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Non-eroding Nozzle Materials for High Temperature Combustion Gases


OUSD (R&E) CRITICAL TECHNOLOGY AREA(S): Hypersonics; Advanced Materials


The technology within this topic is restricted under the International Traffic in Arms Regulation (ITAR), 22 CFR Parts 120-130, which controls the export and import of defense-related material and services, including export of sensitive technical data, or the Export Administration Regulation (EAR), 15 CFR Parts 730-774, which controls dual use items. Offerors must disclose any proposed use of foreign nationals (FNs), their country(ies) of origin, the type of visa or work permit possessed, and the statement of work (SOW) tasks intended for accomplishment by the FN(s) in accordance with the Announcement. Offerors are advised foreign nationals proposed to perform on this topic may be restricted due to the technical data under US Export Control Laws.


OBJECTIVE: Develop ablation resistant, non-eroding, rocket nozzle materials with high temperature strength and compatibility with highly oxidizing propellant combustion gases.


DESCRIPTION: Maximizing missile range requires higher propellant combustion temperatures without nozzle erosion. Controllable solid propellant rockets use a variety of propellants based on their application, and many propellants used for controllable solids produce highly oxidizing combustion gases. While some propellants are seen as reducing, this topic is specifically asking for solutions for highly oxidized propellants. Besides high temperature strength, successful nozzle materials must also be capable of surviving the thermal shock experienced during ignition.


PHASE I: Develop nozzle material solutions with predicted high temperature strength and resistance to high partial pressures of oxidizing species such as oxygen, carbon dioxide, and water. If no testing is performed in Phase I, thermo-structural and chemical reaction modeling should demonstrate ablation and erosion resistance, high temperature strength, and thermal shock resistance for use in rocket motors with combustion temperatures up to 2800°C. Manufacturability of the material solution must be demonstrated.


PHASE II: Demonstrate the survivability of the material solutions developed in Phase I with loads and temperatures representative of those experienced by solid propellant rocket nozzles. High temperature mechanical and thermal material properties of the material solution should be characterized by the end of this effort. Testing must demonstrate thermal shock resistance under temperature rises experienced by rocket nozzles. Identify additional applications for the proposed technology beyond MDA applications.


PHASE III DUAL USE APPLICATIONS: Produce nozzle components that meet the requirements of a propulsion system supplier and demonstrate performance of the nozzle components through static testing. Demonstrate the quality, reproducibility, and production requirements for a developing, prime contractor system.



  1. Non-eroding nozzle throat material for rocket motors with AP-based propellant
  2. Chemical Erosion of Refractory-Metal Nozzle Inserts in Solid-Propellant Rocket Motors
  3. Status of army pintle technology for controllable thrust propulsion


KEYWORDS: Hypersonics; Materials; High Temperature; Propulsion; Nozzle

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