Scalability of aerodynamics from demonstrator to full flight scale of a vertical landing reusable launcher

Marwege, Ansgar; Deutsches Zentrum für Luft- und Raumfahrt e. V. (DLR); Goldyn, Pawel; Klevanski, Josef; Gülhan, Ali

doi:10.5281/zenodo.17200016

Scalability of aerodynamics from demonstrator to full flight scale of a vertical landing reusable launcher

Marwege, Ansgar; Deutsches Zentrum für Luft- und Raumfahrt e. V. (DLR); Goldyn, Pawel; Klevanski, Josef; Gülhan, Ali

Publisher: Zenodo
DOI: 10.5281/zenodo.17200016
Repository: zenodo.org
Language: en

Abstract

In the last decade, the vertical landing of launcher first stages has gained large interest due to the successes of SpaceX with its Falcon 9. In this context, the SALTO (reusable Strategic Space Launcher Technologies & Operations) project, funded by the European Union in the frame of the Horizon Europe programme, is supporting the ESA Themis programme in which a demonstrator for a vertically landing launcher first stage is built. A first version for this demonstrator is propelled by one Prometheus engine with about 120 t thrust and performs simple trajectories. This version is called T1H, where the H stands for the “Hop Test” to be performed. A later version, called the Themis 3 or T3, shall have 3 engines and perform more complex flight trajectories. In SALTO the Hop Test of T1H is performed and technologies for the T3 vehicle are matured. DLR has various tasks in the scope of the SALTO project. One of them is the investigation of the scalability of the technologies developed for the demonstrator to full-scale flight vehicles. This paper will focus on the scalability of aerodynamics between vehicle sizes, e.g. if the grid fins are designed for the trajectory of a specific demonstrator – are these grid fins then representative for a full-scale flight vehicle; are the flying qualities of the complete launcher configurations comparable; etc. For this purpose, two preliminary full-scale launcher configurations have been designed based on the same technologies as used in the SALTO project (mainly the same Prometheus engines and the same assumptions for these). In this context, a DLR in-house toolset named AIOLOS has been used for the preliminary launcher design. The toolset had been developed partly in scope of SALTO, with the aim of considering the launcher reusability already in the first launcher design loop; this has been achieved by expanding the existing launcher design mathematical model for a general reusable or expendable launch vehicle. Based on these full-scale launcher configurations, in this paper, the aerodynamic similarity parameters, the aerodynamic design, and the aerodynamic modelling approaches are compared between the full-scale launcher configurations and the T3 demonstrator.

Full text

SCALABILITY OF AERODYNAMICS FROM DEMONSTRATOR TO FULL FLIGHT SCALE OF A VERTICAL LANDING REUSABLE LAUNCHER FAR 2025 19.05.2025 Ansgar Marwege, Pawel Goldyn, Josef Klevanski, Ali Gülhan Institute of Aerodynamics and Flow Technology Supersonic and Hypersonic Technologies German Aerospace Center (DLR e.V.) SALTO •SALTO - reusable Strategic Space Launcher Technologies & Operations •funded by the European Union in the frame of the Horizon Europe programme •supports the ESA Themis programme •T1H hop tests •Technology maturation for T3 and future launcher configurations 2 SALTO 3 SALTO 4 Hop Test demonstrator with one Prometheus® Demonstrator to be flown in the frame of SALTO. T1H Reference configurations for future launchers with vertical take-off vertical landing. 5 Prometheus® and 9 Prometheus® engines. Future Launcher Configurations MECO Flip over LANDING BURN AERODYNAMIC PHASE REENTRY BURN 2nd stage to orbit MECO Flip over LANDING BURN AERODYNAMIC PHASE REENTRY BURN Reference configuration for future demonstrator version with three Prometheus® engines with more complex trajectory including all flight phases of vertical descent and landing with retro-propulsion T3 High TRL Medium TRL Low TRL T3 vehicle •3 Prometheus engines (approx. 120 t thrust) •Approx. 80 km altitude •Diameter 3.5 m •Length approx. 29 m 5 6 Configurations inspired by Ariane Group NESTS Study Configurations ArianeGroup - The future of European space transportation NESTS (New European Space Transportation Solutions) https://www.youtube.com/watch?v=CqWCc64Xbi0 •MEDIUM Launcher •First Stage: 5 Prometheus (adapted for sea level) •Second Stage: 1 Prometheus (adapted for vacuum) •HEAVY Launcher •First Stage: 9 Prometheus (adapted for sea level) •Second Stage: 1 Prometheus (adapted for vacuum) Methodology •Engine data: RPA •Geometry, masses, dimensions and the overall launcher design: AIOLOS •Trajectory: STRATOS •Aerodynamics: •Ascent: CAC (Calculation of Aerodynamic Coefficients) •Descent: AEDB from the ESA RETPRO project 7 AIOLOS: Goldyn, P., Marwege, A., Riehmer, J., Klevanski, J., and Gülhan, A., "Preliminary Design of Expendable and Reusable Mixed-Staged Launch Vehicles," Journal of Spacecraft and Rockets, Vol. 0, No. 0, 2025, pp. 1-24, 10.2514/1.A36174. Prometheus Engine (DLR rebuild) •Start with 100 t thrust and 100 bar pressure in combustion chamber [1] •Set thrust to 120 t [2] •Assumption: 120 bar achievable in combustion chamber (Vulcain 2 approx. 117.3 bar [3]) 8 [1] Bonhomme, C. et al. "Prometheus: European next generation liquid rocket engine," The 68th International Astronautical Congress (IAC), Adelaide, Australia, 2017. [2] Patureau de Mirand, A. et al. "Ariane Next, a vision for a reusable cost efficient European rocket," 8th European Conference for Aeronautics and Space Sciences, Madrid, Spain, 2019 https://www.eucass.eu/doi/EUCASS2019-0949.pdf. [3] "Website Vulcain 2." http://cs.astrium.eads.net:80/sp/launcherpropulsion/rocket-engines/vulcain-2-rocket-engine.html (accessed 05.12.2015). Engine Lower Stage Engine Upper Stage Engine Characteristics Adapted for sea level Adapted for vacuum – 100 as a good compromise Fuel LCH4/LOX LCH4/LOX Oxidizer Fuel Ratio (OFR) 3.5 3.5 Expansion Area Ratio 20 100 Chamber Pressure 120 bar 120 bar Exit Pressure 0.756 bar 0.094 bar Exit Diameter 1.2 m 2.68 m Isp Vacuum 355.04 s 384.83 s Isp Sea Level 323.19 s 225.59 s Thrust Vacuum 1259.287 kN 1312.668 kN Thrust Sea Level 1146.319 kN 800.142 kN !"# !$# !%# !&# !'# (## #)# (# $# &# *## *)# *(# *$# *&# +,-..I01230425 673,829:80;<=,0>,?9:04@5 MEDIUM •Launch site: Kourou, 5°10’08” N, 52°41’25” W •Orbit inclination same as launch site’s latitude: +5°10’08” •Design Payload: 16.24 t to LEO (200 km) •Payload to GTO: 0.92 t (estimated) •Gross Lift-off Mass (GLOM): 483.74 t •Thrust to Weight Ratio: 1.2 •Structural fraction dry •6.8 % – stage 1 •7.4 % – stage 2 9 Stage1 Stage2 Mass ratio 2.397 5.405 Stage mass share 0.670 0.296 Δvi2953 6368 Mass 324094 143409 Mass propellant ascent 281848 130114 Mass propellant descent 14092 0 Mass propellant total 295940 130114 Mass propellant reserve 2959 1301 Mass propellant dead 5919 2602 Mass payload 159649 16239 Mass gross real (GLOM) 483542 159552 Mass structure real wet 42246 13295 Mass structure real dry 22035 10597 Mass structure margin 2203 1060 Structural frac. real wet 0.130 0.093 Structural frac. real dry 0.068 0.074 Height 30.97 26.37 Outer diameter 4.5 4.5 Thrust (vacuum) 6108160 1312670 Thrust (sea level) 5731600 --- Thrust to Weight ratio 1.209 0.839 Reusability index 1.05 1.00 !"#$% &'()*+ ,-.+'/0'*1 !/23'4 !/23'5 ! "!!!!! #!!!!! $!!!!! %!!!!! &!!!!! '())6*+ !/-*7/*-' +2118 !/-*7/*-' +211 +2-3)98 :'2( ;-.;'<<29/ +2118 :'17'9/ ;-.;'<<29/ +2118 "17'9/ ;-.;'<<29/ +2118 ,2=<.2( +211 ! For definitions refer to: Goldyn, P., Marwege, A., Riehmer, J., Klevanski, J., and Gülhan, A., "Preliminary Design of Expendable and Reusable Mixed-Staged Launch Vehicles," Journal of Spacecraft and Rockets, Vol. 0, No. 0, 2025, pp. 1-24, 10.2514/1.A36174. Trajectory MEDIUM 17 𝑀𝐹𝑅 =! !,#$%"#&$& %&''&#& 𝑞(=1 2𝜌(𝑢( %=1 2𝑝(𝑀( %𝛾( Trajectory MEDIUM 18 Heat Loads: 120 kPa – R0L1: 388 kW/m2 100 kPa – R1L1: 309 kW/m2 100 kPa – R2L1: 374 kW/m2 100 kPa – R3L1: 374 kW/m2 Trajectory HEAVY 19 Fuel Consumption: 100 kPa –R2L3: 29.7 t 100 kPa –R3L3: 28.8 t 120 kPa –R2L3: 26.6 t 120 kPa –R3L3: 26.2 t Trajectory HEAVY 20 Trajectory HEAVY 21 Trajectory HEAVY 22 Heat Loads: 120 kPa – R3L3: 351 kW/m2 100 kPa – R3L3: 327 kW/m2 120 kPa – R2L3: 275 kW/m2 100 kPa – R2L3: 205 kW/m2 Summary and outlook •MEDIUM and HEAVY configurations based on Ariane NESTS Study •MEDIUM with 5 Prometheus engines, HEAVY with 9 Prometheus engines •Re-entry burn: T3 provides especially good similarity for MEDIUM •For MEDIUM Re-entry burn could be skipped •Landing burn: excellent similarity (Momentum Flux Ratio and Thrust Coefficient) 23 Summary Outlook •Computation of (low-fidelity) Aerodynamic Databases (AEDBs) for both configurations •Flight dynamics and mission analysis by DEIMOS •Analysis on structural integration of landing legs and grid fins by MT Aerospace Ansgar Marwege Aerodynamic Models of T3 Monday 4:40 pm – 6:00 pm: session 1.11 Gabriele De Zaiacomo Flight Dynamics of T3 Tuesday 3:40 pm – 4:00 pm: session 2.13 Jan Vos Aerodynamic Database of T3 Monday 3:20 pm – 3:40 pm: session 1.7 Tim Horchler Unsteady Pressure Loads on T3 Monday 11:40 am – 12:00 pm: session 1.3 Mariasole Laureti Aerothermal Loads on T3 Monday 3:20 pm – 3:40 pm: session 1.5 Thank you!