DLR STERN Project

The STERN (STudentische ExperimentalRaketeN) program of the DLR (German Aerospace Center) is a program which gives student groups from German universities the possibility to participate in a real space project. Student groups have the mission to develop an own rocket within three years, which shall reach at least 3km altitude and the speed of sound. All rockets are launched from the Esrange Launch Site near Kiruna, Sweden.

STERN finances the student projects with funds from the BMWi (Federal Ministry of Economics and Energy). During the STERN program, several German student groups participated with solid, liquid and hybrid rockets which were successfully launched from ESRANGE. During the program, each team is controlled and supported by reviewers from industry and DLR during the development of their rocket and undergoes the space-typical review process. The main goals of the STERN program are to train students on real space projects, gather teamwork experience (which is inevitable for space projects) and give the students the possibility to develop new technologies and inventions.

	Compass	N2ORTH
Nominal Thrust	800 N	10,000 N
Length	2.45 m	7.79 m
Diameter	122 mm	259 mm
Tank Volume	5 L	160 L
Dry Mass	10 kg	TBD

Scroll down to learn more about the different subsystems of N₂ORTH and Compass. Every subsystem has its own dedicated team, which develops, builds and tests the corresponding components. For the most part, the subsystems share the same working principle between N₂ORTH and Compass.

Both rockets, N₂ORTH and Compass, feature a lightweight and efficient hybrid rocket engine with a CFRP combustion chamber hull. The propellants used are nitrous oxide (N₂O) as oxidizer and a custom developed HTPB-based fuel.  The propulsion system operates in blow down mode, meaning there is no active propellant feeding system. Instead, the self-pressurization capability of nitrous oxide is used and enhanced with additional helium, which is stored in the same tank.  This principle is called helium supercharging and increases the lift off thrust and ensures a safe pressure difference between tank and combustion chamber.

	HyFIVE Flight Engine	HyLIGHT Flight Engine
Nominal Thrust	800 N	10,000 N
CC Pressure	25 bar	28 bar
Operation Time	10 s	25 s
Specific Impulse	220 s*	224 s*
Fuel Mass	900 g**	20.5 kg** + 4 kg Film Cooling
Dry Mass	950 g	14.5 kg

*: Measurement on test bench; **: Includes additional fuel (safety factor);

In-house Developed Fuel

The in-house developed HTPB based fuel is completely non-toxic during all production steps and castable at room temperature. The grains feature a single-port geometry. Both engines have demonstrated highly efficient combustion (>95% c* efficiency).

Lightweight Engine Design

For our project, we decided to use a filament winded combustion chamber design. All casings are produced within our facilities with our new filament winding machine.

Testing our Engines

All tests are conducted at the DLR Institute of Space Propulsion Lampoldshausen at Test Bench M11.5. Most hot fires are conducted with constant tank pressurization using nitrogen. Additionally, blow down operation is tested together with other rocket systems (tank, valves, GSE) in a separate test before the launch.

Since September 2020, HyEnD has conducted more than 40 tests with the smaller HyFIVE rocket engine. Tests included the evaluation of the fuel regression rate and combustion behaviour as well as the test of different ablative insulation materials, injectors and mixing elements in the post combustion chamber.

First tests with a modular HyLIGHT engine with an aluminum casing were conducted in September 2021. Smooth and efficient combustion was achieved by using a multi-element swirl injector in combination with a mixing element. Different fuel grain geometries and film cooling ablator materials were compared. First tests with the CFRP casing were conducted in early 2022. A total impulse of more than 267 kNs was tested in a single hot fire.

After the successful flight of Compass, the structure and aerodynamics team is working on prototyping, manufacturing and simulating the N2ORTH rocket. This includes the structural Type V (no load-bearing liner) oxidizer tank, which features a filament wound CFRP construction and is manufactured in-house.

The nose cone is constructed from GFRP to satisfy radio transmissivity requirements and modeled after the LV-HAACK Series to minimize drag. As a first for HyEnD, the airframe is also filament wound in-house in order to maximize strength while allowing for the usage of specialized high-temperature epoxy resins.

To withstand shock heating while flying at Mach>3, all wetted parts are laminated with in-house developed high-temperature epoxy systems and protected by additional insulation material

The trajectory and thermal simulations are used to verify range safety and thermal loads by using Dispersion, FEM and CFD analyses.

The fluid system is mainly tasked with the development of the main vale, providing a fueling interface to the GSE and ensuring the safe handling of the oxidizer inside the rocket. Therefore, the following components are necessary:

• Relief Valve
• Fueling Coupling
• Check Valve
• Temperature Sensor
• Pressure Sensor
• Realease Valve
• Main Valve

The focus lies on the development of a lightweight, compact and reliable system. Because of the former two requirements, in-house developments for the main valve and the release valve are chosen over commercially available products.

Hence, the development of the main valve is a substantial task of the fluid system. A pyrotechnically actuated slider design is chosen, which has been examined in a number of test campaigns. This ranges from the first subscale prototype to the prototype of the N2ORTH main valve. Additionally, the concept was flight proven in the Compass rocket, where it performed flawlessly.

For the release valve, a lightweight and reliable concept is developed. Again, the concept was examined during test campaigns with different prototypes.

The system as a whole was successfully tested during the Compass launch and the Compass static hot fire test. These subscale tests are an important milestone for the development of the N2ORTH fluid system.

The N2ORTH recovery system is designed to perform the required safe landing of the entire rocket.

According to our ambitious altitude goal, a supersonic drogue parachute becomes necessary. A hemisflo ribbon design was chosen to meet these requirements. The drogue parachute is ejected at apogee to slow down the rocket to a steady descent rate.
At an altitude of 3000 m the second stage of the recovery system is triggered to further decelerate the rocket allowing the aspired save landing. All opening occurring shock loads are neutralized by polyamide shock absorbing harness components.

The entire recovery system is stored in a department in the upper part of the rocket.

The entire recovery bay is located in the upper part of the rocket. The following picture shows a schematic of the structure and the lines.

The avionics need to fulfil multiple tasks, mainly activating the pyro lines firing the recovery system as well as collecting as much data as possible to allow further analysis after the flight.

It consists of many different modules, each one dedicated to a specific task e.g. gathering trajectory data. The systems can communicate between each other to allow the different modules to send their measurement data to a central microcontroller. The collected data will then be transmitted to the ground station.

In order to obtain pressure data of the engine and tank with a high sample rate, a custom module is designed to read and store the measurements with a sample rate of up to 10 kHz. This enables a detailed frequency analysis after the flight, making it easier to compare the behavior of the rocket in flight with the previous tests and simulations.

Another part of the avionics is dedicated to the powertrain. A stable and redundant power supply is extremely important. A loss of electricity will render the whole mission unsuccessful. The power train is customly designed to prevent a variety of common problems, like overvoltage and overcurrent as well as being able to switch between two sets of batteries and voltage regulators.

The parachute deployment is done via two parallel Telemegas. They will detect the apogee and trigger the recovery mechanism. Both will send additional data regarding the trajectory.

The ground support team is responsible for providing a safe launch platform on the ground. The open-end design provides a platform that can be adjusted and repurposed for future projects.

The GSE control system has three main tasks:

1.            Controlling the fueling process through valves and sensors

2.            Communicating with the rocket on the ground via a CAN-Bus

3.            Ignition of pyrotechnic devices during launch

To accomplish these tasks, the GSE control system is divided into different custom PCBs controlled by a Raspberry Pi. This allows for the use of higher quality shielded connectors while still maintaining a compact design. Furthermore, the pyrotechnic section is galvanically separated from the rest of the system to prevent any risk of damage to the delicate electronics with high currents.

The Raspberry Pi is running C Code and communicates with the control computer over Ethernet. It can run completely independently which allows for automated actions and fail safe in case of communication issues.

The deployment of the oxidizer will be realized by our oxidizer supply system. It is crucial to deliver the right amount to ensure that the engine will operate within the defined parameters. Therefore, the nitrous oxide used as oxidizer will be stored in an intermediate gas tank. The weight of the tank will be measured. Combined with the level indication inside the rocket tank the amount of oxidizer shall be determined.

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