Problem with the thermal barrier coatings is not where that ZrO2 can not withstand temperatures above 1600°C, but at the that the substrate under the layer is destroyed from 900°C. If you have at the layer surface 1600°C, a temperature gradient must be implemented from about 700°C over the layer. That is only possible if the layer is several millimeters thick. So a thick ceramic layer will fall from the substrate in any case by the thermal shock, no matter from which ceramics it is made.

In other words: 1600°C on the ceramic layer surface can be realized only when the one-time applications (one rocket-start). For the multiple applications is 1600°C  too high, and the composition of the ceramic top layer is here not important.

I tend to agree with Vadim Verlotski what the specs for 1600C are outside of state-of-art (at least for steel engines). So some thinking out of box is necessary.

How about multilayer? If the rocket engine body is made of steel, how about cladding it with porous sintered titanium with glazed, oxidized surface? Titanium (unlike zirconium) is nearly compatible with steel by CTE, and pores will both reduce thermal conductivity and thermal stress. At said 700C temperature difference and 15 W/mK conductivity, 50mm thick cladding will do the trick to enable radiation-cooled engine.

Also, significant cooling improvement can be done with expander bleed cycle - either with surface film cooling or if cooling fluid flowing through the pores in protective coating.

Also, JAXA have mastered the nozzles made of Silicon Nitride engines capable of 1500C wall temperature. (though the Akatsuki emergency in 2010 clearly showed the need to manage start-up transients more precisely) 

https://www.mhi.co.jp/technology/review/pdf/e454/e454032.pdf

Thanks Vadim, your answer was very clear.

Thanks Oleg, i like the way you think, focused in possible solutions.

I am considering use of exotic metals alloys Re, Ir, Pt, Nb, Mo,.... For this purpose will be a more reliable based design. 

Thanks again!

Rhenium engines are the old and reliable solution. Cost-performance is awful though. Molybdenum was also used as relatively cheap alternative. Also, Mo-based coatings are widespread.

It is the cost, not the performance, which driven transition to Nb and steel.

Pay attention to recent development of super-alloys like

http://www.tms.org/superalloys/10.7449/2008/Superalloys_2008_131_138.pdf

using in rocket engines? Outside heat protection during re-entry or inside in the rocket combustion chamber? It is not clear to me

Example of rocket engine using heavy metals (rhenium and others)

http://arc.aiaa.org/doi/abs/10.2514/6.1998-3356

It is for the combustion chamber Klaus.

Oleg, thanks again for your contributions. You have expertise in this area. Best thrusters are made actually by Rhenium and Iridium. But the price is very high and only can be applied for small thrusters for this reason.

There was a study on heat protection of a lift-off rocket propulsion chamber consisting of LH cooled copper. Isolation was a thin YSZ layer to protect the copper substrate against the combustion heat of about 2500°C. This layer should be kept below a thickness that limited the allowable surface temperature of the YSZ layer to 1400°C at maximum. This thin inner coating resulted in a higher efficiency alias less LH + oxygen consuption as has been intended.

Re-entry of rockets assumingly need a very different approach regarding heat protection.