106, kao funkcija relativne debljine aeroprofila i zakrivljenosti srednje aerodinamicke tetive, za A=5-6
High-speed airfoils
Increasing the relative thickness of the airfoil and increasing the camber line lead to the increase of the torque of the resulting aerodynamic forces; therefore,a special attention is given to airfoil sections exposed to high speeds.
The position of the maximum camber is not that important for the maximum lift value, it is important that it is between 0.30 and 0.50 of the aerodynamic chord. On the other hand, the position of the maximum thickness has a significant impact, especially if the radius of the leading edge changes at the same time.
Figure 18 indicates the relation between the maximum lift coefficient and the relative thickness of the airfoils of the NACA airfoil series 63 and 64 and Re=6(10)6.
The influence of surface roughness
In theoretical studies of aerodynamic characteristics of an object in an air stream, the object surface is taken to be absolutely smooth. However, in real conditions, there is not an absolutely smooth surface. The following examples will show how uneven airfoil skin influences the occurrence of separation flow, i.e. the value of the maximum lift.
Increasing the roughness of the upper side leads to increased lift for a given angle of attack. On the other hand, increasing the roughness of the lower side reduces the lift for a given angle of attack. The increase of roughness directly affects the thickness of the boundary layer, as well as the flow in it.
Figure 19 - Airfoillift, A = 6, R = 6 • 10 for smooth and rough surfaces Рис. 19 - Подъемная сила крыла, А = 6, R = 6 • 106для гладкой и шероховатойповерхностикрыла Slika 19 - Uzgon aeroprofila, A=6, R = 6 • 106 za glatke i hrapave povrsine
Figure 19 shows an CL/a diagram of the airfoil observed under the same conditions, but with a different state of skin. It is clear that, with increasing roughness of the skin, the maximum value of the lift coefficient reduces together with a slight decrease of the lift gradient. It is noticeable that in this case airfoils behave similary after the decline of lift.
In real conditions, roughness is caused by aircraft production technologies and as such can be controlled and avoided on desired surfaces. During flight, there is damage to aerodynamic surfaces (to a lesser extent due to sand or insects and to a a greater extentdue to impacts of birds, which can lead to significant changes in the airfoil geometry).
A special case is the appearance of ice on leading edges, as a result of specific meteorological conditions of flight. Figure 20 shows the results of flow past the NACA 2212 airfoil, at Re=4(10)4, with and without the effect of ice (Kostic, 2010).
Figure 20 - Cl, Cd with and without simulation of ice on the leading edge of the NACA
2212 airfoil
Рис. 20 - Cl,Cd с и без симуляции льда на передней кромке крыла NACA 2212 Slika 20 - Cl, Cd sa i bez simulacije leda na napadnoj ivici za aeroprofil NACA 2212
Ice on the leading edge of the airfoil significantly reduces the value CLmax from 1.3 to 1.0; it increases the coefficient of resistance due to the increased turbulence of the boundary layer and increases the frictional drag.
In addition to the changes of the final values of the CL and CD, there are changes of their functions as well. TheCD/CL2 ratio is affected, causing directly a change in the position of the aerodynamic center forward, which in flight requires an additional deflection of the rudder height, and thus a change of the angle of attack, which further complicates the problem of the spread of ice-affected areas.
The aerodynamic characteristics of the airfoil after stalling
Reaching the maximum lift with a further increase in the angle of attack leads to a faster or slower lift decline (depending on the observed flow past airfoils and its conditions). When the pitch-stall, or when a propeller pitch is increased, and also when there is a need for a reversible propeller pitch, the angle of attack increases.
In Figure 21, the behavior of the airfoil for given conditions of a=0° up to a=180° is shown (Critzos, 1985).
Figure 21 - Variations of the lift coefficient for a = 0°-180° Рис. 21 - Вариации значений коэффициента подъемной силы при a=0o-Slika 21 - Varijacije koeficijenta uzgona za a=0o-180o
180°
The value of the airfoil lift coefficient is similar, but it can be seen
that:
- After separation, the resultant force is approximately perpendicular to the chord;
- Lift force reaches the second maximum at a«45° ( CD~1.0) so that the normal force is Crf\ 2=1.43.
- When a=90° , CL~0 and CD~ 1.8, which is close to the value of the drag for a flat plate of the same angle of attack (CD«1.95);
- For the value of the angle of attack between 90° and 180°, it is characteristic that the sharp trailing edge has become a leading one, and the rounded leading edge has become a trailing edge, which is characterized by increased lift and reduced drag, but with different values and similar behavior.
Influence of the dynamics of the stall
In the previous analyzes of a stall, we started from the assumption that the change of the angle of attackis is gradual, so that we have considered stationary flow. But in real conditions of exploitation, changes of the angle of attack do not correspond to this assumption, especially if working conditions of helicopter rotors are taken into account, or parts of wing machinery during sudden maneuvers.
Figure 22 illustrates the changes in the lift coefficientat gradual and sudden changes in the angle of attack (Kostic, 2010).
Cn^Cl
Cl
f=16,4Hz k=0,062 Aa=4,83o
M
Figure 22 - Dynamic effects of changes in CL-a(t) Рис. 22 - Динамические эффекты изменений CL-a(t) Slika 22 - Dinamicki efekat promena CL-a(t)
It is noticeable that a sudden increase in the angle of attack leads to a short-term lift increase. Keeping the angle of attack leads to a uniform flow past the airfoil, which leads to a stall.
A sudden increase in the angle of attack followed by return to its initial value shows that the value of the lift coefficient reduces compared to the stationary change of the angle of attack.
The diagram CL-a depends on the mode of airfoil entry into a stall. The cases of separation characteristic for the trailing edge have a more rounded shape of the curve than separation cases that occur at the leading edge.
The reduced frequency of changes of the angle of attack is k=cw/V, where с is the chord line, ш is frequency of changes of the angle of attack, and V represents velocity.
This equation is valid for non-compressible flow. For compressible flow, it is written as\k=^/V0(1-M)1/2.
1,6
Conclusion
The selection of an aerodynamic shape of an object in an air stream is a complex task. It is in relation to the conditions and purpose of flow past an objectas the starting factors. One of the limiting factors in exploitation is turbulence.
The analysis of experimental aerodynamic characteristics of different types of airfoils at different flow conditions categorized according to the Reynolds number has led us to a physical interpretation of the formation of laminar and turbulent flow separation.
We can conclude that the behavior of the airfoil at critical angles of attack:
- directly depends on the type of the airfoil and the local characteristics for the defined conditions of the flow;
- depends on the dynamics of the angle of attack increment at its critical values.
Based on the systematically presented results, it is possible to optimize the airfoil selection and reduce the number of iterations in the process, thus reducing the time and cost of design.
References
Cone, D.D. 1985. Maximum Lift.NASA TN D-657.
Critzos, A. 1985.Maximum Lift NACA 0012, 0015, a=0o-180o,.NACA TN 3361.
Doenhoff, A.E. Foil sections tested in two dimensions by NACA, Low-Turbulence Pressure Tunnel, T Rpt 1283. Retrieved from
www.ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19930090976.pdf
Gretchikhin, L.I. 2014. Force of the air pressure on a moving plate. VojnotehniCki glasnik/Military Technical Courier, 62(4), pp. 38-50.doi:10.5937/vojtehg62-5369
Hoerner, S.F., & Borst, H.V. 1985. Fluid-dynamic lift: Practical information on aerodynamic and hydrodynamic lift.Vancouver: Brick Town : Hoerner.
Kostic, C. 2010.Ispitivanje karakteristika kovita aviona u letu.Beograd: Univerzitet u Beogradu, Masinski fakultet.
McCullough,G.B.1955. Investigation of boundary layer and stalling.NACA TN 1683.
Pinkerton, I. 1996. Maximum lift 4412 Distribution. NACA T Rpt613.
von Karman,T. 2001. Lifting-line theory for a wing in non-uniform flow.Analysis of Maximum Lift,Robert A. Millikan Library, California Institute of Technology, Pasadena, California.
АЭРОДИНАМИЧЕСКИЕ ХАРАКТЕРИСТИКИ КРЫЛА ПРИ КРИТИЧЕСКОМ УГЛЕ АТАКИ
Чедомир Л. Костича, Бошко П. Рашуо6
а Генеральный штаб Вооруженных сил Республики Сербия, РВиПВО, 204 вб,
Батайница, Республика Сербия, б Машиностроительный факультет Белградского университета, г. Белград, Республика Сербия
ОБЛАСТЬ: машиностроение ВИД СТАТЬИ: обзорная статья ЯЗЫК СТАТЬИ: английский
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Резюме:
При разработке летательных аппаратов неследует пренебрегать фактором возникновения чрезвычайных ситуаций, а также условиями близкими критическим, так как при критическом угле атаки, обычный полет может перейти в аварийный режим.
В данной работе приведены важные факторы, влияющие на состояние крыла в течение полета при критическом угле атаки. На основании доступных испытательных результатов и расчетов, произведен анализ поведения крыла в зависимости от условий обтекания (в соответствии с классификацией чисел Маха и Рейнольд-са), формы крыла, динамики угла атаки. В работе также описано обтекание крыла при критическом увеличении угла атаки, и влияние неровностей на поверхности крыла в подобных условиях.
В данной работе также представлены физические параметры критического режима полета, в т.ч. снижения подъемной силы и сваливания. Подробно описан процесс обтекания крыла с отрывом потока и выполнена классификация типов профелей с учетом срыва струи и поведения крыла в критическом режиме полета.
Цель работы заключается в выявлении наиболее значимых характеристик обтекания крыла при критическом угле атаки и анализе, проведенного на основах современной аэродинамики, в результате которых предлагается практическая рекомендация для выбора аэродинамического профиля. Результаты данной работы могут представлять интерес как для летчиков и инженеров, так и для образовательных и исследовательских учреждений.
Ключевые слова: аэродинамика, форма и вид крыла, динамика флюидов, производительность крыла, критический угол атаки.
АЕтОР!ЫАМ!СКЕ КАРАКТЕгавПКЕ АЕтОРтОРНА рт кттюмм ыарарММ 1Ю1_ОУ!МА
Свбот1г Ц. КоБйс3, Воёко Р. РаёиоЬ
а Оепега^аЬ Уо]Бке БгЬуе, ^¡РУО, 204. уЬг, Ва1а]пюа, РериЬНка ЭгЬуа, ь итуегейе! и Веодгас1и, Маё^к fakultet, Веодгас1, РериЬИка ЭгЬуа
ОВЬАЭТ: таё^^о
УРЭТА С1_А1\1КА: pгegledni с1апак
^!К С1_А1\1КА: engleski
SaZetak:
РпИкот копэ&икСр vazduhoplova пе эте ее хапетагШ ропайап}е ееИсе и екэ^етпщ odnosno ЬНэко ек^гетпт uslovima 1е^, kakav ¡е е рп кгШспт napadnim ид^^, gde эе normalan е moZe ^ко ргеоЬтШ и р^исеп е.
U ovom radu naglaseni su bitni cinioci koji uticu na ponasanje aeroprofila u letu pri kriticnim napadnim uglovima. Na osnovu dostupnih eksperimentalnih rezultata i proracuna izvrsena je analiza ponasanja aeroprofila u zavisnosti od uslova strujanja (kategorisanih prema Mahovom i Rejnoldsovom broju), oblika aeroprofila, dinamike prirastaja napadnog ugla, opstrujavanja aeroprofila sa povecanjem napadnog ugla nakon dostizanja kriticne vrednosti, kao i uticaja povrsinskih neravnina aeroprofila pri kriticnim napadnim uglovima. Prikazano je i fizicko tumacenje pada uzgona i pojave prevucenog leta. Detaljno je opisan nastanak otcepljenja strujanja oko aeroprofila i izvrsena kategorizacija aeroprofila prema tipu otcepljenja i njegovom ponasanju pri kriticnim napadnim uglovima opstrujavanja.
Cilj ovog rada jeste da na savremenim osnovama aerodinamike prikaze i obrazlozi problematiku i najznacajnije osobenosti opstrujavanja aeroprofila pri kriticnim napadnim uglovima, da da prakticne preporuke prilikom izbora aeroprofila, a moze biti koristan i pilotima, inzenjerima i istrazivacima.
Kljucne reci: aerodinamika, performanse aeroprofila, dinamika fluida, oblik aeroprofila, kriticni napadni ugao.
Paper received on / Дата получения работы / Datum prijema clanka: 09. 04. 2015. Manuscript corrections submitted on / Дата получения исправленной версии работы/ Datum dostavljanja ispravki rukopisa: 24. 04. 2016.
Paper accepted for publishing on / Дата окончательного согласования работы / Datum konacnog prihvatanja clanka za objavljivanje: 26. 04. 2016.
© 2016 The Authors. Published by Vojnotehnicki glasnik / Military Technical Courier (www.vtg.mod.gov.rs, втг.мо.упр.срб). This article is an open access article distributed under the terms and conditions of the Creative Commons Attribution license (http://creativecommons.org/licenses/by/3.0/rs/).
© 2016 Авторы. Опубликовано в "Военно-технический вестник / Vojnotehnicki glasnik / Military Technical Courier" (www.vtg.mod.gov.rs, втг.мо.упр.срб). Данная статья в открытом доступе и распространяется в соответствии с лицензией "CreativeCommons" (http://creativecommons.org/licenses/by/3.0/rs/).
© 2016 Autori. Objavio Vojnotehnicki glasnik / Military Technical Courier (www.vtg.mod.gov.rs, втг.мо.упр.срб). Ovo je clanak otvorenog pristupa i distribuira se u skladu sa Creative Commons licencom (http://creativecommons.org/licenses/by/3.0/rs/).
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