Space Launch in Nigeria and Technology

Overview

For a long time, the capabilities of Nigerian small satellites have seemed to transition from technology demonstrators and academic interest to a more operational mission. As a tradition these classes of satellite have relied on parasitic launch capacity, which, of course, is currently not adequate when businesses, as well as government, depend on timely mission deployment and success. Launch methods for Nigerian small satellites have, therefore, been faced with barriers due to the lack technological know-how and skills.

Nigeria joined Spacefaring Nations on the 27th of September, 2003 with the launch of its first satellite “NigeriaSat-1” into the Low Earth Orbit (LEO). This Satellite carries Remote sensing payload; it also has a Ground Sampling Distance of 32m that seemed to be a fulfillment of one of Government thrust towards solving socio-economic problems, as well as its realization of sustainable development plans through the applications of space science and technology. Moreover NigeriaSat-1 project has included expertise acquisition in the area of Satellite Technology where 15 Nigerian Engineers/Scientists underwent an 18 months Know How Technology Surrey Satellite Technology Limited (SSTL) in order to be conversant with the technology. Alongside their SSTL counterparts, the trainees designed and built the NigeriaSat-1. There has been a full commission NigeriaSat-1 from the Ground Control Station in Nigeria and the spacecraft indicating that it is in perfect working condition. This spacecraft has been known and acknowledged for sending down excellent images of the earth. Nigerian Ground Station Engineers have been controlling NigeriaSat-1. The NigeriaSat-1 has been considered as one of the Spacecrafts that helps in Disaster Monitoring Constellation (DMC) which is made up of the following member states, United Kingdom, China, Turkey, Thailand and Algeria. The Disaster Monitoring Constellation provides real-time data with a global coverage

Problem

Even though small satellite launch in Nigerian, just like in other places, have been tried to be solved primarily by piggyback launch where satellites tend to use spare capacity on cluster launches or commissioned launches which usually share launcher capacity by pooling the requirement for many satellites. These methods, however, are not providing comfortable conditions for an in-time launch arrangement in most cases, (G. Webb, A. da Silva Curiel (2008).

Discussion

The proposed technological capacity for Nigeria Space Program includes design of a Micro Launch Vehicle (payload capacity: 70-75 kg to LEO. The key component to building a low-cost launcher is Low cost propulsion. Currently, WEPA-Technologies is working on developing an expendable, but much cheaper propulsion unit and 35 kN liquid propellant rocket engine (LPRE). Among the potential uses could be in booster applications for sounding rockets or Micro Satellite Launch Vehicles. Taking into account such widely known use of solid propellant booster stages in sounding rockets, by substituting with LPRE stages would bring out an improved variability of payload size as well as the maximum altitude. Moreover, use of liquid propulsion systems would assist in avoiding several safety and legal restrictions that are connected with handling of solid propellants.

Micro Satellite Launch Vehicle

Currently, the standard launch services for microsatellites tend to be carried out through secondary payload rides. In the process of the procedure, the microsatellite tends to be carried together with a much larger satellite. As a result, each satellite’s launch fees due for each vary significantly. Importantly, that significant launch parameters are usually tailored sufficiently in order to meet the needs of the most expensive payloads based on the launch schedule and orbit. Again the required lead time until launch date regularly tends to be in excess of a period of one year, and there are existing significant restrictions in regard to the secondary payload present. Therefore, “Responsiveness” based on the short lead time to launch is unavailable.

The approach used in regard to propulsion units has been based on proven technologies that have been successfully used in flight vehicles of the pioneer period in the USA and former USSR. There is a desire in upgrading by modern standard construction materials and production techniques, however; it is important to avoid modern ‘high-performance materials. Again, the preference should on use of standard, mass production parts and manufacturing of the key components are supposed to be done in-house.

Some of the key points of the launch vehicles include: Majorly designed to use three stages with the aim of avoiding high performance propulsion units as well as advanced vehicle construction issues as per the requirement during a two-stage approach use; Instead of scale up there is preference of numbering up or clustering of propulsion technology; Maximum likely use of identical parts in every stage; easy to handle fuel combinations (LOX, HTP, EtOH), environmental friendly. The purpose is to help in driving the costs during storage, filling as well as in general handling procedures during the entire stages of systems life cycle, (Caceres, M., 2013).Based on conceptional design for the present design study, payload range of 50 – 100 kg was considered since it is common for most Micro Satellite projects. Since the present design study was of a general nature, the focus was on an average value, i.e. 70 – 75 kg to be delivered to LEO. Micro Satellite being preferred as the nature of the payload, but there was room for choosing any other kind of equipment.

Definition of LEO is done to the point of reaching a final upper stage velocity of at least 7, 5 km/ s – since any detailed aerodynamic and trajectory optimization was not conducted, and individual stage velocities calculation has been performed by use of ideal rocket equation. For this inevitable losses to be compensated following the drag and gravity, whereas additional 2, 0 km / s were added (therefore minimum velocity to enter LEO increased up to 9, 5 km / s). The propulsion base case uses LOX and EtOH. While LOX tend to be a well proven oxidizer that is present in the rocket business since it start, there are some specifics of its own that should undergo consideration: isolation requirement (tanks, pumps, sensors) evaporative losses during proportional time, chill down requirements of engine as well as turbopump unit and reliability issues that concerns ignition of upper stages within vacuum.

These listed points tend to result to a significantly reduced degree of system complexity as well as increased operational reliability. Therefore, there is a possibility of multiple burn periods: such will majorly lead to the advantage of achieving more exact orbital insertion of payloads or to some level realize various orbital heights for delivering multiple payloads. The above noted HTP upper stage advantages are only attainable if storable propellants such as dinitrogen tetroxide /hydrazine derivatives (e.g. UDMH).are applied.

These propellants to be known to be highly toxic and therefore needs far-reaching safety protocols that should be followed. A lot of costs is significantly incurred. Such disadvantages can be put to a stop through the use of HTP based upper stage systems. Determining the dimensions and weight were done by first calculating a launcher using LOX / EtOH within every stage.

Current Development Activities

Currently, there is consideration of low-cost propulsion systems as being a key component in realizing low cost Micro Satellite Launch Vehicles, (Andrews, D., 1990).Achieving low-cost propulsion systems can be through: using low-cost materials and manufacturing technologies: – low-level operational parameter (temperatures, chamber pressures, RPM of TPU); focusing on expendable systems; and simplified design of rocket engines and turbo pump units

Design parameter

It is categorically meant to create a design minimizing manufacturing, engineering, and testing effort. Due to that, there are uses of low-level operational parameters.

Summary of key points of the TPU are as follows: –

-Arrangement: Turbine – EtOH – Oxidiser

-Total mass: ~ 35 kg

– Propellant system:LOX / EtOH resp. HTP / EtOH

-Exit pressure: 7,5 MPa

-Gas generator (open cycle)

-Max. 30,000 RPM

Observation/recommendation

The primary solution to solving the implementation of the Nigeria Space Program is capacity building. In that there is need for more local-based engineers to be trained in all aspects of the space launch program so that they won’t have to outsource for specialists to handle the various aspects of system engineering, project management, manufacture, test, assembly or integration as well as final system testing of a spacecraft (GodstimeKadiri James, Joseph Akinyede, Shaba Ahmad Halilu, 2011). Achievement of capacity building has been realized in the building to flight specification and launching of Nigerian Sat-X, unlike the NigeriaSat-1 TM.

Conclusion

The aim of sending the Nigerian satellites to space is for them to provide imagery and thus provide observation continuity through the DMC payload. The images from the satellites will serve as a catalyst to the growth of Nigeria National Geospatial Data Infrastructure) NGDI program. In addition, the satellites launch acts as a chance for capacity building for the chosen 25 Nigerian engineers to get the skills for them to be able to control and manage the high-resolution minisatellites in space. Possibly in the future Nigerian engineers should be able to manage the entire space launch program, and deliver the satellite in accordance to recommended flight specifications. The space launch program is not only considered a major benefit to Nigeria but it helps with towards its strive of being a modern and industrialized economy.