In this section, the applied heat transfer models are described. The sensitivity of the methods to changes in geometry is tested and the impact of surface heating on drag is assessed.
3.2. Sensitivities
The aforementioned local discretization of the heat transfer calculation depends on
.
Section 2.2 focuses on correlations of the total and component wise
. However, to calculate
more knowledge of the geometry is required. For example, two fuselages with the same
have different
distributions if their slenderness ratios (
) differ. To account for these effects, the geometric model of the components is refined. For cylindrical components (fuselage, nacelle) the sensitivity of
is studied:
With length (
l) and diameter (
d). Wing and tail components are modelled as single section trapezoids with no leading edge sweep. Their geometries, specifically the span-wise chord distribution can be fully defined with the help of their
, aspect ratio (
) and taper ratio (
) [
20]. The following sensitivity studies are conducted around TO conditions. It is one of the most critical conditions for TMS design, because of the low air flow velocities, high ambient temperatures (
) and large cooling demand (
) due to maximum propulsive power. Unless otherwise specified, the values in
Table 4 are assumed for the wing sensitivity studies. The values are not specific to any aircraft but generally lie inside the range of the given data.
is the critical Reynolds number and
the average surface temperature. In a real cooling application, the surface temperature would most likely not be uniform but have a gradient in the direction of a hot side flow underneath the surface. However, in this first approximation an average
is assumed for simplification.
3.2.1. Transition Location
Flat plate heat transfer correlations distinguish between laminar and turbulent flow. They rely on the knowledge of a critical location (
) where transition occurs. Usually
is defined by
which according to Reference [
24] is between
to
depending on free stream turbulence and surface roughness. Detailed transition modelling is a complex research area and beyond the scope of this work. However, a sensitivity study with varying
and
ranging from slow taxiing
to representative TO
is conducted. The results of the theoretically available cooling power (
) as well as the relative
compared to the
at the lowest
for each
(
) are displayed in
Figure 3.
An increased
results in increased
because of the increased effects of forced convection. Shifting
to higher values, results in a decrease in
. Turbulent flows favour heat transfer more than the structured flow in laminar regions because of the increased particle mixing within the boundary layer. With increased
the portion of
with laminar flow increases. For very low
increasing
beyond
results in laminar flow on the entire surface. A further increase in
has no additional effect. The transition point has a large influence on
. For 3D wings, transition is more complex than defining an
and assuming instantaneous transition. This study does not accurately account for real transition effects. The results in
Section 4 assume fully turbulent flow areas downstream the transition location. Therefore, the results of this study cannot directly be used for concepts with enhanced laminarity such as natural laminar flow (NLF) wings. Covering these advanced aerodynamic concepts is part of future work.
3.2.2. Wing Aspect Ratio
is varied from 6 to 18—a range that includes all aircraft used for the correlations in
Section 2 and also leaves margin for possible future aircraft with increased
. The results of
as well as the relative
compared to the
at the lowest
for each
(
) are displayed in
Figure 4. For better comprehension of the trends in
Figure 4, the effect of increasing
on the local
distribution for the lowest and largest
are illustrated with heat maps in
Figure 5.
increases with , because the forced convection increases, due to increasing . Depending on , increases or decreases with increasing . More specifically for the lowest of , decreases with increasing . For all other used in the study increases with . Two counteracting effects are the reason:
In general,
decreases along
x because of the increasing thickness of the thermal boundary layer (
). Therefore, higher
favours heat transfer because for the same area, the average chord length is lower (cf.
Figure 5 bottom graphs).
The front section of the wing is laminar, which results in small
. A higher
increases the span and, thus, the laminar portion of the plate’s total area (cf.
Figure 5 top two graphs). The
depends on
. For low
the transition occurs further downstream, which means that this second effect contributes more.
Tripling the aspect ratio results in depending on . The sensitivity is too weak for the expected precision of this study that aims to determine the order of magnitude of the surface cooling power. Hence, it is not regarded in the following studies.
3.2.3. Wing Taper Ratio
is varied between and . The same range as in the previous sections is applied. Variations in do not exceed with slight advantages for the non-tapered wings (). The aforementioned effect of increasing flow length is positive for heat transfer of tapered wings near the wing tip but negative near the root, which leads to its equalization after integration over the entire span. As with the sensitivity, the effect is too small to be further considered in this work.
3.2.4. Fuselage Slenderness Ratio
For a fuselage with m, is varied from 5 to 15 within the same range as the previous sensitivity analysis. Regardless of , the change in from the lowest to the highest value is around −8%, again due to the increasing flow length with increasing . The effect is also within the expected precision of this study. For further investigations, is used, which is conservative as it is one of the highest values found in today’s aircraft for example, for the Airbus A340-600.
3.3. Drag
For any aircraft component, which contributes to the aircraft’s drag, local surface temperature can influence the aerodynamics of air passing the component surface at a certain velocity and with certain fluid characteristics. The two main occurring effects depending on the fluid’s initial state are:
As the skin friction coefficient (
) is significantly smaller in laminar than in turbulent flow, total skin friction drag (
) of a surface can be decreased by moving the transition location downstream that is, by increasing the laminar length. During the last centuries, laminar flow control approaches have been studied intensively as a means to decrease drag. As such, surface temperature alteration can be employed to decrease the growth rate of unstable disturbances in the fluid and thus, to repress transition from laminar to turbulent flow [
27]. The application of this method was shown in experiments by for example, References [
27,
28]. Two different approaches apply [
29]:
In the two-dimensional case, Tollmien-Schlichting instabilities, which dominate the laminar boundary layer, are mitigated by cooling of the near wall boundary layer. In accordance with theory, flat plate experiments showed that the cooling of a surface leads to an increase of
and a downstream movement of
[
28]. The effect is reversed when the surface is heated: the destabilizing effect of the temperature increase in the boundary layer dominates and
moves upstream [
27].
However, the stabilizing effect of cooling can also be utilized when a portion of the surface is heated at strategic locations. For a two-dimensional case, it was shown that heating a portion of a surface where stable laminar flow is present (preferably the leading edge) followed by a cool that is, unheated, “relaxation” surface downstream can lead to a preferable downstream movement of
. The heated wall has to be situated in the region where Tollmien-Schlichting waves start to develop in the laminar boundary layer. The temperature of the near wall boundary layer is increased and when the fluid reaches the cooler wall further downstream, the temperature of the boundary layer is higher than the wall temperature. The boundary layer is cooled down and the growth rate of the unstable disturbances is decreased. The transition point moves downstream. If the surface is heated in an unstable flow region, the effect is reversed [
27,
29,
30].
In three-dimensional airflows, however, cross-flow instabilities determine the boundary layer. Dovgal et al. showed that in this case, a temperature increase of the near wall boundary layer fosters cross-flow instabilities no matter if the whole surface or only a part of the surface is heated. The transition location moves upstream resulting in an increased
[
27,
30]. Thus, for any three-dimensional aircraft component, localized and global surface heating in the laminar flow region facilitates laminar to turbulent transition and increases
.
In contrast, when the boundary layer is fully turbulent, different mechanisms govern the flow: Heating of the near wall boundary layer reduces the turbulent
. Kramer et al. conducted wind tunnel experiments and flight tests in 1999. They found that an increase of the near wall boundary layer temperature leads to a decrease of
, which in turn leads to a reduced local skin friction force [
31]. For a body similar to a fuselage, they showed that the heating of the fore body leads to a higher drag reduction than the heating of the aft body, whereas the heating of the whole body has the highest drag reduction potential. The findings are supported by a numerical evaluation of the effect of heating on the turbulent boundary layer flow over slender and bluff fuselage-like bodies conducted by Lin and Ash in 1986 [
32]. The following theoretical deviation of
as a function of wall heating for a smooth flat plate is based on the deviation proposed by Reference [
31].
The length Reynolds number is defined as:
For
, the turbulent
for a flat plate of length
x can be expressed with
,
by Reference [
33]:
Total skin friction drag of a flat plate with the length
x and total area
A for a turbulent boundary layer is defined as [
33]:
Assuming a constant heated surface temperature (
) along
x, the temperature ratio of unheated air (
) and
is defined as:
Applying the ideal gas law leads to
and the dynamic viscosity of air can be simplified to
. Thus:
When the wall is heated (
),
decreases with increased temperature. In consequence,
increases. However, the change in
has a larger effect on
than the change in
:
and therefore if
. The higher the wall temperature compared to the ambient temperature, the higher the drag decreasing potential. All simplified relations are depicted in
Figure 6.
Wall heating not only has an impact on skin friction drag but also effects the pressure drag. The turbulent boundary layer velocity profile thickens, because [
34]:
For a flat plate, the pressure gradient is zero at all locations. For a slender body (fuselage) or lifting surface (wing, tail planes), however, the pressure gradient varies in stream wise direction. Therefore, for a three-dimensional curved body, the heating of the wall has an effect on the (not-separated) pressure drag as shown by Lin and Ash. The heating of the wall increases the turbulent displacement thickness (
), which in turn leads to a slight increase in pressure drag [
32]. In addition, the boundary layer shape factor is increased. Thus, the adverse pressure gradient is increased, causing an earlier flow separation [
32,
35]. The effect of wall heating on pressure drag is small compared to the effect on skin friction drag [
32].
In summary, in regions in which the boundary layer is laminar, an increased temperature leads to an earlier transition and, thus, to an increase in total . To make use of the beneficial effect of wall heating on the turbulent drag force, the surface of aircraft components should preferably be heated only in regions in which a fully turbulent boundary layer is present. This means that for example, the fuselage nose (cockpit area) or wing leading edge (slats etc.) should not be used for heat disposal. For aircraft concepts that unite different technologies, which emit excessive heat and aim at an increased laminar flow control, detailed studies have to be conducted, compromising excessive heat disposal and drag reduction approaches.