Aircraft and Helicopter Icing Experimentation/Simulations
Experimentation: Icing simulation in wind tunnel
Icing can be reproduced in wind tunnel : a spray bar system injects water droplets, equivalent in diameter and concentration as those found in natural clouds. Water is injected in air stream at negative temperature. Airs tream can be cooled (Centre d'Essais des Propulseur de Saclay-FRANCE) or can be extracted from outside in winter (wind tunnel S1MA, Modane - FRANCE).
Tests are used both to study the ice-deposit growing mechanism and to validate the codes. Two fields are currently under study :
Freezing drizzle or freezing rain conditions, caracterized by droplets larger than 0.2 mm. Those conditions led to a air-plane crash in 1992 near Chicago, and have been subjected to research relating to their formation process.
Icing deposits called "lobster tail deposits", produced on swept wings.
Moulding of a "lobster tail deposit"
produced on a swept wing, in the S1MA wind tunnel
Numerical simulation of ice deposits and protection systems.
To be allowed to fly in icing conditions, an aircraft has to :
possess a protection system minimising icing effects and permitting it to keep on flying
be able to maintain suffisant flying conditions, in order to reach the nearest airfield, in case of protection systems failure
The proof of these capabilities has to be performed by aircraft manufacturers. Several methods can be considered : flight tests in icing conditions, icing tunnel test or numerical simulation.
Two-dimensional (2D) and three-dimensional (3D) numerical codes have been developed in order to simulate ice accretion. Actually, they are based on four different codes :
the first one determines the flow field around the airfoil
lthe second one computes the water droplet trajectories , giving the local catch efficiency intensity on the airfoil
the third one gives the heat transfert coefficient using a boundary layer calculation which takes into account the roughness of the profile due to ice deposit
the fourth one computes the thermodynamic balance at the wall. This balance enable to compute locally the ice accretion rate
Figure 2 shows an example of ice deposit growth computed with ONERA's 2D code. Results has to be compared with an ice deposit obtained in icing wind tunnel (Figure 3).
Fig. 2 : Ice accretion simulation on a profile
(section of an airplane wing)
ONERA's 2D code has been widely distributed to manufacturers for airplanes certification. It enables to determine the shape of ice deposit occuring in utmost icing conditions rarely met during flight tests. Then, some flight tests are performed using repliqua of the computed ice deposits pasted on the wings. Thus airplane behaviour in case of protection systems failure could be analyzed.
ONERA code has been compared to similar codes developed at NASA and DRA (Defense Research Agency - GB) and gives similar results in comparison with experiments.
After the computation of ice deposit on a profile, de-icing or electrothermal anti-icing system can be simulated with MAD 2D code. Thermal transfer in the wall and in the ice is computed till ice melting. This code has been compared to similare ones (Fig. 4) and has been validated in icing tunnel with infrared thermography measurements
Fig. 4 : Surface temperature profile during an electrothermal de-icing sequence
Results comparison of NASA, DRA and ONERA codes
Prospects
Numerical codes developed at ONERA can be used to define optimal configuration test in order to reduce flight test cost. Because of the increasing sensitivity to ice of new planes, due to thinner airfoils and less powerful engines, thus requiring a protection system optimisation, the need of such computer models will increase in future.
Current and future studies are attempted to improve existing models. ONERA is currently developing a 3D ice accretion code, because two-dimensional approximations are not sufficiently realistic in some cases. Existing codes will also be adapted to different parts of a plane (air intake, guide vanes...).
