i wonder if the coeff of thermal expansion (CTE) of the composite is much greater than aluminium and may allow more aircraft growth and contraction
Coeff of thermal expansion is negative for carbon fibre. Coeff of thermal expansion of the CFRP depends on the layup, it might be negative in one direction and positive in the other one, it might be very close to zero in a balanced layup of fibres in a matrix.
However, a wound fibre fuselage barrel may increase its diameter (fibre direction) in the cold air at altitude, while increasing additionally due to the pressurisation loop stress, so overall diameter increase may be more than the one of an aluminum fuselage. But I seriously doubt that this is a challange for the wiring, it should be designed to withstand airframe deformation.
And, prior to getting to my final points regarding thermal insulative and conductivity properties of aluminum alloys and CFRP's as I promised that the outset of this overly long input, let me give some relevant metallic vs. CFRP cross plied properties for others to mull over. CF is widely touted, and oft over-touted, regarding its strength, but let us be very careful here. A decent CFRP will have a tensile composite strength, if a UD material at 60 % fiber volume, of around 280- 320 KSI,( note that is the sigma 1-1 equivalent to metallics).
A quasi-isotropic composite in the same CFRP will have a tensile value of around 90-120 KSI. Now, let us look a couple of weaknesses without boring you all with the hygroscopic nature of that nasty epoxy.
The shear strength of metallics is typically 60% of the tensile, so for a decent steel, for example, say in the 240 ksi range we will have a shear strength of around 145 KSI. Now look at composites and up pops that nasty epoxy again. Whereas Steel will have a SBS or ILSS strength of 144 KSI, a quasi-isotropic composite will have an ILSS of 8-10 KSI static on a good day and, if we allow for fatigue et al, we are down in the 4 ksi area. And finally another key nasty is the short transverse tensile strength ( equivalent to metallic sigma 3-3), which is entirely epoxy dependent and no CF failure is involved, there we find on a good day around 3-4 KSI with fatigue knocking that down to around 1-1.5 kSI wihich is close enough to zero in this composite engineer's mind.
This is the reason why a composite aircraft structures design should look totally different from a metallic one. Both material types have their superior strengths and their enormous weaknesses. The key to composite deign is to understand this and to design accordingly. For example in a pressurized fuselage skin the loading is almost entirely in-plane, so sigma 3-3 or ILSS is "close enough to zero" to perfectly meet the load carrying capability of a CFRP layup. On the other hand any item loaded in all three dimensions (e.g. a landing gear main fitting) will never be seriously proposed to be made from CFRP. As stated before, composites are more or less artificial wood, the structures design should be fairly the same (except that nails my not be a good idea...). Unfortunately in the large aeroplane world (maunfacturers, operators, authorities) there is little knowledge of wooden aeroplanes, hence there is an awful lot of misconception going on.
Having designed, build, tested to destruction, certified and repaired a CFRP glider wing, I often wonder what those "metal guys" are trying to build from CFRP...